Fluid reaction surfaces (i.e. – impellers) – With heating – cooling or thermal insulation means – Changing state mass within or fluid flow through working...
Reexamination Certificate
2001-04-18
2002-11-05
Ryznic, John E. (Department: 3745)
Fluid reaction surfaces (i.e., impellers)
With heating, cooling or thermal insulation means
Changing state mass within or fluid flow through working...
Reexamination Certificate
active
06474947
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
The present invention relates to a film cooling hole structure of a gas turbine moving blade in which arrangement of film cooling holes is optimized so as to enhance a cooling efficiency of the moving blade.
2. Description of the Prior Art
In a gas turbine moving blade known in the art, cooling air is flown in a serpentine cooling passage provided in the blade for effecting a convection cooling, and also cooling air is injected from film cooling holes onto a blade outer surface for effecting a film cooling.
FIGS.
6
(
a
) and
6
(
b
)are cross sectional views of one example of a gas turbine moving blade cooling structure in the prior art, wherein FIG.
6
(
a
) shows an entire portion of the cooling structure and FIG.
6
(
b
) shows a cross sectional view taken on line B—B of FIG.
6
(
a
). In FIG.
6
(
a
), numeral
30
designates a moving blade, whose interior is sectioned by ribs
36
,
37
,
38
,
39
to form a leading edge side cooling passage
31
, a serpentine cooling passage comprising cooling passage portions
32
,
33
,
34
in a blade central portion, and a trailing edge side cooling passage
35
, when the passage portions
32
,
33
,
34
communicate with each other in this order.
Cooling air represented by arrows
40
in a blade base portion enters the cooling passages, wherein the cooling air flowing in the leading edge side cooling passage
31
cools a blade leading edge portion and flows out of leading edge side holes as represented by arrows air
40
a, the cooling air flowing in the cooling passage portions
32
,
33
,
34
cools the blade central portion and flows out of film cooling holes provided in a blade surface for effecting a film cooling of the blade surface as air represented by arrows
40
b,
and the cooling air flowing in the trailing edge side cooling passage
35
cools a blade trailing edge portion and flows out of a blade tip portion as represented by arrows air
40
c
and also flows out of a multiplicity of cooling holes provided in a blade trailing edge as air represented by arrows
40
d.
FIGS.
5
(
a
) and
5
(
b
) are cross sectional views of another example of a gas turbine moving blade cooling structure in the prior art, wherein FIG.
5
(
a
) shows an entire portion of the cooling structure and FIG.
5
(
b
) shows a cross sectional view taken on line A—A of FIG.
5
(
a
). In FIG.
5
(
a
), numeral
20
designates a moving blade, whose interior is sectioned to form a leading edge side cooling passage
21
, a serpentine cooling passage comprising cooling passage portions
22
,
23
,
24
, and a serpentine cooling passage comprising cooling passage portions
25
,
26
,
27
on a rear side-thereof, wherein the cooling passage portions
22
,
23
,
24
and
25
,
26
,
27
communicate with each other in this order, respectively.
Cooling air represented by arrows
41
in a blade base portion enters the cooling passages, wherein the cooling air entering passage (A) flows into the leading edge side cooling passage
21
and flows out of leading edge side holes as air represented by arrows
41
a,
the cooling air entering passage (B) flows into the cooling passage portion
22
to then flow through the cooling passage portions
23
,
24
and flows out of film cooling holes provided in a blade tip portion as air represented by arrows
40
b,
and the cooling air entering passages (C), (D) flows into the cooling passage portion
25
to then flow through the cooling passage portions
26
,
27
and flows out of a multiplicity of cooling holes of a blade trailing edge portion as air represented by arrows
41
d.
Thus, the blade is so constructed as to be cooled effectively in its entirety.
FIG. 4
is an enlarged explanatory view of portion x of FIG.
5
(
a
) showing a film cooling hole structure in a cooling passage turning portion of the gas turbine moving blade in the prior art. The cooling passage portions
22
,
23
are sectioned by a rib
51
and communicate with each other at a turning portion in the blade tip portion. In the blade tip portion, there are provided a multiplicity of film cooling holes
50
. When the cooling air represented by arrow
41
flowing in the cooling passage portion
22
flows into the adjacent cooling passage portion
23
sectioned by the rib
51
as shown by arrow
41
e,
it does not flow along the rib
51
in the turning portion but separates therefrom as shown by arrow
41
f,
which results in causing a separation area
52
where a heat transfer rate is reduced. Further, as shown by arrow
41
g,
there arises a stagnation area
53
in a corner of the cooling passage portion
22
, and the heat transfer rate is low in the stagnation area
53
also. Thus, there is caused a cooling non-uniformity in the cooling passage.
In the mentioned prior art gas turbine moving blades of FIGS.
5
(
a
),
5
(
b
),
6
(
a
) and
6
(
b
), there are provided the leading edge side cooling passage, the serpentine cooling passage of the blade central portion and the trailing edge side cooling passage and the cooling air is flown therethrough for blade cooling and the cooling air is also injected from the film cooling holes onto the blade outer surface for effecting a film cooling. However, the positions of the film cooling holes are not necessarily optimized, so that there arises the stagnation area of the cooling air in the cooling passage and also there is caused the separation phenomenon of the cooling air from the rib surface in the turning portion of the serpentine cooling passage. The stagnation area and separation area are areas where the heat transfer rate is reduced, thereby the cooling of the blade interior becomes non-uniform and this is one of the reasons for the cooling efficiency being reduced.
Thus, the present invention is made with a first object to provide a gas turbine moving blade cooling structure in which film cooling holes provided in a cooling passage are devised to be arranged so as to eliminate a stagnation area and a separation phenomenon of cooling air to thereby realize a uniform cooling in the cooling passage, and to enhance a cooling efficiency by eliminating an area where a heat transfer rate is low.
FIGS.
8
(
a
) and
8
(
b
) are cross sectional views of still another example of a gas turbine moving blade cooling structure in the prior art, wherein FIG.
8
(
a
) shows an entire portion of the cooling structure and FIG.
8
(
b
) shows a cross sectional view taken on line B—B of FIG.
8
(
a
). In FIG.
8
(
a
), numeral
30
designates a moving blade, whose interior is sectioned by ribs
36
,
37
,
38
,
39
to form a leading edge side cooling passage
31
, a serpentine cooling passage comprising cooling passage portions
32
,
33
,
34
in a blade central portion, and a trailing edge side cooling passage
35
, wherein the cooling passage portions
32
,
33
,
34
communicate with each other in this order. In each of these cooling passages, there are provided turbulators
48
for making a flow of cooling air therein turbulent to accelerate a convection to thereby enhance a heat transfer effect of the cooling air.
Cooling air represented by arrows
40
in a blade base portion enters the cooling passages, wherein the cooling air flowing in the leading edge side cooling passage
31
cools a blade leading edge portion and flows out of leading edge side holes as air represented by arrows
40
a,
the cooling air flowing in the cooling passage portions
32
,
33
,
34
cools the blade central portion and flows out of film cooling holes provided in a blade surface for effecting a film cooling of the blade surface as air represented by arrows
40
b,
and the cooling air flowing in the trailing edge side cooling passage
35
cools a blade trailing edge portion and flows out of a blade tip portion as air represented by arrows
40
c
and also flows out of a multiplicity of cooling holes provided in a blade trailing edge as air represented by arrows
40
d.
FIGS.
7
(
a
) and
7
(
b
) are cross sectional views of still another example of a gas turbine moving blade cooling s
Mitsubishi Heavy Industries Ltd.
Ryznic John E.
Wenderoth , Lind & Ponack, L.L.P.
LandOfFree
Film cooling hole construction in gas turbine moving-vanes does not yet have a rating. At this time, there are no reviews or comments for this patent.
If you have personal experience with Film cooling hole construction in gas turbine moving-vanes, we encourage you to share that experience with our LandOfFree.com community. Your opinion is very important and Film cooling hole construction in gas turbine moving-vanes will most certainly appreciate the feedback.
Profile ID: LFUS-PAI-O-2968835