Patent Application
20120195766
The invention relates to turbine engine blades made of ceramic matrix composite (CMC) material, i.e. comprising fiber reinforcement that is densified with a ceramic matrix.
The intended field is that of gas turbine blades for aeroengines or for industrial turbines, and more particularly the blades of rotor wheels of compressor or turbine stages, such as the blades of a low pressure (LP) turbine.
Proposals have already been made to produce turbine engine blades out of CMC. Reference may be made in particular to the following documents: US 2011/0293828, US 2011/0311368, and US 2008/0229567.
Turbine engine blades are exposed to excitations that are produced by their environment, in particular the wake effect from nozzles (stationary vanes) or by unbalance. These excitations give rise to vibratory stresses that it is desirable to damp in order to avoid a blade breaking.
With wheels that have metal blades, it is known to damp vibration between adjacent blades by means of parts that are engaged freely in facing housings formed in the edges of outer platforms of the blades, the parts being pressed against the walls of the housings by centrifugal force in order to damp vibration. Reference may be made to document U.S. Pat. No. 3,752,599.
Still in the context of wheels having metal blades, it is known to provide damping by pre-twisting the blades about axes extending substantially in the longitudinal directions of the blades, i.e. axes that are substantially radial relative to the axis of the wheel. The twisting prestress is maintained by mutual engagement between the outer platforms of adjacent blades in the wheel, via contact zones that are situated on opposite sides of the outer platforms in the circumferential direction. A coating of wear-resistant material is then conventionally applied to the contact zones by welding, e.g. a coating of the material sold under the name Stellite®.
Such a solution cannot be envisaged for blades made of CMC because of the difficulty firstly in binding a metal coating to a CMC material, and secondly of guaranteeing that said binding is long-lasting, given the difference between the coefficients of thermal expansion of the metal coating and of the CMC material.
An object of the invention is to provide a solution to the problem of damping the vibratory stresses to which the blades of rotor wheels of a turbine engine are subjected.
In a first aspect, the invention provides a turbine engine rotor wheel including a plurality of blades of composite material comprising fiber reinforcement densified with a ceramic matrix, each blade having a first portion constituting a blade airfoil and root and being formed as a single piece together with at least one second portion forming an outer platform,
in which wheel the blades are prestressed in twisting about longitudinal axes and are held under prestress by mutual engagement via contact zones between the outer platforms of adjacent blades, the contact zones situated on opposite sides of the outer platform of a blade being defined by at least one insert integrated in the outer platform.
The term “insert integrated in the outer platform” is used herein to mean an insert that is made integrally with the outer platform.
The invention is remarkable in particular by inserts being integrated in the outer platforms of CMC blades in order to define mutual contact zones between the outer platforms of blades that are mounted with twisting prestress in the rotor wheel. This makes it possible for a material to be selected for these zones in contact that is other than a CMC material and that presents good resistance to wear by friction, and that also advantageously presents good behavior at high temperatures and good resistance to oxidation.
The inserts may be made of a carbon-based material, in particular of a material selected from monolithic graphite and a carbon/carbon (C/C) type composite material, such materials having not only good resistance to wear but also a coefficient of thermal expansion that is close to that of CMC type materials.
Each outer platform may be provided with two inserts arranged on opposite sides of the outer platform, or with a single insert extending from one side of the outer platform to the other.
The twisting prestress angle that is applied to each of the blades is preferably less than 7°, or indeed less than 5°, which angle may be greater when the longitudinal size of the blade is large.
In another aspect, the invention provides a turbine engine including at least one rotor wheel as defined above.
In yet another aspect, the invention provides a turbine engine blade made of composite material comprising fiber reinforcement densified with a ceramic matrix, the blade having a first portion constituting a blade airfoil and root and being formed as a single piece together with at least one second portion forming an outer platform, in which blade the outer platform presents contact zones on opposite sides for coming into contact with the outer platforms of adjacent blades when the blade is integrated in a rotor wheel of a turbine engine, and the contact zones are defined by at least one insert integrated in the outer platform.
The or each insert is made of a carbon-based material, in particular essentially a material selected from monolithic graphite and a C/C type composite material.
The outer platform may be provided with two inserts arranged on either side of the outer platform or with a single insert extending from one side of the outer platform to the other.
In yet another aspect, the invention provides a method of fabricating a turbine engine blade as defined above, the method comprising:
in which method the or each insert defining a contact zone is inserted into the portion of the fiber blank that forms the second preform portion.
The or each insert may for example be inserted in a zone of non-interlinking between two layers of three-dimensional weaving.
Other features of the invention appear on reading the following description made by way of non-limiting indication and with reference to the accompanying drawings, in which:
The invention is applicable to various types of turbine engine blade, in particular compressor blades or turbine blades for various spools of a gas turbine, e.g. a blade for a rotor wheel of an LP turbine, such as the blade 10 of
In well-known manner, the blade 10 comprises an airfoil 20, a root 30 formed by a portion of greater thickness, e.g. presenting a bulb-shaped section, and extended by a tang 32, an inner platform 40 situated between the tang 30 and the airfoil 20 and an outer platform 50 in the vicinity of the free end of the airfoil 20.
The airfoil 20 extends in a longitudinal direction between the inner platform 40 and the outer platform 50, and in cross-section it presents a curved profile of varying thickness between its leading edge 20a and its trailing edge 20b. The root 30 is extended by the tang 32 so as to connect to the inner (or bottom) face of the inner platform 40.
At its radially inner end, the airfoil 20 is connected to the outer (or top) face 42 of the inner platform 40 that serves to define the inside of the flow passage for the gas stream through the turbine. The lower platform is terminated by overlap nibs 44 and 46. In the example shown, the face 42 of the lower platform is sloping, so as to form generally a non-zero angle α relative to the normal to the longitudinal direction of the blade. Depending on the profile desired for the inside surface of the gas stream flow passage, the angle α may be zero, or the face 42 may have a profile that is generally non-rectilinear, e.g. that is curved.
At its outer radial end, the airfoil 20 is connected to the outer platform 50 via an inner (or bottom) face 52 of the outer platform that defines the outside of the flow passage for the gas stream. The outer platform defines a depression or bathtub 58. Along the upstream and downstream edges of the bathtub 58, the outer platform carries wipers 60 with ends that are suitable for penetrating into a layer of abradable material carried by a turbine ring (not shown) for reducing the clearance between the tip of the blade and the turbine ring. In the example shown, the face 52 of the outer platform extends substantially perpendicularly to the longitudinal direction of the blade. In a variant, and depending on the profile desired for the outer surface of the flow passage for the gas stream, the face 52 may be sloping so as to form generally a non-zero angle relative to the normal to the longitudinal direction of the blade, or the face 52 may have a profile that is generally non-rectilinear, e.g. that is curved.
In the example shown, the airfoil 20, the root 30, the inner platform 40, and the outer platform 50 are made as a single piece of CMC material, with the exception of inserts that are integrated in the outer platform and that are made of a different material, as explained below.
The rotor wheel 60 of the LP turbine is built up by assembling blades 10 on a turbine disk 62 (
The blades 10 are mounted with prestress in twisting, or pre-twisting, along respective axes that extend substantially in the longitudinal directions of the blades, i.e. axes that extend radially relative within the wheel 60. The blades are held under twisting prestress by the outer platforms 50 of adjacent blades engaging mutually via contact zones that are situated on the opposite sides 50a and 50b of the outer platforms, i.e. sides that are opposite in the circumferential direction of the wheel 60.
In the example shown, the mutual engagement of the outer platforms 50 is achieved with locking by means of complementary shapes in relief that are formed in or on the sides 50a and 50b, e.g. a notch 52a and a projecting portion 52b that are both substantially V-shaped and situated substantially in the middle portions of the sides 50a and 50b. Other complementary shapes in relief could naturally be envisaged.
The contact zones in the notches 52a and the projecting portions 52b are defined by respective inserts 54a and 54b that are integrated in the outer platform 50 (see
As shown in
The outer portions 542a and 542b of the inserts 54a and 54b define the flanks 520a and 520b and possibly also adjacent portions of the sides 50a and 50b, in particular at the bottom of the notch 52a and at the tip of the projecting portion 52b, with the remainder of the sides 50a and 50b being defined by the CMC material of the outer platform 50.
The CMC material of the blade 10 is formed by fiber reinforcement that is densified with a ceramic matrix. The reinforcement is formed of carbon fibers or of ceramic fibers, in particular carbon fibers and/or refractory oxide fibers such as fibers of silicon carbide (SiC) or essentially of SiC. The ceramic matrix may be made of a refractory carbide, nitride, or oxide, e.g. SiC. The twisting prestress needs to remain within the elastic deformation range of the CMC material. Given the usual mechanical properties of this type of material, a twisting angle of less than 7°, or indeed of less than 5° is preferable for an LP turbine blade of an airplane turbojet, in particular depending on the longitudinal dimension of the blade, as mentioned above.
The material of the inserts 54a and 54b is selected from materials that present better resistance to friction wear than does the CMC material of the blade, while nevertheless being compatible therewith, i.e. having a coefficient of thermal expansion that is substantially equal to or close to that of the CMC material and that is suitable for being inserted during the process of fabricating the blade. For the inserts 54a and 54b, it is possible to select in particular a carbon-based material such as monolithic graphite or a composite material of the carbon-carbon (C/C) type, i.e. comprising fiber reinforcement made of carbon fibers and densified by a carbon matrix, possibly including ceramic particles dispersed in the matrix and forming a minority fraction of the matrix. Such C/C materials are well known in the friction field, in particular for airplane brake disks.
The inserts 54a and 54b may be cut by being machined from a block of graphite or of C/C type composite material. A block of C/C type composite material may be obtained by superposing plies of a fabric of carbon-precursor fibers, e.g. fibers of preoxidized polyacrylonitrile (PAN) and by bonding them together, e.g. by needling, and then transforming the precursor into carbon by heat treatment in order to obtain a carbon fiber preform, and then densifying the preform with a carbon matrix. The densification may be performed by a liquid technique or by a gaseous technique. The liquid technique consists in impregnating the preform with a carbon-precursor resin and then transforming the resin into carbon by curing and by pyrolysis. The gaseous technique consists in depositing a matrix of pyrolytic carbon by chemical vapor infiltration (CVI) by using a gas containing at least one carbon-precursor gas such as methane or propane. These densification methods are well known.
The CMC blade 10 is obtained by a process comprising making a fiber blank by three-dimensional (or multilayer) weaving, by shaping the fiber blank in tooling in order to obtain a fiber preform constituting the fiber reinforcement of the CMC material, and by densifying the fiber preform with a ceramic matrix. The prefabricated inserts 54a and 54b are inserted in the stage of making the fiber preform prior to the preform being densified. To this end, zones of non-interlinking are provided, e.g. while weaving, between two layers of yarns in the portion of the preform that corresponds to the outer platform in order to be able to insert the portions 540a and 540b of the inserts 54a and 54b.
A process for obtaining a blade 10 is described with reference to
The blank 100 comprises three portions 102, 104, and 106 that are obtained by three-dimensional weaving, and only the envelopes of those three portions are shown in
The three portions 102, 104, and 106 are in the form of strips that extend generally in a direction X that corresponds to the longitudinal direction of the blade that is to be made. In its portion that is to form an airfoil preform, the fiber strip 102 presents thickness that varies in determined manner as a function of the thickness of the profile of the airfoil of the blade that is to be made. In its portion that is to form a root preform, the fiber strip 102 presents extra thickness 103 that is determined as a function of the thickness of the root of the blade that is to be made.
Various modes of three-dimensional weaving are described in particular in document US 2010/0144227, the content of which is integrated herein by reference. The variation in the thickness of the strip 102 in its portion that is to form an airfoil preform may be obtained by varying the numbers of layers of yarns, while the extra thickness in the portion that is to form a root preform may be obtained by inserting an insert.
The fiber strip 102 is of width l selected as a function of the length of the (flat) developed profile of the airfoil and of the root of the blade that is to be made, whereas each of the fiber strips 104 and 106 presents a width L greater than l that is selected as a function of the developed lengths of the platforms of the blade that is to be made.
The fiber strips 104 and 106 are of substantially the same width, and each of them is of a substantially constant thickness that is determined as a function of the thicknesses of the platforms of the blade that is to be made. Each of the strips 104 and 106 comprises a first portion 104b or 106b that extends along and beside a first face 102b of the strip 102, a second portion 104a or 106a that extends along and beside the second face 102a of the strip 102, and a third portion 105b or 107b that extends along and beside the first face 102b of the strip 102.
The portions 104b and 104a of the strip 104 are connected together by a connection portion 140c that extends transversely relative to the strip 102 at a location corresponding to the location of the inner platform of the blade that is to be made. The connection portion 140c passes through the strip, making an angle α relative to the normal to the longitudinal direction of the fiber blank. Similarly, the portions 106b and 106a of the strip 106 are connected together by a connecting portion 160c that extends transversely relative to the strip 102 and that is substantially parallel to the connection portion 140c (and might possibly be spaced apart therefrom).
The portions 104a and 105b of the strip 104 are connected together by a connection portion 150c that extends transversely relative to the strip 102 at a location corresponding to that of the outer platform of the blade that is to be made. In the example shown, the connection portion 150c passes through the strip 102 substantially perpendicularly to the longitudinal direction X of the fiber blank. Similarly, the portions 106a and 107b of the strip 106 are connected together by a connection portion 155c that extends transversely relative to the strip 102 and that is substantially parallel and adjacent to the connection strip 150c.
Depending on the shape desired for the outer platform of the blade, the connection portions 150c and 155c may pass through the strip 102 while making a non-zero angle relative to the normal to the longitudinal direction X of the blank, as for the inner platform. In addition, the connection portions 140c, 160c, and/or 150c and 155c may be of profiles that are curvilinear instead of being rectilinear as in the example shown.
The strips 102, 104, and 106 are woven simultaneously by three-dimensional weaving without interlinking firstly between the strip 102 and the portions 104b, 104a and 105b of the strip 104, and secondly between the strip 102 and the portions 106b, 106a and 107b of the strip 106. Advantageously, a plurality of successive blanks 100 are woven continuously in the X direction. Similarly, there is no interlinking between the strips 104 and 106.
The fiber strip 102 is cut at one end in the extra thickness 103 and at another end a little beyond the connection portions 150c and 155c so as to provide a strip 120 of length that corresponds to the longitudinal dimension of the blade that is to be fabricated with an enlarged portion 130 formed by a portion of the extra thickness 103 that is situated at a location corresponding to the position of the root of the blade that is to be fabricated.
Furthermore, cuts are made at the ends of the portions 104b and 105b of the strip 104, at the ends of the portions 106b and 107b of the strip 106, and in the portions 104a and 106a thereof so as to leave segments 140a and 140b on either side of the connection portions 140c and 160c, and also segments 150a and 150b on either side of the connection portions 150c and 155c, as can be seen in
Because of the non-interlinking between the strip 102 and the portions 104b, 104a and 105b of the strip 104, and also between the strip 102 and the portions 106b, 106a and 107b of the strip 106, the segments 140a, 140b, 150a, and 150b can be folded out perpendicularly relative to the strip 102 without cutting yarns so as to form plates 140 and 150, as shown in
The inserts 54a and 54b are put into place by being inserted between the portions 151 and 152 of the strips 104 and 106 that form the plate 150 and that are not interlinked by weaving. Beforehand, cuts are made in the sides 150a and 150b of the plate 150 that are to form the sides 50a and 50b of the outer platform, so as to give them the desired shapes in order to obtain the complementary shapes in relief that are desired for these sides 50a and 50b (
A strip preform 200 of the blade that is to be fabricated is subsequently obtained by molding, with the strip 102 being deformed so as to reproduce the curved profile of the airfoil of the blade (
In a variant, instead of the strips 104 and 106, it is possible to use a single strip, e.g. 104, by making provision while weaving the strip 104 for zones of non-interlinking so as to make it possible to form the portion of the preform corresponding to the wipers of the outer platform and so as to make it possible to insert the wedge-shaped portions of the inserts 54a and 54b.
The fiber preform may be densified with a ceramic material in known manner. A first consolidation step with a first matrix phase may be performed using a resin that is a precursor of the ceramic consolidation matrix. The preform may be impregnated with resin prior to being shaped, e.g. during the step of
In the embodiment of
In another variant, a single insert may be provided that extends from one side of the outer platform to the other by being inserted between the two layers forming the plate 150.
The description above relates to a blade made of CMC material with integral platforms constituting a single piece together with the airfoil and the root. In a variant, the blade may be made with an outer platform that forms a single piece together with the blade and the root, while the inner platform is fitted separately thereto.
| Number | Date | Country | Kind |
|---|---|---|---|
| 11 50821 | Feb 2011 | FR | national |
