US 3754839 A
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United States Patent 1191 Bodman 1451 Aug. 28, 1973 F ILAMENT REINFORCED ROTOR ASSEMBLY  Inventor: Robert R. Bodman, Riviera Beach,
 Assignee: United Aircraft Corporation, East Hartford, Conn.
 Filed: May 1, 1972  Appl. No.: 249,336
 US. Cl. 416/195, 416/218, 416/230, 416/241, 416/190  Int. Cl. F04d 29/26, FOld 5/24  Field of Search 416/218, 189, 230, 416/241 A, 244 A, 195, 190
 References Cited UNITED STATES PATENTS 6/1963 Wamken 416/244 x 7/1968 Blackhurst et 31.. 416/218 x 10/1968 Stoffer 416/230 3/1970 Stoffer et 81.... 416/230 x 1/1971 HOWfilCl et a1 416/230 x 3.625.634 12/1971 Stedfeld 416/230 X 3,279,967 10/1966 Martin et al. 416/189 L'X 3.601.500 8/1971 Palfreyman 416/190 3,656,864 11/1970 Wagle 416/190 FOREIGN PATENTS OR APPLICATIONS 223,227 12/1958 Australia 416/190 Primary Examiner-Everette A. Powell, Jr. Attorney-Charles A. Warren [5 7 ABSTRACT This invention relates to a rotor assembly reinforced with a filament wound composite ring. The rotor assembly comprises a rotor including a plurality of radially extending circumferentially spaced blades. The composite ring surrounds the blade tips and is closely spaced therefrom by radially flexible support means. During rotor operation the blade tips grow radially outwardly and come into centrifugal load carry relationship to the composite ring, whereupon the ring carries a portion of the centrifugal loads thus reducing the strength requirements of the blades and rotor.
8 Claims, 2 Drawing Figures FILAMENT REINFORCED ROTOR ASSEMBLY BACKGROUND OF THE INVENTION 1. Field of Invention This invention relates to the use of circumferentially wound filaments to reinforce a rotor assembly.
2. Description of the Prior Art I The use of circumferentially wound filaments to reinforce a rotor assembly is well known in the prior art as evidenced by US. Pat. No. 3,393,436 to Blackhurst et al, and British Pat. No. 1,252,544 issued April 9, 1970 to General Motors Corporation. The chief advantage of these filaments is their high tensile strength and lightweight; when these filaments are wound about a rotatable body their high tensile strength translates into a high hoop strength giving the filaments the ability to carry large centrifugal loads.
Two basic problems are encountered with the use of these filaments. One is the difference in thermal and centrifugal expansion rates between the filaments and noncomposite materials; the other problem is that many of these filaments, depending upon the material from which they are made, deteriorate in certain environments, such as in a high temperature oxygen environment as is present in gas turbine engines.
It is often desirable to add a tip shroud and/or a tip seal to the ends of rotor blades to improve efficiency and to reduce vibration; often, however, blades cannot withstand the additional centrifugal loads created by the added mass located near their tip. This is particularly true in the turbine area of a gas turbine engine where temperatures and rotational speeds are very high. This problem becomes more acute as turbine inlet temperatures increase.
SUMMARY OF THE INVENTION It is an object of the present invention to provide a lightweight filament reinforced rotor assembly.
It is a further object of the present invention to reduce stresses in the blades of a rotor assembly.
Accordingly, the present invention contemplates an annular filament wound composite ring surrounding the tips of a set of rotor assembly blades, radially spaced therefrom by radially flexible support means and adapted to come into centrifugal load bearing relationship to the tips of said blades at operating speeds and temperatures for carrying a portion ofthe centrifugal loads of the rotor assembly.
Since the composite ring carries a portion of the centrifugal loads of the rotor assembly, the mass of the disc used to carry the blades may be substantially reduced, possibly to the point of only requiring a thin drum; also, by locating the composite ring around the tips of the blades, the stresses within the blades themselves are reduced to the point where tip shrouds and/or seals may be located near the tips of the blades without overstressing the blades. The overall effect of this invention is to substantially reduce the total weight of a rotor stage by reducing the strength requirements of the rotating parts.
In order to protect the composite ring from a contaminating environment, if one exists, as it would if such a ring were used in the turbine area of a gas turbine engine, and to prevent direct contact between said composite material and said blades, the ring is encapsulated by suitable means such as an annular tube of noncomposite material; if necessary, the tube can be filled I companying drawing.
BRIEF DESCRIPTION OF THE DRAWING FIG. 1 is a partial side elevation view, partly in section, showing a turbine rotor assembly utilizing one embodiment of the present invention.
FIG. 2 is a partial view in perspective looking in direction A in FIG. 1.
DESCRIPTION OF THE PREFERRED EMBODIMENT As an example of a rotor assembly utilizing one embodiment of the present invention, consider the gas turbine engine turbine rotor assembly generally represented by'the numeral 10 in FIG. 1. The rotor assembly 10 comprises a disc 12 and a plurality of radially extending blades 14 each having a root l5 and a tip 16, said blades being circumferentially spaced around the periphery of said disc 12 and attached thereto by any suitable means such as by the well-known fir tree root configuration which is contemplated in this embodiment. Suitable blade locks 17 may also be provided. The root attachment for the blade is not intended to be a part of the present invention.
The rotor assembly 10 also comprises an annular ring 18 made from one or more circumferentially wound carbon filaments embedded in a carbon matrix material. Choice of the filament and matrix material depends upon the particular environment in which the invention is used and may also be, for example, graphite filaments in a resin matrix, boron filaments in an aluminum matrix, saphire filaments in a nickel matrix, or any other suitable filament-matrix combination. The ring 18 is positioned around the tips 16 of the turbine blades 14 and is radially spaced therefrom to allow for differences in thermal and centrifugal growth rates of the ring 18, the blade 14, and the disc 12. As temperature and rotational speed increase the blade tips 16 move toward the ring 18 fastener than the ring moves away from the blade tips; the inner diameter of the ring is sized to result in the ring coming into the centrifugal load bearing relationship to the blade tips when the rotor assembly reaches operating speeds and temperatures. Prior to that time, the blades and disc are sufficiently strong to carry the loads imposed upon them.
In this embodiment each blade is provided with a tip shroud 26. As best shown in FIG. 3, adjacent tip shrouds 26- are in abutting relationship to each other helping to damp blade vibration while at the same time forming the outer wall 28 of the engine gas path. At one end of each shroud 26 is a radially extending labyrinth type seal segment 30. The adjacent seal segments 30 form an annular labyrinth seal ring 32 which cooperates with an outer turbine casing, not shown. The annular composite ring 18 is connected to the rotor assembly through these seal segments 30 by means of an axially extending annular support ring 34. The support ring 34 serves to position the composite ring 18 concentrically about the blade tips 16, and is flexible enough in a radial direction to allow for the differential rates of radial growth between the blade tips and the composite ring 18. The flexibility is accomplished, in this example, by making the connecting ring 34 from thin sheet metal and by giving the ring a Z-shaped cross section as shown in FIG. 1. In some situations it may be desirable that the ring 34 be slotted in an axial direction in several locations such that the composite ring 18 is supported by a plurality of fingers rather than by a full annular ring.
A barrier 36 is provided between the composite ring 18 and the blade tips 16, because direct contact between the blade tips and the composite ring might damage the filaments within the composite ring. In the present embodiment the barrier 36 is an annular tube which also serves to encapsulate the composite ring to protect it from a contaminating environment, such as the high temperature oxygen environment in the turbine section of a gas turbine engine. Further protection may be afforded the composite ring 18, if necessary by filling the annular tube with an inert gas.
As hereinbefore stated, a composite ring, such as the ring 18, surrounding the tips of rotor blades is particularly useful when tip shrouds and tip seals are desired; indeed, in the turbine section of a jet engine, where temperatures and rotational speeds are extremely high, such a composite ring may sometimes be a necessity if tip shrouds are used. In any event, this invention is attractive, notwithstanding the use of tip shrouds, simply to allow a reduction in the mass of the disc supporting the blades. Under certain conditions the requirement for a disc as a centrifugal load carrying member may be eliminated altogether by this invention; in that instance the composite ring would, of course, have to be sufficiently strong to take all the centrifugal loads.
Although the invention has been shown and described with respect to a preferred embodiment thereof, it should be understood by those skilled in the art that various changes and omissions in the form and detail thereof may be made therein without departing from the spirit and the scope of the invention.
Having thus described a typical embodiment of my invention, that which we claim as new and desire to secure by Letters Patent of the United States is:
1. A filament reinforced rotor assembly comprising:
a plurality of radially extending blades circumferentially spaced about the periphery of said rotor and attached thereto;
an annular filament wound composite ring closely surrounding the tips of said blades and radially spaced therefrom when said rotor is at rest; and
radially flexible support means disposed on said rotor, concentrically positioning said ring with respect to said blades and adapted to permit said ring to come into centrifugal load bearing relationship to said blades during rotor operation due to the different centrifugal and thermal growth rates of said blades and said composite ring.
2. The filament reinforced rotor assembly according to claim 1 including barrier means between said ring and said blades to prevent contact between said ring and said blades.
3. The filament reinforced rotor assembly according to claim 1 wherein said support means includes barrier means between said ring and said blades to prevent contact between said ring and said blades.
4. The filament reinforced rotor assembly according to claim 2 wherein said barrier means is an annular tube encapsulating said ring.
5. The filament reinforced rotor assembly according to claim 3 wherein said barrier means is an annular tube encapsulating said ring.
6. The filament reinforced rotor assembly according to claim 5 wherein said annular tube is filled with inert gas.
7. The filament reinforced rotor assembly according to claim 1 wherein said support means includes an axially extending radially flexible annular ring connected to said blades and adapted to provide said radial flexibility to said support means.
8. The filament reinforced rotor assembly according to claim 7 wherein each of said blades includes a tip shroud and a radially extending seal segment, and said composite ring is connected to said seal segments by said axially extending radially flexible annular ring.