U.S. patent number 3,732,031 [Application Number 05/047,046] was granted by the patent office on 1973-05-08 for cooled airfoil.
This patent grant is currently assigned to General Motors Corporation. Invention is credited to Charles E. Bowling, George B. Meginnis, Ronald P. Schwedland.
United States Patent |
3,732,031 |
Bowling , et al. |
May 8, 1973 |
COOLED AIRFOIL
Abstract
A cooled turbine blade or vane for high-temperature machines. To
obtain maximum strength and oxidation resistance and the best
cooling characteristics, the airfoil has a core of cast super alloy
covered by a porous facing of wrought super alloy sheet, the core
providing passages to transmit the cooling fluid to the facing. The
effective cooling and resistance to corrosion are due primarily to
the facing and the strength and resistance to centrifugal,
buffeting, or ballooning loads are contributed primarily by the
cast core.
Inventors: |
Bowling; Charles E. (Speedway,
IN), Meginnis; George B. (Indianapolis, IN), Schwedland;
Ronald P. (Indianapolis, IN) |
Assignee: |
General Motors Corporation
(Detroit, MI)
|
Family
ID: |
21946780 |
Appl.
No.: |
05/047,046 |
Filed: |
June 17, 1970 |
Current U.S.
Class: |
416/97R; 416/228;
29/889.721; 416/96A; 416/231R; 416/193R |
Current CPC
Class: |
F01D
5/184 (20130101); Y10T 29/49341 (20150115) |
Current International
Class: |
F01D
5/18 (20060101); F01d 005/18 () |
Field of
Search: |
;416/90,95,97,226,229,231,232 |
References Cited
[Referenced By]
U.S. Patent Documents
Primary Examiner: Feinberg; Samuel
Claims
We claim
1. A cooled flow-directing member for use in a high-temperature
turbomachine comprising, in combination, a rigid hollow core of
airfoil configuration and a porous facing of corresponding airfoil
configuration bonded to and covering the core and defining the
exterior of the member, the core having distributed perforations to
conduct cooling fluid from within the core to the facing, the core
having a smooth outer surface and the facing having an inner
surface abutting the core outer surface, the said facing bearing a
two-dimensional array of spaced bosses extending from its inner
surface abutting the outer surface of the core so as to provide
passages for flow of the cooling fluid between the bosses generally
parallel to the said surfaces, the perforations communicating with
the said passages, the core reinforcing and supporting the facing
against loads imposed on the facing and the facing being effective
to shield the core from the working fluid of the turbomachine and
to cool the core, the facing defining pores distributed over a
substantial portion of the area of the airfoil out of register with
the said perforations, the said pores connecting the said passages
to the exterior of the facing for transpiration cooling of the
member, the core being of a cast super alloy having greater
strength than the facing at high temperature and the facing being
of a ductile wrought super alloy sheet having greater resistance to
corrosive attack by high temperature gases than the core.
2. A member as defined in claim 1 in which the facing is a
laminated structure of more than one layer.
3. A member as defined in claim 1 in which the core and facing are
of nickel base or cobalt base alloys.
Description
The invention herein described was made in the course of work under
a contract or subcontract thereunder with the Department of
Defense.
DESCRIPTION
Our invention relates to cooled airfoils, particularly such as are
used as vanes or blades in high temperature turbomachines. Our
invention is primarily intended for high temperature gas turbines
but is applicable to other environments where airfoils having a
high degree of resistance to very hot motive fluids and a
considerable degree of strength are required.
In the preferred embodiment, a vane or blade (hereafter referred to
as an "airfoil") comprises a hollow cast airfoil cross-section core
and a formed porous sheet metal facing covering and bonded to the
core. A cooling fluid, usually air, is supplied through
perforations in the core to the interface between core and facing,
one of which is formed to provide passages generally parallel to
the outer surface of the airfoil. The facing defines pores through
which the cooling fluid emerges from the surface of the airfoil for
transpiration cooling.
By virtue of the principles of the invention, the strength
characteristics of cast super alloys employed for the core or
backbone of the airfoil give it the requisite strength. The
superior adaptation to cooling, formability, and resistance to
oxidation of the wrought super alloy sheet used for the surface of
the airfoil give better cooling and oxidation resistance than can
be obtained in an all-cast structure.
The outer facing made of thin sheets with the circulation of air
through and from the sheets provides very effective and efficient
cooling of a degree of efficiency which cannot be approached in the
present state of the art with a cast blade lacking the facing. On
the other hand, the materials which are available for the facing do
not have the high strength at high temperatures of the materials
specified for the core. In prior practice, with a porous sheet
metal blade, the maximum temperature of the outer layer has been
about 1,650.degree.F. By providing the cast core, the temperature
of the facing may be as high as 2,100.degree.F. at the outer
surface. This makes possible effective and efficient cooling even
with cooling air at a temperature of around 1,300.degree.F., which
is the level of cooling air available under some conditions in
modern gas turbine engines. By the more efficient cooling and the
higher temperature level of the outer surface, the amount of
cooling air can be approximately halved with respect to prior
technology involving either a cast blade or a formed sheet metal
transpiration cooled blade or vane. Viewed another way, the
improved cooling and resistance to temperature are capable of
providing a turbine blade usable in a gas turbine operating at the
very high motive fluid temperatures (about 3,700.degree.F. turbine
inlet and about 3,300.degree. at the first rotor stage) resulting
from stoichiometric combustion of JP fuels. It will be appreciated
by those skilled in the art how significant these facts are and it
will also be appreciated from the succeeding detailed description
that this advance in the art is made possible by a structure which
is feasible to fabricate in the present state of the art.
The principal objects of the invention are to improve the
efficiency of turbomachines; to improve the temperature tolerance
and life at high temperature of turbomachinery; to provide airfoils
for turbomachines which have superior cooling properties; to
provide a highly satisfactory cooled airfoil structure which is
feasible from the standpoint of manufacture; and to provide an
airfoil such as a turbine vane or blade having a porous sheet metal
facing and a perforated hollow cast metal core bonded to the
facing.
More specifically, the primary objective of the invention is to
comb the high temperature strength of a cast super alloy as a
supporting core with the extremely high cooling efficiency of a
laminated porous material outer skin. Thus, one can use the
inherent 200.degree. to 300.degree.F. temperature gradient which is
found across the wall from outside to inside to advantage. Allowing
the inner fibers of the blade (cast core) to operate at their
stress rupture limit (approximately 1,800.degree.F.) and letting
the outside sheet operate at its oxidation limit (approximately
2,100.degree.F.) one can achieve a maximum allowable temperature at
around 1,950.degree.F. which is 100.degree. to 150.degree. better
than the present day state of the art.
The nature of our invention and its advantages will be better
appreciated by reference to the succeeding detailed description of
preferred embodiments of the invention and the accompanying
drawings thereof.
Referring to the drawings, FIG. 1 is an axonometric view of a
turbine blade of known overall configuration.
FIG. 2 is a sectional view of the same taken on the spanwise
extending plane indicated by the line 2--2 in FIG. 1.
FIG. 3 is a transverse section of a turbine airfoil.
FIG. 4 is an enlarged view of the leading edge portion of FIG.
3.
FIG. 5 is a sectional view similar to FIG. 4 of a modified form of
the invention.
FIG. 6 is a fragmentary cross sectional view of a still further
form of the invention.
Referring now to the drawings, the rotor blade 2 illustrated in
FIGS. 1 and 2 includes an airfoil or flow-directing portion 3, a
platform 4, a stalk 6, and a root 7 of multiple dovetail
configuration. Such a blade is mounted by the root in a mating slot
in the turbine rotor rim and the platforms 4 of adjacent blades
meet to define an annular inner boundary of the motive fluid flow
path through the turbine rotor stage. The airfoil may be of any
suitable cross section; FIG. 3 illustrates a typical section of the
type normally employed in turbomachinery. It may have a leading
edge 8, a trailing edge 10, a concave face 11, and a convex face
12. As will be particularly apparent from FIG. 3, the airfoil
(blade or vane) is a composite structure made up of a core 14
overlaid by a facing 15 which covers the core, the facing being a
laminated structure made up of an outer layer 16 and an inner layer
18. The core is hollow so as to define a spanwise extending air
passage 19 through the airfoil, this passage connecting with a
passage 20 in the blade stalk which is supplied from externally of
the blade through any suitable opening, as is well known to those
skilled in the art. The core 14 is a casting of a nickel base super
alloy, such alloys being known to those skilled in the art under
such trade names as Mar M 246, Udimet 710, and Inco 713C. These are
high strength materials well adapted to support the loads from
blade stresses. In the structure illustrated, the platform, stalk
and root portion of the blade is bicast to the airfoil portion 3,
the airfoil portion having a ribbed base 22 which is interlocked
with the platform and stalk when the latter is cast around the
airfoil.
The outer end of the blade core as cast has holes 24 which provide
for core location and removal of the core by leaching. After this
is done, the holes are closed in any suitable way. As illustrated,
a plug 23 in the form of a rivet is inserted and is retained by
flaring out the outer end of the plug. At the trailing edge 10 of
the airfoil, the air passage 19 discharges through a narrow slot 26
extending generally from end to end of the airfoil. Adjacent the
trailing edge, the two faces of the cast core 14 are close together
and are joined by unitary cast pins or spacers 27.
The blade, including the pins 27, may be cast by known techniques
of the type described in British Pat. No. 872,705 of Hamilton L.
McCormick, published July 12, 1961. Preferably, the leading and
trailing edge portions of the core are substantially a single
crystal as described in U.S. Pat. No. 3,008,855 of Swenson, Nov.
14, 1961.
Referring now to FIGS. 3 and 4 for a more detailed explanation of
one variety of blade structure according to our invention, it will
be seen that the core 14 defines a multitude of perforations 28
disposed preferably in a generally rectangular two-dimensional
array over the major part of the concave face 11 and roughly the
first half of the chord of the convex face 12 in the specimen
illustrated. These perforations may be formed as part of the
airfoil casting process or may be formed by machining, as desired.
The outer surface of the core 14 may be formed with a surface
relief so that the facing is partly spaced from the core, this
relief being preferably in the form of a two-dimensional array of
bosses 30 on the outer surface of the core distributed over the
area through which the perforations 28 extend. These bosses might
be cast on the core or could be the result of some machining
process such as photochemical etching to cut away the core surface
to a desired depth between the bosses 30, but preferably the core
outer surface would be left smooth and the spacers would be etched
on the facing.
In the structure of FIGS. 3 and 4 the facing 15, as previously
pointed out, comprises an outer layer 16 and an inner layer 18. The
inner layer 18 is formed with the surface relief on its outer
surface by a two-dimensional array of bosses 31 and with an array
of pores 32 through the layer. The outer layer also has a
two-dimensional array of pores identified as 34. The pores 34 are
out of register with pores 32, which in turn are out of register
with perforations 28. Because of the arrangement of pores and the
surface relief at each interface between the core and layer 18 and
between layer 18 and layer 16, cooling air can flow from the air
passage 19 through perforations 28, pores 32, and pores 34 to the
exterior surface of the airfoil. This provides for transpiration
cooling of a considerable part of the surface of the airfoil. The
trailing edge portion may be cooled through convection by cooling
air escaping from passage 19 through the slot 26 scrubbing the
inside surfaces of the core and the spacers 27.
While transpiration cooling might be employed at the leading edge
of the blade, it is preferred to adopt an impingement mode of
cooling involving a modification of the facing structure. As shown
clearly in FIG. 4, the inner layer 18 is removed locally to provide
a series of chordwise channels adjacent the leading edge of the
blade and over the forward portion of the concave face 11 so as to
define a series of cooling air ducts 35 extending from the forward
edge of the convex face around to the edge 36 of the inner layer on
the concave face of the airfoil. The cooling air is delivered from
the passage 19 through a spanwise-extending row of nozzles 38
defined by holes through the core which direct the cooling fluid
against the leading edge and causes it to flow between the outer
layer and the core through the channels 35 to a row of outlets 39
adjacent the edge 36. The direction of flow is such as particularly
to scour the leading edge portion of the facing and increase the
effectiveness of convection cooling of the facing at this
point.
Referring particularly to FIG. 3, it may be seen that the facing 15
tapers toward the trailing edge of the blade past the perforations
28 and this part of the facing may be imperforate. Relative
distribution of the cooling air to various areas of the facing may
be controlled by size and spacing of the pores and perforations and
the depth of surface relief.
While the facing as shown comprises two layers, it could comprise
more, and could consist of a single layer, with other means than
that shown in FIG. 4 employed for cooling the leading edge. As will
be apparent, if the outer layer 16 is omitted, cooling air may flow
from the perforations 28 under the layer 18 and out the pores 32
for transpiration cooling. Thus it is possible to have a single
outer protective layer over the core.
FIG. 5 illustrates a modification of the cooled airfoil of the
previously described figures. It may be the same as the structure
shown in FIGS. 3 and 4 except for the changes noted below. The core
14' has a smooth outer surface and the relief between the core and
the facing 15' is provided by photoetching or otherwise relieving
the inner surface of the inner facing layer 18' to provide bosses
40 on this layer. In connection with this, it is also preferred to
provide the relief on the inner surface of the outer layer 16'
rather than the outer surface of the inner layer of the facing.
There is an advantage to this modification in that it is not
necessary to relieve the surface of the core and, assuming the
perforations and nozzles in the core are provided during the
casting operation, the photoetching or other machining may be
accomplished only on the layers or layer of the facing.
FIG. 6 is an enlarged sectional view comparable with those of FIGS.
4 and 5, except showing a portion of the concave face of a still
further modification of the cooled airfoil of our invention.
The structure of FIG. 6 involves a core of the same nature as that
illustrated in FIG. 4 but a facing which embodies the principle
described and claimed in the copending patent applications, of
common ownership with this application, of Thomas H. Mayeda, Ser.
No. 879,094 [U.S. Pat. No. 3,700,418] and George B. Meginnis, Ser.
No. 879,110, both filed Nov. 24, 1969. The particular structure
illustrated is similar to one illustrated in the Meginnis
application. The facing is such as to cause the air flowing from
the facing to flow at an acute angle to the face of the blade in a
direction downstream with respect to the flow of motive fluid so as
to minimize interference between the cooling air flow and the
motive fluid flow. Specifically, the core 14 is overlaid by an
inner layer 42, an intermediate layer 43, and an outer layer 44,
these being bonded together and bonded to the core. The inner layer
may have the same configuration as the inner layer 18 described in
connection with FIGS. 3 and 4. The intermediate layer 43 is a thin
sheet having no surface relief but having pores 46 communicating
with the air space defined by the relief on the face of the inner
layer 42. The outer layer 44 has pores defined by intersecting
staggered pits 47 on the inner surface of the layer and 48 on the
outer surface of the layer disposed so that the flow through the
pores 46 is deflected in a downstream direction; that is, upwardly
as illustrated in FIG. 6.
In general, it is desired that the facing in any of the forms
described be relatively thin and, in the case of a typical turbine
airfoil having a chord of about one or two inches, it is
contemplated that the facing have a total thickness of about
fifteen to twenty thousandths of an inch. Suitable materials for
the facing include various wrought high temperature resistant
nickel alloy sheets, specifically, materials known as Haynes alloy
188 and Hastelloy X. The layers of the facing may be bonded
together by processes of diffusion bonding after the surface relief
and pores have been machined by photoetching or otherwise. The
facing is preferably diffusion bonded to the core, which may be
accomplished by heat and pressure exerted through a suitable pad to
allow the pressure to be distributed over the curved face of the
core during the bonding operation. The bonding operation may take
place in connection with a creep forming operation by which the
contour of the facing is completed. It is preferred to effect the
diffusion bonds between facing layers and between facing and core
simultaneously. It should be understood that other modes of
attachment of the facing to the core are contemplated but that in
general welding is not suitable in the present state of the art for
the materials specified and brazing is not regarded as being as
satisfactory as diffusion bonding. Explosive welding might be
feasible. The core is ordinarily cleaned and polished and nickel
plated before diffusion bonding the facing to it.
While reference has been made to suitable known commercial high
temperature alloys in the preceding specification, it is to be
understood that the invention is applicable to such alloys as have
the desired properties, either present or as discovered in the
future. The principal characteristic of the cast material is high
strength at high temperatures to resist centrifugal, side, or
ballooning loads, and the principal characteristics of the facing
are its ductility and its resistance to oxidation and sulfidation
at extremely high temperatures, the highest temperature in the
airfoil being at the surface.
In a rotor blade, the facing will be running in compression and
will unload all its weight onto the load-carrying core because of
the greater thermal expansion of the facing resulting from its
higher temperature.
The detailed description of preferred embodiments of the invention
for the purpose of explaining the principles thereof is not to be
considered as limiting or restricting the invention, as many
modifications may be made by the exercise of skill in the art.
* * * * *