U.S. patent number 3,826,084 [Application Number 05/411,124] was granted by the patent office on 1974-07-30 for turbine coolant flow system.
This patent grant is currently assigned to United Aircraft Corporation. Invention is credited to Bruce R. Branstrom, Frank W. Huber.
United States Patent |
3,826,084 |
Branstrom , et al. |
July 30, 1974 |
**Please see images for:
( Certificate of Correction ) ** |
TURBINE COOLANT FLOW SYSTEM
Abstract
In a turbine engine compressed air is delivered through a
diffuser to a burner section and then through a turbine section. A
portion of the compressed air from said diffuser is removed by an
annular center body manifold and directed by hollow struts inwardly
to an annular passageway where the compressed air flow is taken to
the forward part of the turbine section. The flow from the annular
passageway is then directed into a passageway having directing
vanes for imparting a desired velocity and direction to the
existing flow so as to be compatible with the rotating turbine disk
onto which it flows. This air is then directed to cool blades on
that disk and also passed through that disk to be directed to
turbine blades on the next disk. The invention herein described was
made in the course of or under a contract with the Department of
the Air Force.
Inventors: |
Branstrom; Bruce R. (Riviera
Beach, FL), Huber; Frank W. (Palm Beach Gardens, FL) |
Assignee: |
United Aircraft Corporation
(East Hartford, CT)
|
Family
ID: |
26708755 |
Appl.
No.: |
05/411,124 |
Filed: |
October 30, 1973 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
Issue Date |
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32687 |
Apr 28, 1970 |
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Current U.S.
Class: |
60/806; 60/751;
415/144; 416/95; 415/115; 415/175; 416/97A; 416/97R |
Current CPC
Class: |
F01D
5/081 (20130101); F02C 7/18 (20130101) |
Current International
Class: |
F01D
5/02 (20060101); F01D 5/08 (20060101); F02C
7/16 (20060101); F02C 7/18 (20060101); F02c
007/18 () |
Field of
Search: |
;60/39.66,39.07,39.31,39.32 ;415/165,175,176,144,145,115 |
References Cited
[Referenced By]
U.S. Patent Documents
Foreign Patent Documents
Primary Examiner: Hart; Douglas
Attorney, Agent or Firm: McCarthy; Jack N.
Parent Case Text
This is a continuation, of application Ser. No. 32,687, filed Apr.
28, 1970 now abandoned.
Claims
We claim:
1. In combination, a turbine engine including an annular inlet
section having a first outer wall and a first inner wall, an
annular combustion section having a second outer wall connected to
said first outer wall and a second inner wall connected to said
first inner wall, a turbine section, said turbine section including
a turbine rotor, said turbine rotor including a turbine disk
mounted for rotation along with turbine blades mounted therearound,
an annular manifold located immediately downstream of said annular
inlet section, said annular manifold having an outer surface and an
inner surface, said annular manifold having an annular inlet for
receiving flow from said inlet section, said annular manifold being
axially in line with the center of said annular inlet section with
said annular inlet of said annular manifold being forwardly into
the center of the annular inlet section for receiving flow passing
therethrough, said annular combustion section having a third inner
wall spaced from said second inner wall providing a first annular
passageway, hollow struts extending between the inner surface of
said annular manifold and the adjacent end of said first annular
passageway, a rotarting annular passageway extending forwardly from
said disk, said first annular passageway having its downstream end
located adjacent the rotating annular passageway so as to direct
cooling air therefrom into the rotating annular passageway, passage
means in said disk being in communication with said rotating
passageway for directing flow to said blades for cooling.
2. A combination as set forth in claim 1 wherein projections extend
from the outer surface of said annular manifold outwardly towards
the second outer wall of the annular burner section, each
projection having an opening extending into the end thereof, means
being provided at a plurality of locations for extending through
said second outer wall into the openings in said projections for
supporting said annular manifold.
3. A combination as set forth in claim 2 wherein said means for
extending into the openings are individually removable pins.
4. A combination as set forth in claim 1 wherein a burner is
located in said annular combustion section, the rearward part of
said burner being connected between said second inner wall and said
second outer wall and positioned to exhaust into said turbine
section, the forward end of said burner having an arm means
projecting outwardly therefrom, said arm means having a second
opening therethrough, said second opening being aligned radially
outwardly from an opening in a projection extending from the outer
surface of said annular manifold, said pin means extending through
said second opening into its cooperating opening in a projection.
Description
BACKGROUND OF THE INVENTION
While cooling air has been directed to turbine blades for cooling,
and compressor air has been used for this purpose, the arrangement
set forth herein is a different device for doing so as will be
later described.
SUMMARY OF THE INVENTION
This turbine coolant flow system is arranged to provide low coolant
temperatures with a minimum of expended horsepower to provide
maximum engine efficiency. The coolant is expanded through nozzle
means in a tangential plane paralleling the direction of turbine
disk rotation.
This system provides reduction in the amount of coolant required,
an increase in turbine horsepower, and a reduction in the first
disk seal leakage.
The requirement for coolant is reduced due to reduction in cooling
air temperature. The turbine horsepower is increased since the disk
does not have to pump the coolant up to wheel speed to have it flow
through the holes in the disk. First disk seal leakage is reduced
because the pressure ratio across the seal is reduced.
The preswirl nozzle means accelerates the air flow and directs it
in a tangential direction approximately paralleling the rotation of
the first disk permitting the lower temperature and pressure.
Minimizing cooling and leakage air quantities by providing cooler
air at lower pressures allows more air to remain in the engine main
stream.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 is a sectional view taken through an engine casing showing
the diffuser section, burner section, and turbine section
incorporating the turbine coolant flow system.
FIG. 2 is an enlarged view taken along the line 2--2 of FIG. 1
showing the directional vanes which direct coolant flow on to the
first turbine disk.
FIG. 3 is a velocity diagram showing the relative velocity of the
air leaving the vanes and of the disk.
DESCRIPTION OF THE PREFERRED EMBODIMENT
As shown in FIG. 1 an annular diffuser 2 is shown discharging into
an annular burner section 4 which contains a plurality of burners 6
therearound. Compressor means (not shown) discharges into the
annular inlet of the diffuser 2, this can be from a compressor
section as shown in U.S. Pat. No. 3,301,526. The outer wall 8 of
the annular burner section is a continuation of the outer wall of
the diffuser 2. Pin means 10 are provided at a plurality of
locations around the engine where the outer diffuser wall and
burner wall 8 meet to fix the forward end of each burner 6 in
position. An arm 12 extends forwardly from each burner having
opening means through which a pin is positioned which is fixed to
the outer burner wall. The inner wall 14 of the burner section is a
continuation of the inner wall of the diffuser 2. This inner wall
14 has a cooperating wall 15 which forms an annular passageway 17
therewith extending from a point adjacent the outlet of the
diffuser to the forward part of the turbine section 11. Spacer
members 13 are used to properly space the walls 14 and 15.
The rearward end of the outer burner wall 8 is fixed to the outer
wall 16 of the turbine section 11. A plurality of vanes 18 extend
inwardly from the outer wall 16 and are connected at their inner
ends by a composite annular member 20. The composite annular member
20 is connected to the rear end of the inner wall 14 by a solid
wall member 28. The annular member 20 includes one labyrinth member
22 extending rearwardly as a cylindrical member and having inwardly
extending lands for a purpose to be here-in-after described. An
annular duct member 24 is positioned radially inwardly from
cylindrical member 22 and extends axially having a passageway 26
therethrough. Passageway 26 has a plurality of vanes 27 therein
which takes the air which is directed substantially axially thereto
and turns it a predetermined amount for a purpose to be
here-in-after described.
The forward part of the duct 24 is connected at its outer edge to
member 20 by an outwardly extending flange and is connected at its
inner edge to wall 15 by a solid wall member 30. The solid member
30 also has a second labyrinth member 32 extending rearwardly at a
position radially inwardly from the duct 24 as a cylindrical member
and having inwardly extending lands for a purpose to be
here-in-after described. Solid wall member 28 and solid wall member
30 form an annular connecting means between the rear end of
passageway 17 and the inlet to passageway 26.
The rear end of each burner is supported by flange members 3 and 5
which extend forwardly from the forward part of the turbine section
11. These flange members are adapted to engage cooperating members
fixed to the rearward part of each of the burner cans. While one
means has been shown, the burner can may be fixed in position by
any means desired.
Immediately downstream of vanes 18 is the first stage of a turbine
rotor comprising a rotor disk 40 with blades 42 mounted thereon.
The rotor disk 40 is mounted on a flange 41 which extends from a
shaft 43 which is mounted for rotation within said engine. Said
shaft being mounted for rotation and sealed where necessary by
conventional means. The blades may be attached to the rotor by any
means desired which permits a flow of fluid through the blade (as
shown by the arrows). The rotor disk 40 has a forward side plate 44
fixed thereto and to the blades, and a rearward side plate 46 fixed
thereto and to the blades. The forward side plate 44 has a circular
flange means 48 extending forwardly therefrom with its free end
being positioned between labyrinth member 22 and the outer wall of
the annular duct member 24. This free end has a sealing surface on
its outer side which mates with the lands on the member 22. A
similar flange 49 extends forwardly from the rotor disk 40 with one
side of its free end being positioned adjacent the lands of the
labyrinth member 32 so as to provide a sealing action at that
point. These flange members 48 and 49 are spaced from the annular
member 24 and form a rotating annular passageway 50
therebetween.
A plurality of holes 52 extend from the passageway 50 between the
connection of flange means 48 and 49 on the rotor disk 40 to the
space formed between the side plates 44 and 46. Holes 54 extend
from the passageway 50 between flanges 48 and 49 through the disk
to a point on the rearward side of the disk. Openings (not shown)
extend through the platform in each blade 42 to permit fluid flow
from the space between the side plates 44 and 46 to the interior of
the blades 42. Fluid in blades 42 is permitted to flow therefrom
through openings 56 located in the trailing edge of the blades.
Downstream of the first stage of the turbine rotor is the second
stage comprising a rotor disk 60 and blades 62. Disk 60 is also
fixed to the flange 41. A web member 64 interconnects the outer
edge of each of the rotor disks and has a plurality of lands
extending outwardly therefrom to provide a labyrinth seal for
sealing in a manner to be here-in-after described. An annular
chamber 70 is formed between the disks 40 and 60 and web member 64.
The rotor disk 60 has a forward side plate 66 fixed thereto and to
the blades, and a rearward side plate 68 is fixed thereto and to
the blades. A plurality of holes 72 extend between the chamber 70
and the space formed between the side plates 66 and 68. Openings
(not shown) extend through the platform in each blade 62 to permit
fluid flow from the space between the side plates 66 and 68 to the
interior of the blades 62. Fluid in blades 62 is permitted to flow
therefrom through openings located in the outer end of the
blades.
Vanes 80 extend inwardly from the outer wall 16 of turbine section
11 between blades 42 and 62. Said vanes are fixed at their outer
end and have a sealing member at their inner ends which cooperates
with the lands on the web member 64. Vanes 90 are positioned
downstream of the blades 62 and seal means 92 are provided between
the side plates 68 and inner structure of the vanes 90.
Openings 94 are located in flange 41 so that fluid bleeding by
labyrinth lands on member 32 can pass to the rear of turbine disk
60 and back into the flow path.
An annular manifold 100 is positioned around the forward part of
the burner section 4 with an annular opening 101 directed forwardly
so as to pick up air from the compressor discharge diffuser. This
location provides the coolest and cleanest mid-span air for
cooling. The annular manifold 100 has projections 102 extending
from the top thereof, each with an opening, to receive the pin from
pin means 10. The annular manifold has a plurality of thin hollow
struts 104 to connect the annular manifold to the forward end of
the annular passageway 17. An opening 105 in wall 14 opens the
interior of each hollow strut into passageway 17.
It can now be seen that air leaving a compressor located forwardly
of the diffuser 2 will flow through diffuser 2 into the burner
section 4. A predetermined amount of compressor air is picked up by
the annular manifold 100 and directed into annular passageway 17 by
the plurality of hollow struts 104 through holes 105. The flow then
passes through passageway 17 into the annular connecting means
formed by solid wall member 28 and solid wall member 30 and
directed to the inlet of passageway 26. The flow is then expanded
through the vanes 27 and exits from the passageway 26 in a
predetermined direction and at a desired velocity to be compatible
with the rotation of the turbine rotor disk 40. A portion of flow
then passes through holes 54 into annular chamber 70 and a portion
flows through openings 52 into the area between side plates 44 and
46 where it then flows through the blades 42. The air in chamber 70
then flows through openings 72 to the area between side plates 66
and 68 and then passes through blades 62.
Calculations made for one construction showed that flow losses were
low in the system and approximately 96 percent of compressor
discharged total pressure was supplied to the inlet of annular duct
member 24. The cooling air was then expanded through the full
annulus of vanes 27 to a pressure equal to approximately 65 percent
of compressor discharged total pressure. Actually the swirl
velocity imparted to the cooling air is greater than the speed of
the rotating passageway 50 and work is then done on the rotating
disk. In this expansion process, the gas temperature inside the
chamber 50 dropped approximately 120.degree.F. This cooling air
then progressed through the flow system. As stated here-in-before
seal leakage is minimized because the disk pressure level has been
dropped to approximately 65 percent of the compressor discharged
total pressure.
* * * * *