U.S. patent application number 09/161104 was filed with the patent office on 2002-10-24 for fuel mixing in a gas turbine engine.
Invention is credited to RAKHMAILOV, ANATOLY.
Application Number | 20020152753 09/161104 |
Document ID | / |
Family ID | 22579835 |
Filed Date | 2002-10-24 |
United States Patent
Application |
20020152753 |
Kind Code |
A1 |
RAKHMAILOV, ANATOLY |
October 24, 2002 |
FUEL MIXING IN A GAS TURBINE ENGINE
Abstract
A gas turbine engine has a device for admitting a rotating fluid
flow from an annular space of the casing to the inlet portion of a
combustor to form a rotating fluid flow in the inlet portion of the
combustor. The rotating fluid flow is formed in the annular space
of the casing by supplying a fluid from a compressor to the blades
of the turbine rotor disk.
Inventors: |
RAKHMAILOV, ANATOLY;
(BATAYSK, RU) |
Correspondence
Address: |
PETER A SULLIVAN
HUGHES HUBBARD & REED
ONE BATTERY PARK PLAZA
NEW YORK
NY
100041482
|
Family ID: |
22579835 |
Appl. No.: |
09/161104 |
Filed: |
September 25, 1998 |
Current U.S.
Class: |
60/772 ;
60/804 |
Current CPC
Class: |
F01D 9/023 20130101;
F01D 9/04 20130101; F04D 27/0215 20130101; F23R 3/54 20130101; F02C
7/08 20130101; Y02T 50/60 20130101; F01D 5/081 20130101; F01D 9/065
20130101 |
Class at
Publication: |
60/772 ;
60/804 |
International
Class: |
F02C 003/14 |
Claims
I claim:
1. A method of operation of a gas turbine engine having a
compressor for producing a fluid flow, a casing, a combustor in
said casing, said combustor having an inlet portion, a turbine
rotor disk with blades, and an annular space in said casing, said
annular space surrounding said blades, said method comprising:
supplying fuel and combustion air to said combustor to prepare a
heated fluid; supplying said heated fluid directly from said
combustor to said blades; supplying said fluid flow from said
compressor to said blades to form a rotating fluid flow in said
annular space; feeding at least a part of said rotating fluid flow
into said inlet portion of said combustor.
2. The method of claim 1, wherein said fuel is supplied into said
rotating fluid flow within said inlet portion of said
combustor.
3. A method of operation of a gas turbine engine having a
compressor for producing a fluid flow, a casing, a combustor in
said casing, said combustor having an inlet portion, a turbine
rotor disk with blades, and an annular space in said casing, said
annular space surrounding said blades, said method comprising:
supplying said fluid flow from said compressor to said blades to
form a rotating fluid flow in said annular space; feeding at least
a part of said rotating fluid flow into said inlet portion of said
combustor; feeding said fuel into said rotating fluid flow within
said inlet portion of said combustor preparing a heated fluid in
said combustor by burning said fuel and air in said combustor;
supplying said heated fluid directly from said combustor to said
blades.
4. A gas turbine engine, said gas turbine engine comprising: a
compressor for producing a fluid flow; a fuel source; a combustion
air source; a casing; a combustor in said casing, said combustor
having an annular inner wall and an inlet portion, said combustor
communicating with said fuel source and with said combustion air
source to prepare a heated fluid; a turbine rotor disk with blades
said blades positioned immediately downstream of said combustor for
receiving said heated fluid from said combustor; an annular space
in said casing, said annular space surrounding said blades; a zone
upstream of said turbine rotor disk, said zone communicating with
said compressor for supplying said fluid flow from said compressor
to said blades to form a rotating fluid flow in said annular space;
a means for admitting said rotating fluid flow from said annular
space to said inlet portion of said combustor, whereby a rotating
fluid flow is formed in said inlet portion of said combustor.
5. The gas turbine engine of claim 4, wherein said fuel source
communicates with said inlet portion of said combustor.
6. The gas turbine engine of claim 4, wherein said means for
admitting said rotating fluid flow from said annular space to said
inlet portion of said combustor comprises: an annular guide wall
that is installed in said combustor in a spaced relation to said
annular inner wall of said combustor, said annular guide wall
defining with said annular inner wall of said combustor said inlet
portion of said combustor; said inlet portion of said combustor
communicating with said annular space of said casing.
7. The gas turbine engine of claim 6, wherein said fuel source
communicates with said inlet portion of said combustor.
8. A gas turbine engine, said gas turbine engine comprising: a
compressor for producing a fluid flow; a fuel source; a combustion
air source; a casing; a combustor in said casing, said combustor
having an annular inner wall and an inlet portion, said combustor
communicating with said combustion air source to prepare a heated
fluid by burning said fuel with said combustion air; a turbine
rotor disk with blades said blades positioned immediately
downstream of said combustor for receiving said heated fluid from
said combustor; an annular space in said casing, said annular space
surrounding said blades; a zone upstream of said turbine rotor
disk, said zone communicating with said compressor for supplying
said fluid flow from said compressor to said blades to form a
rotating fluid flow in said annular space; an annular guide wall
that is installed in said combustor in a spaced relation to said
annular inner wall of said combustor, said annular guide wall
defining with said annular inner wall of said combustor said inlet
portion of said combustor; said inlet portion of said combustor
communicating with said annular space of said casing; said fuel
source communicates with said inlet portion of said combustor.
Description
[0001] This application cross-references three copending U.S.
patent applications, each of which was filed on Sep. 25, 1998, as
U.S. patent application Ser. No. 09/______, 09/______ and
09/______, respectively, each of which copending U.S. applications
is incorporated herein by reference.
[0002] The invention relates to the field of gas turbine engines,
and more specifically, to an improved gas turbine engine using a
rotating fluid flow train to feed the combustor and enhance
air/fuel mixing and emissions.
BACKGROUND OF THE INVENTION
[0003] A type of prior art gas turbine engine has a compressor, a
fuel source, a combustion air source, a casing, and a combustor to
prepare a heated fluid from fuel and combustion air. The combustor
is connected to the fuel source, the combustion air source and the
compressor. Practically the entire fluid flow from the compressor
is directed to the combustor. The engine has a turbine rotor disk
with blades that receive the heated fluid from the combustor. As
the turbine rotor disk rotates during engine operation, the heated
fluid flow coming from the combustor has to be directed at an angle
to the blades to ensure smooth entry conditions. This is done using
stator vanes that are positioned at a certain angle and direct the
heated fluid from the combustor to the turbine rotor disk in a
manner compatible with rotor disk rotation. This gas turbine engine
is disclosed in U.S. Pat. No. 3,826,084 to Branstrom et al.
[0004] The stator vane angle normally is chosen to accommodate the
most optimum and prevailing turbine rotor disk operating conditions
(speed). This solution is quite acceptable for gas turbine engines
that have more or less stable operating conditions, such as when
used for power generation. In applications where the load upon the
gas turbine engine is steady, the turbine rotor disk rotates at a
stable speed, and the entry angle for the blades remains unchanged
thus minimizing losses. If, on the other hand, this gas turbine
engine is used to power a vehicle, the situation is radically
different. In that application, the turbine rotor disk speed will
vary within a broad range depending on vehicle load. Consequently,
the entry angle also varies within a broad range under load
fluctuations, which leads to greater losses. This problem could not
be solved by using the conventional approach with the stator vanes.
It is possible to use controllable stator vanes to change the entry
angle at the blades, but it is a very complicated and expensive
solution given the high temperatures downstream of the combustor
and space limitations. As a result, the gas turbine engine would
have high losses in vehicle applications. Moreover, the stator and
vanes occupies an additional space and makes the engine design more
complicated and expensive. The use of controllable vanes makes the
engine less reliable.
[0005] The problems indicated above are solved in the gas turbine
engine of this invention.
SUMMARY OF THE INVENTION
[0006] It is an object of the invention to provide a gas turbine
engine of the above type that has a higher efficiency.
[0007] Another object of the invention is to provide a more compact
gas turbine engine that has a simpler design.
[0008] Another object of the invention is to improve the emission
characteristics of the gas turbine engine.
[0009] A gas turbine engine has a device to admit a rotating fluid
flow from an annular space in the casing to the inlet portion of a
combustor to form a rotating fluid flow in the inlet portion of the
combustor. The rotating fluid flow is formed in the annular space
of the casing by supplying a fluid from a compressor to the blades
of the turbine rotor disk.
[0010] Other objects and advantages of the invention will become
apparent from the following detailed description of preferred
embodiments and accompanying drawings.
DETAILED DESCRIPTION OF THE DRAWINGS
[0011] FIG. 1 shows a diagrammatic view of a gas turbine engine
according to the invention.
[0012] FIG. 2 is a sectional view of an embodiment of the annular
space (Leonid to supply a sketch.)
DETAILED DESCRIPTION OF THE DRAWINGS
[0013] With reference to FIG. 1, a gas turbine engine has a casing
10, a compressor 12 for supplying a compressed fluid, a turbine
rotor disk 14 mounted downstream of compressor 12 installed on the
turbine rotor, a combustor 16 to prepare a heated fluid to be
supplied to turbine rotor disk 14. Combustor 16 has a port 18 to
admit fuel supplied from a fuel source (not shown). Combustor 16
defines a combustion zone 20 in which the heated fluid is formed.
Combustion air is supplied from an air source (not shown) as shown
by arrows A to an inlet portion of the combustor in which port 18
is provided.
[0014] The inlet portion of the combustor shown at 19 is defined by
an inner annular wall 22 of combustor 16 and by an annular guide
wall 24 that extends within the combustor in a spaced relation to
annular inner wall 22. Annular guide wall 24 is installed by
brackets 26 in such a manner that a space 28 is left for fluid
passage.
[0015] A part of the fluid from compressor 12 is supplied to
turbine rotor disk 14, bypassing combustor 16, as shown by arrows
B, through passage 30 in casing 10 and reaching a zone 32 upstream
of turbine rotor disk 14. Vanes 34 can be provided in passage 30 to
make this fluid flow compatible with the turbine rotor disk 14
rotation. These vanes will function in an optimum manner only under
certain turbine engine operating conditions. Since the quantity of
fluid that is fed to the turbine rotor disk 14 and the velocity of
this fluid are not very high, losses that would occur under
non-optimum conditions would be relatively low. This fluid is
admitted to turbine rotor disk 14 and envelops the blades 15. The
fluid from the compressor 12 passes through a passage 36 of the
blade 15 and leaves the passage 36 to reach an annular space 38
that is defined in casing 10 and surrounds blades 15. As the blades
15 rotate, the fluid from the compressor 12 leaves blade passage 36
having obtained a rotation that forms a rotating fluid flow in
annular space 38. This rotating fluid flow is admitted through
space 28 to inlet portion 19 of combustor 16 to form a rotating
fluid flow there. As fuel is fed through port 18, it is entrained
in a rotary motion by the rotating fluid flow in the inlet portion,
and intense stirring and mixing of fuel and fluid will take place
to prepare a good quality fuel mixture. The rotating fluid flow
entrains air that is fed as shown by arrow A, moves into combustion
zone 20, and imparts a spin to the heated fluid when it is formed
in combustion zone 20. The direction of this rotating flow is the
same as the turbine rotor disk direction of rotation and the
velocity of this rotating flow steadily follows turbine rotor disk
14 rotation velocity (with a very short lag). The heated fluid
formed in combustor 16 will move to the turbine blades 15 in a
manner that is almost entirely compatible with rotation of the
turbine rotor disk. Consequently, losses in this zone, which
account for most of the losses in the turbine flow duct, are
minimized.
[0016] Another advantage of the invention is that the fluid from
the compressor that goes through passage 36 and reaches blade 15
cools the blade and the adjacent wall of casing 10.
[0017] The intensive mixing and stirring of fuel, air, and the
fluid that comes from the compressor in inlet portion 19 provides
almost ideal conditions to prepare a fuel mixture. This high
quality fuel mixture provides better conditions for combustion and
improves the emission characteristics of the engine.
[0018] Another advantage of the invention is the method of
preparation of the fuel mixture. The quantity of fuel supplied for
small-power gas turbine engines is rather low. It is very difficult
to prepare a homogeneous fuel mixture with a ratio of fuel to air
and fluid of 1:15 to 1:30. The fuel mixing method that is used here
solves this problem. When fuel is entrained in a rotary motion by
the rotating fluid flow admitted to the inlet portion of the
combustor, fuel atomizing, mixing and stirring in the rotating flow
are very thorough and intensive. This thoroughness assures a high
degree of homogeneity of the fuel mixture.
[0019] FIG. 2 shows an embodiment of the space 28 with annular
guide walls 24 attached by brackets 26. This space 28 can take the
form of an arc slit cut in a flanged portion of the annular guide
wall or in the form of spaces between the adjacent brackets (not
shown).
* * * * *