Method for maintaining good performance in gas turbine during air flow distortio

Humenik, F. M.

Norgren, C. T.

Roudebush, W. H.

English

NASA Technical Reports Server

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION 
WASHINGTON. D.C. 20546 
It1 I’LY to  
A l T N  Of.  GP 
To: USI/Scientific & Technical Infomation Division 
Attention: Mia8 Winnie M. Morgan 
FROM: G P / O f f i c e  of Assistaqt General Counsel for 
Patent Matters 
I 
SUBJECT: Announcement of NASA-Owned U. S. Pa ten t s  i n  STAR 
I n  accordance w i t h  the procedures agreed upon by C o d e  GP 
and C o d e  U S I ,  the attached NASA-owned U. S. Patent is being 
forwarded hor abstracting and announcement i n  NASA STAR. 
The following information is providedt 
U. S. Patent No. 
Government o r  
Corporate Employee 
Supplementary Corporate 
Source (if applicable) : fl/ . 
NASA Patent C a s e  No. 
NOTE - I f  this patent covers an invention made by a corporate 
employee of a NASA Contractor, the following is applicable% 
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  
inventor (author) appears a t  the heading of Column No. 1 of 
the Specification, following the words , with sespect to  
Y e s  0 N o  B 
1 Enclosure 
Copy of Patent aited abov 2 
s 
E 
(CODE) 
2l3 
(CATEGORY) 
1 X (NASA CR OR TMX OR AD NUMBER) 
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https://ntrs.nasa.gov/search.jsp?R=19710019439 2020-03-17T03:15:12+00:00Z
3,581,492 
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SHEET 1 OF 3 
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I V  FIG. I 
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56 
56 
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FIG. 6 INVENTORS 
C A R L  T. NORGREN 
WILLIAM H. ROUOEBUSH 
FRANCIS M. HUMENIK BY 
SHEET 2 OF 3 
PRIOR ART COMBUSTOR 
/30 
FIG. 2 
SIDE ENTRY COMBUSTOR 
FIG. 3 
BY 
INVENTORS 
C A R L  T. NORGREN 
WILLIAM H. ROUDEBUSH 
FRANCIS M. HUMENIK 
ATTORNEYS 
c 
PATENTED JUN 1 1971 31581,492 
SHEET 3 OF 3 
-1 
- -  i 
, 
INVENTOR 
CARL T. NORGREN 
WILLIAM H. ROUOEBUSH 
BY FRANCIS M. HUMENIK 
ATTORNEYS 
t 
Y 
ited States Patent I 1 1 1  3,581,492 
lnventon Carl T. Norgren; 
William H. Roudebush, North Olmsted; 
Francis M. Humenik, Parma, Ohio 
Appl N o  839,994 
Filed July 8,1969 
Patented June 1,1971 
Assignee The United States of America as represented 
by the Administrator of the National 
Aeronautics and Space Administration 
GAS TURBINE COMBUSTOR 
8 Claim, 6 Drawing Figs. 
U.S. CI ........................................................ 60139.36, 
60/39.65,43 1/352 
Int. CI ......................................................... F02c 3/06 
Field of &r& ............................................ 60/39.36, 
39.65; 431/352 
SIDE ENTRY COMBUSTOR 
1561 References Cited 
UNITED STATES PATENTS 
2,993,337 7/1961 Cheeseman .................. 60/39.65 
3,338,051 8/1967 Chamberlain et a1 ......... 60/39.65 
Primary Examiner-Carroll B. Dority, Jr. 
Attorneys-N. T. Musial, G. E. Shook and G. T. McCoy 
2,841,958 7/1958 Stokeset a1 .................. 60/39.65 
ABSTRACT Maintaining performance in a gas turbine com- 
bustor during periods of airflow distortion. A diffuser duct is 
between a liner forming the combustor and a housing sur- 
rounding the liner. Entry ports in the combustor liner have 
scoops which extend the full height of the diffuser duct to 
maintain the proper airflow distribution,regardless of inlet air- 
flow distortion. 
e 
3,58 ; ,492 
A t 
GAS 'FUZBEHE COKBUSTOB 
ORISEN OF THE INVENTEON 
The invention &%Pibed herein was ma& by employees of 
the United States Government and may be manufactured and 
used by or for the Government for government& purFses 
without the payment of any royalties thereon or therefor. 
BACRGROLTN3 OF THE INVENTBQN 
This invention is concerned with maintaining good per- 
fmmance in a w turbine combusm during periods of airflow 
distortion. The invention is particularly directed to providing 
Bow profile from the compressor is distorted. 
mounted on a spindle ZQ. A combustor $8 enclosed by a hous- 
ing %9 i s  positioned between the compressor 112 and the tur- 
bine 84. Air from the compressor 14 i s  heated in the com- 
bustor 18 and then directed through the turbine 14 in a 
manner known in the art. 
One form or" combustor has a torus-shaped chamber 20 
between the housing 19 and an inner casing 2: which encir- 
cles the spindle If Fuel for combustion is injected into the 
chamber 2@ through a plurality of nozzles 22. Swlrlers 24 may 
be used with the nozzies 22 for improved combustion. 
The prior art combustor BI is shown in greater detail in FIG. 
2. air from the $2 through an inlet 
3Q a liner 2T,  This is divided between an inner 
annulus between the inner casing 23 and the liner 27. The 
outer duct 30 comprises the annular space between the hous- 
ing E3 and the liner 27. 
A portion of the air in the inlet passage 26 is also directed to 
the swirlers 24 through snouts 31. The compressed air in the 
inner duct 28 is admitted into the combustion chamber 28 
through suitable ports 32 and 34. In like manner, compressed 
air in the outer duct 3? passes through ports 36 and 38. 
The iryrows in the inlet passage 26 illustrate a distorted radi- 
25 al air flow profile. The arrows in the ducts 2P and 30 show the 
a of these dis&,rtions the amount of air entering 
31 will not be the Same as 
produce local overheating. 
present invention in which air for the combustion process en- 
proper through each entry port even if the duct 21 and an outer duct $9, The duct 2251 is in the form of an 
Certain combustion problenis have been encountered in gas 
deceleration, off design operation of the engine or local en- 
vironmental conditions. %'ne problems have become critical in 
advanced engines which operate near the maximum capability 20 . 
ofthe materials from which the engine is constructed. 
turbine engines due io flow distofiions caused by 
SUMMARY QFTWE INVENTION 
bustor construction in accordance with the present invention 
wherein air from the compressor IS ducted into a single annu- 
velocity profile across the duct reflects the velocity profile at 
the compressor exit. 
Air is zdmitted into the combustion chamber through entry 
completely to the "'l 
'!lese problems have been so'ved i~ a gas 
distoitlotls of the in these ducts. 
di%!Selser that is adjacent the COmbUStiOll chamber The the combustlon ,.hambe,. 20 through the ports 32, 3$,3Q, and 
a inlet 
30 Consequently, combustion problems occur which may 
prts in a predetermined pattern. A scsOP is attached to each A side entry combustor constructed in accordance with the 
POru, and each 
height of the diffuser duct. fn this manner the mass flow 
the radial flow profile from the compressor becomes inverted, 
the airflow into each port will remain the same. Thus ttle com- 
bustion chamber will contain the same fuel air distribution 
which assures that no hot temperature streaks will occur due 40 
to Iwal decrease in airiiow. 
ters subtantially from a single annulus is shown in s. scmP and into the combustion chamber is always 35 Here again compressed air from the compressor 12 flows 
through an inlet passage 26 toward a liner 27. m i s  air then Pro~rt ional  to the 'Ow "Om the compressor' Even if 
passes into an annuler diffuser duct 4@ that is adjacent to the 
combustion chamber 20 between the liner 2'7 and the housing 
29. The radial air profile across the diffuser duct 40 shown by 
the arrows in FiG. 3 will reflect the radial air profile at the 
compressor exit in the passage 24. 
Air from the diffuser duct 40 is admitted into the com- 
bustion chamber 28 in a predetermined entry pattern through 
Et is, therefore, an object of the present invention to main- primary entry ports 42 md secondary entry ports 44 in the 
tain good peflomance in a gas turbine combustor with no 45 liner 27. Fuel is supplied to the nozzles 22 from pipes 45. A 
b d  overheaeing during perids  offlow distortion caused by portion of the air in the inlet passage 24 is directed to the swir- 
acceleration, deceleration, off-design operation, and local en- lers 24 at the nozzles 22 through inlet diffusers &. 
vironmentd conditions. An important feature of the invention is the provisions of 
Another object of the invention is to accommodate flow scoops 48 at the primary entry ports 42. The scoops 48 extend 
distortions of d r  to a gas turbine combustor so that the engine 50 to the full height of the diffuser duct 40 from the liner 27 to 
can be operated near the maximum CaPabib' ofthe materials. the housing 19. Each scoop 48 has an open front which faces 
These and other objects of the invention will be apparent toward the flowing air. A curved backwall 49 on each SCOOP from the specification which follows and from the drawings 4g extends from the downstream edge of each prima- 
wherein like numerals are used throughout to identify like 55  ry entry port 42 the full height of the duct 4o to the housing 
parts. 19. Vertical sidewalls 51 extend from opposite edges of each 
primary port 4.2 to the curved backwall 49. If desired, the wall 
49 may be a straight plate which extends diagonally from the 
downstream edge of the porp 42 to the housing 19 opposite the 
In this manner the mass flow through each scoop 48 and 
mass flow from the compressor. Even if the radial flow profile 
shown by the arrows in FIG. 3 from the compressor becomes 
65 inverted the airflow into each primary entry port 42 will 
remain the same. Thus the combustion chamber 26) will have 
the same fuel air distribution which, in turn, will assure that no 
hot temperature streaks will occur due to local decreases in 
aidlow. 
The operation of the side entry combustor shown in FIG. 3 
is illustrated in greater detail in FIG. 4. The combustion 
chamber 28 encircles the spindle 16 which mounts the com- 
pressor at one end and the turbine at the other. Fuel passes 
through pipes 45 and is injected into the chamber 28 through 
0BJBc"FS OF THE IPdVEPlTlON 
DESESCRIPTTION OF THE DRAWINGS 
FIG. 1 is an axial quarter section view of a typic& gas tUr- 
bine engine showing the relative positions of its various COm- 60 upstream edge of the same primary port. 
ponents; 
bustor; 
structed in accordance with the present invention; 
accordance with the invention; 
lustrating an alternate embodiment of the invention; and 
present invention. 
c FIG. 2 is an enlarged sectional view of a prior art COm- into the combustion chamber 28 is always proportional to the 
FIG. 3 is an enlarged sectional view of a combustor con- 
FIG. 4 is a perspective view of a combustor constructed in 
FIG. § is a perspective view of a portion of a combustor il- 
F16. 5 is a plan view of still another embodiment of the 70 
B)ESCR~TION THE 
Refebning now bo the drawings there is shown in FIG. 1 a gas 
r 
EMBODIMENT 
turbine engine I@ having a compressor 12 and a Furbine 2.4 75 nozzles as previously described. 
I 
3 
3,581,492 
Compressed air from the compressor enters the diffuser 
duct 40 as shown by the arrows A. A portion of this air is inter- 
the primary entry port 42. These 
rcepted air into the combustion 
chamber 20. Still another portion of the air enters the com- 5 
bustion chamber 28 through the secondary ports 44. 
A small amount of air passes through the inlet diffusers 46 
into a cooling duct 50 as shown by the arrow B. This cooling 
air passes through cooling apertures 52 into the combustion 
chamber 20. The remaining air from the diffusers passes to the IO 
swirlers 24 which surround the nozzles 22. 
DESCRIPTION OF ALTERNATE EMBODIMENTS 
In the alternate embodiment shown in FIG. 5 scoops 54 are 
provided at each of the secondary ports 44. The scoops 54 
likewise extend to the full height of the diffuser duct 40. 
In the embodiment shown in FIGS. 4 and 5 the secondary 
are staggered relative to the primary entry ports 42. 
The embodiment shown in FIG. 6 has the primary and secon- 
dary ports aligned. Each primary port is provided with a scoop 
56 that is aligned with another scoop 58 on the secondary 
port. As seen in FIG. 6 the secondary scoop 58 is larger than 
the primary scoop 56 to obtain the proper proportioning of 
the air. 
The embodiment shown in FIG. 6 also provides a set of 
secondary ports 60 which are staggered relative to the primary 
entry ports. This results in combinations of entry ports which 
may be sized to suit desired combustor performance. When 
scoops are attached to entry ports, the scoops extend to the 
full height of the diffuser duct. 
While several embodiments of the invention have been 
shown and described it will be appreciated that various struc- 
tural modifications may be made without departing from the 
spirit of the invention or the scope of the subjoined claims. For 
example, it is contemplated the secondary ports 44 may be 
staggered with respect to the primary ports 42 as shown in 
FIG. 5 yet have different size scoops as shown in FIG. 6. What- 
ever the arrangement all the scoops must extend to the full 
height of the diffuser duct to obtain at all times correct pro- 
portional amount of air which is to be used for combustion 
purposes. This amount of air is independent of the compressor 
outlet radial velocity profile. 
We claim: 
1. A combustor for a gas turbine engine of the type having a 
compressor and a turbine mounted on a spindle, said com- 
bustor comprising 
a curved liner forming a torus-shaped chamber encircling 
said spindle between said compressor and said turbine, 
said liner having a plurality of spaced ports therein, each 
of said ports having an upstream edge toward said com- 
pressor and a downstream edge away from said compres- 
sor; 
means for injecting fuel into said torus-shaped chamber, 
a housing surrounding said liner and spaced therefrom to 
form an annular diffuser duct, said duct being in commu- 
nication with said compressor for receiving compressed 
air therefrom whereby compressed air is admitted to said 
a plurality of scoops in said diffuser duct adjacent some of 
said ports and substantially surrounding said lines for in- 
tercepting a portion of the flow of compressed air, each of 
said scoops having a pair of spaced sidewalls extending 
from said liner toward said housing, a backwall connect- 
ing the outermost edges of each of said pairs of sidewalls 
whereby the space between the sidewalls at adjacent 
ports is substantially unobstructed to the flow of com- 
pressed air, said backwall extending the full height of said 
diffuser duct from said downstream edge of said port to 
said housing substantially opposite the upstream edge of 
said port, and an open front between said sidewalls ex- 
tending from said upstream edge of said port to said hous- 
in whereby the mass airflow through said open front and 
a8acent port into said combustion chamber is propor- 
tional to the mass flow in said diffuser duct. 
2. Apparatus as claimed in claim 1 wherein the ports are in 
at least two adjacent rows, one of said rows being upstream 
from the other, 
P 
torus-shaped chamber through said ports. and 1 
15 
20 
25 
primary entry ports being in said one row, and 
secondary entry ports being in said other row. 
3. Apparatus as claimed in claim 2 including scoops on the 
4. Apparatus as claimed in claim 2 including scoops on the 
5. Apparatus as claimed in claim 2 wherein the primary 
35 entry ports and secondary entry ports are aligned in the 
direction of airflow. 
6. Apparatus as claimed in claim 2 wherein the primary 
entry ports and the secondary entry ports are staggered in the 
direction of the airflow. 
7. Apparatus as claimed in claim B wherein backwall por- 
tion of the scoop comprises 
a member curving outwardly from said downstream edge of 
8. Apparatus as claimed in claim 1 wherein the backwall 
a substantially straight member extending diagonally out- 
ward from said downstream edge of the adjacent port to a 
housing substantially opposite the upstream edge of said 
Port. 
30 
primary entry ports. 
secondary entry ports. 
40 
the adjacent port to the housing. 
45 portion of the scoop comprises 
50 
55 
60 
P 
65 
70 
75 


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