A swirl can full annulus combustor, operable over a wide range of exit average temperatures is used in a high performance gas turbine engine for advanced aircraft. A large number of swirl can modules are mounted in an array in each combustor

Jones, R. E.

Niedzwiecki, R. W.

English

NASA Technical Reports Server

REPLY TO
ATTN OF:
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
WASHINGTON. D.C. 20546
SEP o 5 1973
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TO: KSI/Scientific & Technical Information Division
Attention: Miss Winnie M. Morgan
PROM: GP/Office of Assistant General Counsel for
Patent Matters
» i- SUBJECT: Announcement of NASA-Owned U.S. Patents in STAR
• i
£1, In accordance with the procedures agreed upon by Code GP
^ and Code KSI, the attached NASA-owned U.S. Patent is being
0iforwarded for abstracting and announcement in NASA STAR.
The following information is provided:
U.S. Patent No. : 3, % 5 3»
Government or
Corporate Employee
Supplementary Corporate
Source (if applicable)
NASA Patent Case No.
fb"T> oer-/\'/ne./»
/. £ dO-
NOTE - If this patent covers an invention made by a corporate
employee of a NASA Contractor, the following is applicable:
Yes / / No / XV
Pursuant to Section 305(a) of the National Aeronautics and
Space Act, the name of the Administrator of NASA appears on
the first page of the patent; however, the name of the actual
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the Specification, following the words "... with respect to
an invention of ,
Elizabeth A. Carter
Enclosure
Copy of patent cited above lui
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https://ntrs.nasa.gov/search.jsp?R=19730021933 2020-03-17T09:09:18+00:00Z
United States Patent
Jbnes et al.
tin 3,748,853
[45] July 31, 1973
[54] SWIRL CAN PRIMARY COMBUSTOR
[75] Inventors: Robert E. Jones, Berea; Richard W.
Niedzwiecki, Cleveland Heights,
both of Ohio
[73] Assignee: The United States of America as
represented by the Administrator of
the National Aeronautics and Space
Administration, Washington, D.C.
[22] Filed: Oct. 27, 1971
[21] Appl. No.: 192,970
[52] U.S. Cl... 60/39.65, 60/39.72, 60/39.74 R,
60/39.66,431/9,431/173
[51] Int. Cl F02c 1/00
[58] Field of Search 60/39.65, 39.23,
60/39.29, 39.72, 39.74 R; 431/9, 173
[56] References Cited
UNITED STATES PATENTS
3,267,676 8/1966 Sneeden 60/39.65 X
2,934,894 5/1960 Lang 60/39.65 X
2,999,359 9/1961 Murray 60/39.65
Primary Examiner—Clarence R. Gordon
Attorney—N. T. Musial,John R. Manning et al.
[57] ABSTRACT
A swirl can full annulus combustor operable over a
wide range of exit average temperatures is used in a
high performance gas turbine engine for advanced air-
craft. A large number of swirl can modules are
mounted in an array in each combustor.
4 Claims, 3 Drawing Figures
501
PATENTED JUl 311973 3.748,853
22
12
FIG. I
FIG. 3
INVENTORS
ROBERT E.JONES
RICHARD W. NIEDZWIECKIBY
^X^V^^vtX^ *-"•
ATTORNEYS
. • ' 3,748,853
SWIRL CAN PRIMARY COMBUSTOR Another object of the invention is to maintain high
STATEMENT OF GOVERNMENT OWNERSHIP combustion efficiency in a gas turbine engine combus-MAlbMbNl Of UOVtKNMbNl UWNtKJ>HlP tor over a wide range of exit average temperatures.
The invention described herein was made by employ- A further object of the invention is to provide an im-
ees of the United States Government and may be man- 5 proved combustor having a length about half that of
ufactured and used by or for the Government for gov- conventional combustors.
ernmental purposes without the payment of any royal- A still further object of the invention is to reduce ni-
ties thereon or therefor. trogen oxide levels by accelerated mixing of combus-
PRIOR ART tion gaSCS and addit'onl air-10 These and other objects of the invention will be ap-
This invention is concerned with a combustor having parent from the specification which follows and from
improved performance for use in an improved gas tur- the drawing wherein like numerals are used throughout
bine engine. The invention is particularly directed to a to identify like parts.
full annulus swirl can primary combustor operable over , „„
a wide range of combustion average exit temperatures 15 DESCRIPTION OF THE DRAWINGS
with high efficiency. FIG. 1 is a sectional view of a high temperature corn-
Certain advanced aircraft need improved engines for bustor constructed in accordance with the present in-
propulsion at high cruising speeds. Gas turbine engines vention;
of minimum weight have been proposed for take-off FIG. 2 is an enlarged end view of a swirl can module
and acceleration to the high cruise Mach numbers of taken along the line 2 — 2 in FIG. 1; and
such aircraft. Problems have been encountered in these FIG. 3 is a sectional view taken along the line 3 — 3
engines because present combustors operate at rela- in FIG. 2.
lively low exit temperatures and use complicated liners INSCRIPTION OF THF PRFFFRRFD
to admit mixing air into the hot combustion gases. DESCRIPTION OF THE PREFERRED
The performance of conventional combustors de- EMBODIMENT
creases when they are operated at high exit average Referring now to the drawings there is shown in FIG.
temperatures. Combustion efficiency especially de- 1 a combustor 10 used with a gas turbine engine having
creases at high exit temperatures approaching stoichio- a compressor and a turbine mounted on a spindle. The
metric temperatures. Also large amounts of combustor
 3Q combustor 10 is positioned between the compressor
total airflow are required to cool the complicated lin- and the turbine so that air from the compressor is
ers. This limits exit temperatures because the air is not heated in the combustor and directed through the tur-
available for combustion. bine in a manner well known in the art. A prior art
Acoustic instability may be a problem at high exit combustor in a gas turbine engine is shown in U.S. Pat.
temperatures in conventional combustors. Smoking 35 No. 3,581,492.
problems may be encountered at the fuel-air ratios re- The combustor 10 has an outer wall 12 spaced from
quired for high exit temperatures. The complicated an inner wall 14. Both the walls 12 and 14 encircle the
liner configurations required in these combustors ere- gas turbine engine spindle. Air from the compressor is
ate durability problems. Long combustor lengths result supplied to the space between the walls 12 and 14
in greater liner surface areas to be cooled. The conven- 40 through an inlet air duct 16. This air moves past dif-
tional combustors also utilize atomizing fuel nozzles fuser inserts 18 positioned in the end of the duct 16.
which require high pressure fuel pump systems. According to the present invention a plurality of swirl
....... . „,, ,»„ ^ ,,T, ,»,,,,^,~,~,., can modules 20 are mounted in the combustor 10 asSUMMARY OF THE INVENTION
 shown fa F,G ^ By way of eMmp,e JM swir, ^
These problems have been solved by a full annulus 45 modules 20 were mounted in three consecutive circles
combustor utilizing many swirl can modules. Fuel is in- in a 42 inch diameter combustor. Lines 22 from a suit-
jected into the combustor at each module. Combustion able supply of liquid fuel are connected to the swirl can
air enters each module and mixes with the fuel. The modules. Valves 23 control the flow of fuel to mani-
fuel-air mixture passes through a swirler and is ignited folds 25 which are connected through choke orifices 27
downstream. Secondary combustion air flows axially SO to the lines 22. Three manifolds 25 were utilized in the
past the modules, recirculates in their wakes, and com- aforementioned 42 inch diameter combustor. One
pletes the combustion reaction. Essentially all of the manifold was connected to the swirl can modules 20 in
combustion air is passed through the module array and each of the three consecutive circles. Thus fuel is in-
burning zone. jected at 120 locations in the aforementioned combus-
Premixed fuel and air are injected into the combustor 5S tor compared to approximately 30 for a conventional
at many more locations than in a conventional combus- combustor.
tor thereby providing a better opportunity for homoge- Referring now to FIGS. 2 and 3 each swirl can mod-
neous mixture of fuel and air. Fuel flow is readily con- ule 20 has a hollow cylindrical housing 24 which serves
trolled at each module and varying amounts of fuel can as a carburetor. The end of the can facing the air duct
be supplied to compensate for nonuniformity of air 16 is open at inlet 26 as shown in FIG. 3 to enable corn-
flow profiles. bustion air to enter the can. Fuel enters the can adja-
OBJECTS OF THE INVENTION *" "" * fr°m *" ^  Une " *" "
It is, therefor, an object of the present invention to ,. A swirl plate 28 is mounted on the opposite end of
provide an improved primary combustor for a high per- the can housing 24 from the inlet opening 26. The swirl
formance gas turbine engine that is operable up to and plate 28 has a number of vanes 30 which are angularly
beyond the stoichiometric temperature. disposed relative to the exit plane of the can. An angle
3,748,853
between 10° and 35° has been satisfactory for the vanes
30 in the swirl plate 28.
A flame stabilizer plate 32 having a hexagonal pe-
riphery encircles the swirl plate 28. This flat plate flame
stabilizer 32 is likewise mounted on the end of the can 5
24.
In operation, combustion air enters the can housing
24 through the inlet 26 where it is mixed with fuel from
the line 22. The fuel-air mixture then passes through
the swirl plate 28 and is ignited downstream of the flat 10
plate stabilizer 30 by a suitable igniter 34. The fuel-air
mixture burns in a combustion chamber 36 formed by
a film cooled liner 38. Nozzle fouling problems are
minimized because the fuel orifices in the lines 22 are
outside of the hot zone. IS
Secondary combustion air flows axially past the swirl
can module 20 into the combustion chamber 36 where
it recirculates in the wakes of the modules and com-
pletes the combustion reaction. Because of the large
circumference of the stabilizer plate 32 the interfaced 20
mixing area between the bypass air and the hot gases
in the wake of each module is increased to its maximum
value. The housing 24 of each module 20 does not ex-
tend into the burning zone within the combustion
chamber 36. The mixing of diluent air and combustion 25
products occurs because of recirculation and eddy cur-
rents. The liner 38 has nearly all its metal surfaces dis-
placed from the flame streamline thereby reducing the
liner coolant flow required. A plurality of air entry slots
40 are provided on the leading edge of the liner 38 to 30
supply air to the hottest zones in the combustion cham-
ber 36. This air is not used entirely for cooling because
the entry slots are upstream and the air can be used to
combine with fuel.
The length of the combustor 10 is much less that that 35
of a conventional combustor. By way of example the
aforementioned 42 inch diameter combustor had an .
overall length of 20 Vi inches. The length of the com-
bustor was only 11% inches from the plane of fuel en-
try. With such a short length very little air is required 40
for cooling the liners 38 and essentially all of the com-
bustion air passes through the array of swirl can mod-
ules.
The flow of fuel can be easily controlled into each
module 20 by adjusting the valves 23 and the orifices 45
27. Varying amounts of fuel can be supplied to each
module 20 to compensate for nonuniformity of air flow
profiles. Also, it is not necessary to utilize all modules
at all burning conditions. Under certain conditions,
such as during engine idle, one row of modules can be so
ignited and good performance can be maintained at off
design conditions.
While a preferred embodiment of the invention has
been described it will be appreciated that various struc-
tural modifications can be made without departing
from the spirit of the invention or the scope of the sub-
joined claims. For example, although the flame stabi-
lizer 32 is flat and has a hexagonal periphery other con-
figurations have been utilized. Flame stabilizers in the
form of cones have been used. Also the stabilizer may
be in the form of a star.
What is claimed is:
1. A combustor for a high performance gas turbine
engine comprising
a pair of spaced walls forming a full annulus,
a duct for supplying compressed air to an annular
space between said spaced walls,
a plurality of spaced modules mounted in said annu-
lar space adjacent to said duct whereby said com-
pressed air is directed toward said modules so that
a first portion of said compressed air passes
through said modules and a second portion passes
around said modules in the space therebetween,
each of said modules comprising
a hollow cylindrical housing having an open end for
receiving said first portion of said compressed air
forming primary combustion air,
means for injecting fuel tangentially into said hous-
ing to provide a mixture of air and fuel, and
a plurality of vanes at one end of said housing op-
posite said open end for swirling said mixture of
air and fuel prior to mixing with secondary air,
said secondary air comprising said second por-
tion of said compressed air that flows around said
modules, said first and said second portions com-
prising substantially the entire amount of said
compressed air supplied from said duct,
a film cooled liner on the oposite side of said modules
from said duct forming a combustion chamber for
receiving said mixture of air and fuel and said sec-
ondary combustion air, and
means for igniting said mixture of air and fuel in said
combustion chamber.
2. A combustor as claimed in claim 1 including
means for separately controlling the flow of fuel to cer-
tain of said modules whereby said modules are selec-
tively ignited thereby enabling said engine to be oper-
ated at off design conditions.
3. A combustor as claimed in claim 1 including
means for stabilizing combustion flame adjacent said
vanes.
4. A combustor as claimed in claim 3 wherein the
means for stabilizing combustion flame comprises a
substantially flat plate extending outward from said
vanes.
55
60
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Swirl can primary combustor

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