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Publication numberUS3549444 A
Publication typeGrant
Publication dateDec 22, 1970
Filing dateDec 28, 1967
Priority dateDec 28, 1967
Publication numberUS 3549444 A, US 3549444A, US-A-3549444, US3549444 A, US3549444A
InventorsHarry S Katz
Original AssigneeHarry S Katz
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Filament wound blade and compressor
US 3549444 A
Abstract  available in
Previous page
Next page
Claims  available in
Description  (OCR text may contain errors)

Dec. 22, 1970 H. s. KATZ 3,549,444

FILAMENT WOUND BLADE AND COMPRESSOR Filed Dec. 28, 1967 I 2 SheetsShet 1 I NVENTO Amen/1572 ATTORNEY United States Patent 3,549,444 FILAMENT WOUND BLADE AND COMPRESSOR Harry S. Katz, 785 Pleasant Valley Way, West Orange, NJ. 07052 Filed Dec. 28, 1967, Ser. No. 694,227 Int. Cl. B6511 81/02 U.S. Cl. 156-175 1 Claim ABSTRACT OF THE DISCLOSURE A filament wound structure such as a compressor and the method for producing the latter in which a pair of compressor blades is formed by means of a continuous filament in which each filament turns an angle greater than 45. Each pair of compressor blades is made as an integral unit and secured around the periphery of a filament wound disc. A shroud comprised of filaments is wound and secured to the tips of the blades to form a high strength compressor assembly free of internal stresses.

In the prior art molding of a single high performance filament-reinforced-plastic structure, such as an aircraft compressor blade, the filaments at the blade tip or root have been either discontinuous or bent sharply on a small radius. If the filaments are discontinuous, such as when individual unidirectional fabric plies are placed into a compression mold, the final structure has proven to be weaker than continuous filament wound structures.

If, on the other hand, the filaments are continuous at the blade tips and roots, the sharp bending radius required to form a relatively thin blade tends to produce filament abrasion, breakage and a great deal of difficulty is experienced in trying to maintain the tension of the filaments uniform and at the proper level in the area of the sharp bend. A prime failure mode has been a tensile failure at the blade root, which points to the factors discussed above and to the problem of joining the blades to another structure such as a disc without introducing localized execessive stresses at the joint. Others have tried to circumvent the problem by using a small diameter metal rod, wound into the root of the blade, however, the latter technique has not proved to be successful because the radius of curvature of the filaments was too small whereby poor control of filament tension and interfilament abrasion resulted. My invention has greatly improved the techniques of the prior art by producing a basic multiple blade structure with better filament control during molding, and better capacity to eliminate high-localized-stress-failure-initiation points.

A number of high strength filaments are currently available, including whiskerized graphite, graphite, S- glass, boron, silicon carbide, E-glass, and carbon. The function of the matrix of a composite material is mainly to transfer the stresses onto the high strength filament, and to separate the filaments to prevent interfilament abrasion and breakage. Metals, ceramics, or plastics can be utilized as the matrix, with the latter being the preferred matrix for my invention. An epoxy resin has been the main tested mtaerial. For higher temperature applications, above 500" F., another resin such as a polyamide or polybenzamidazole may be used.

The use of filament reinforced composites is growing rapidly because of their extremely high strength-toweight ratio. The standard filament has been E-glass, with a tensile strength of 500,000 p.s.i. and strength-todensity ratio of 5.4 million. S-glass has a ratio of 7.8, boron 5.4, and graphite 4.9 million. The ratio for the highest strength steel is 2.1 million. In most aircraft structures, the stiffness or modulus is a critical factor.


The modulus-to-density ratios are E-glass 110, S-glass 140, graphite 720, and boron 650 million.

These properties make it apparent that composite offer great advantages for structures such as rotor and stator stages of compressors. In an aircraft engine compressor blade and disc assembly, a saving of weight of more than 20% can be achieved by using the filament wound process instead of the conventional metallic structures presently widely used. The deterrent to more rapid adaptation of composites for this type of application has been the difiiculty in attaining a large percentage of the potential strength and performance of the composite in the final molded assembly. In order to improve the molding procedure and performance of the final assembly, I have invented an entirely different and subtle improvement for forming a structure such as a compressor disc and blades assembly.

My invention produces a structure that has a greater strength-to-weight ratio and yields better engine performance, due to better filament control during molding and better capacity to eliminate high-localized-stress failureinitiation-points. The nucleus of this procedure is the forming of at least two blades in a continuous loop as an integral structure. The mandrel for forming the loop is preferably a segmented or collapsible metal mold, although a soluble mandrel may be used. The mandrel wrap may be with filaments or unidirectional tape, cut to an appropriate contour for the blade cross-section. Angle or crossplies may be inserted between the longitudinal layers in order to build up a proper orientation for the predetermined stresses to which the blade will be subjected. The segmented mandrel serves to link the uncured dual blade components together onto a ring that has been wrapped with a layer of uncured resin-impregnated filaments. Side plates are bolted onto this assembly of a complete circle of uncured blades. The shroud ring is wrapped over top quadrants of the dual blade loops. The preferred thickness of this loop has been about inch. The assembly is then cured in an oven. A typical cure cycle has been a successive treatment of 3 hours at 1 hour at 250 F., and 1 hour at 350 F. After cooling to room temperature, the side plates and segmented mandrels were removed. Prior to assembly, a mold release had been sprayed onto the metal mold parts to facilitate disassembly. The result is a finished disc and blades assembly, either a rotor or stator stage. If required by the duty cycle of the blades, an erosion resistant coating, such as a polyurethane may be applied.

In order to make my invention clear to those skilled in the art, I have described my invention as applied to a specific embodiment. It is to be understood, however, that my invention can be applied to a multitude of structures. The preferred embodiment disclosed herein is a high performance aircraft rotor or stator section of a compressor in which:

FIG. 1 is a perspective view of an interconnected pair of blades;

FIG. 2 is a perspective view of a mandrel;

FIG. 3 is a perspective view of filaments wrapped around the mandrel;

FIG. 4 is a fragmentary elevational view of a plurality of mandrels assembled;

FIG. 5 is a fragmentary elevational view similar to FIG. 4 with a disc ring added;

FIG. 6 is a fragmentary elevational view similar to FIG. 5 with a shroud added;

FIG. 7 is a section on line 77 of FIG. 6;

FIG. 8 is a section similar to FIG. 6 of an alternate construction of shroud side plates.

Referring to FIG. 1, there is shown a pair of interconnected stator or rotor blades 10 Which have been formed utilizing a continuous filament by using the techniques of my invention. The blade subassembly includes a pair of blades 11 and 12 respectively which are integrally joined at their respective upper tip and lower root portions by interconnecting members 13 and 14. The subassembly 10 forms a unitary filament wound molded structure of extremely high strength and in which internal localized stress-failure-points are virtually eliminated.

The blade subassembly 10 is preferably formed by wind ing a continuous loop of filament such as S-glass or any other known filament material around the periphery or faces of a mandrel 15. The mandrel 15 upon which the loop is wound is preferably a segmented or collapsible metal mold, although other materials comprising the mold could be used including a soluble mandrel. As shown in FIG. 2, the mandrel 15 is composed of a plurality of segments 16. I show nine segments, however, any number of segments can be used since the number used is merely a matter of choice. The segments 16 of the mandrel include any known joining means thereon such as a bolt 17 inserted through a flange 17a. The outside or face portions 18 and 19 of the mandrel are shaped to conform to the stator or stator blades of a compressor. The face 18 is concave and dimensioned so as to form one side of one blade 11. The face 19 is convex and is similarly dimensioned to form the opposite side of a second blade 12. The upper and lower faces 20 and 21 of the mandrel 15 are arched slightly in the longitudinal direction so as to enable curved interconnecting members 13 and 14 to be formed. The angle formed between the blades and interconnecting member is preferably of the order of 90; however, the angle should be greater than 45 to avoid the deleterious affects of interfilament abrasion. Thus, a sharp turn, necessitated by a smaller angle at this juncture is avoided.

Thus, in order to {form the blade subassembly depicted in FIG. 1, the segments 16 of the mandrel 15 are joined together to form the mandrel assembly depicted in FIG. 2. A suitable wrapping material such as E-glass, S-glass, boron, graphite or any other suitable filament wrap or pre-impregnated tape is chosen. Assuming a continuous filament material such as S-glass is chosen, it is then wrapped, using any known wrapping machine to apply appropriate tension continuously around the periphery of the mandrel 15 as is clearly shown in FIG. 3 until a predetermined thickness has been reached. If it is desired to form blades having edges which are thinner at the edges than in the central body portion of the blades, the inner portion of the blades can have narrower pre-impregnated tape applied thereto. Since the narrower tape does not extend to the opposed side edges of the blades they will of course be thinner at the blade edges than in the central body portion of the blades.

After it has been determined to what extent and in which locations stresses occur on the blades, angle or crossplies can be inserted between the longitudinal layers in order to build up a proper orientation for the predetermined stresses to which the blades will be subjected. A resin is then applied to the filaments Wound on the mandrel by a vacuum-pressure impregnation, unless the preferred pre-impregnated tape has been used. Upon completion of the step of wrapping the filaments or applying pre-impregnated tape to the mandrel, an end plate 23 which can be segmented as the mandrel 15 and formed of any suitable material is joined by the flanges 17a and bolts 17 thereon to corresponding flanges on the mandrel 15. The end face of the end plates 23 which butt against the concave face of the blade 11 are convex and the end face of the end plate 23 which abuts the convex face of the blade 12 is concave. Since the end plates are firmly butted against the corresponding. faces of the mandrel 15, the proper contour of the blades 11 and 12 is insured. The end plates 23 function to space the wrapped mandrel subassemblies from each other so as to provide the necessary predetermined spacing. between the blades which is generally the same dimensions as the mandrel 15 if equal spacing between the blades is desired. The latter step -is repeated until the desired number of blade subassemblies are completed necessary to form the rotor or stator stage of the compressor.

-An inner disc ring 24 formed of metal or other suitable material is wrapped with a continuous filament over its circumference and then impregnated with a resin. Alternatively pre-impregnated tape can be used and heated sutficiently to form an uncured resinous mass. The previ ously wrapped mandrels 15 which have been joined together to form a circle are bolted to the flanges 17a on the inner ring 24 or are otherwise joined thereto in any other known manner. The wrapped mandrels 15 are placed on the circumference of the ring 24 before the resin has cured so that the blades will be firmly bonded to the filament wrap on the inner ring 24.

Side plates 26 and 27 are clamped to the opposite side surfaces of the Wrapped mandrels which have been joined around the ring 2 as shown in FIG. 6. A shroud 28 is formed between the side plates 26, 27 by wrapping a continuous filament or by adding pre-impregnated tape having unidirectional filaments therein between the opposed faces of the side plates 26 and 27 to a depth of approximately inch. The top interconnecting portion 13 of each blade wrapped on a mandrel is thereby integrally joined to the shroud. The resin is, of course, in the uncured state immediately after the wrapping is completed and is allowed to cure so as to bond the interconnecting portions 13 to the shroud 28.

An alternative method of applying the shroud is to use a mandrel 15 having integral extended wall portions 25 which are essentially walls of approximately to 4 thickness. The filaments or pre-impregnated tape is then wrapped between the wall portions 25 whereby the wall portions serve as guides for the filaments. The end plates 23 can similarly have wall portions 25 applied thereto if it is desired to have a continuous wall circumscribing the outer periphery of the turbine.

The compressor assembly is now cured in an oven. A typical cure cycle time is a successive treatment of approximately 3 hours at F., 1 hour at 250 F. and thence 1 hour at 350 F. After cooling the assembly to room temperature, the side plates (if the latter are used) and the segmented mandrels are removed. A mold release is applied onto the mold parts to facilitate disassembly. The result is a finished disc and blade assembly comprising either a rotor or stator stage of a compressor. If required by the duty cycle of the blades an erosion resistant coating, such as polyurethane may be applied. The resulting structure is extremely strong and because I have eliminated all discontinuity in filaments and I have ellminated all sharp radius turns by forming a two bladed subassembly, there are no areas of localized stress built into the compressor structure, and the blades can withstand a high degree of vibrational, torsional and impact stresses.

While I have shown and described a particular embodiment of the present invention, it will be apparent to those skilled in the art that various changes and modifications may be made without departing from the invention in its broader aspects. It is therefore contemplated in the appended claims to cover all such changes and modifications as fall within the true spirit and scope of the invention.

What is claimed as new and desired to be secured by Letters Patent of the United States is:

1. A method for forming a multi-blade turbine comprising the steps of,

(a) wrapping a filament coated with resin continuously around a mandrel having faces thereon shaped to conform to that of compressor blades and other faces shaped to permit said blades to be interconnected by said filament,

6 (b) connecting the wrapped mandrel to an end plate, (h) removing the mandrels, and the disc after the asand sembly has cured. (c) repeating the step of connecting the wrapped mandrel to an adjacent end plate until a closed circle of References Cited inteegprrlllracted wrapped mandrels and end plates 5 UNITED STATES PATENTS (d) applying a filament coated resin continuously 219501083 8/1960 Compton at 156*18O around a disc having joining means thereon, 3,057,767 10/1962 Kaplan 156245 (e) connecting the circle of mandrels at points along 3,403,844 10/1968 stofiel' 156*173 its inner diameter to the disc containing an uncured resin thereon, 10 JOHN T. GOOLKASIAN, Primary Examlner (f) applying a filament coated with resin along the D. LFRITSCH, Assistant Examiner outer circumference of said mandrels so as to form ashroud, U.S. Cl. X.R.

(g) permitting the assembly to cure, and 15 156-180 182, 433; 264-136

Patent Citations
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US2950083 *Jul 23, 1954Aug 23, 1960Thompson Ramo Wooldridge IncBlade assembly
US3057767 *Apr 1, 1958Oct 9, 1962Poly Ind IncMethod of making compressor blades
US3403844 *Oct 2, 1967Oct 1, 1968Gen ElectricBladed member and method for making
Referenced by
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US3675294 *Dec 22, 1970Jul 11, 1972Secr DefenceMethod of making a bladed rotor
US3758232 *Apr 23, 1971Sep 11, 1973Secr DefenceBlade assembly for gas turbine engines
US3905722 *Mar 12, 1973Sep 16, 1975Rolls Royce 1971 LtdFluid flow machines
US4046489 *Oct 8, 1975Sep 6, 1977Eagle Motive Industries, Inc.Aerodynamic fan blade
US4098559 *Jul 26, 1976Jul 4, 1978United Technologies CorporationPaired blade assembly
US4191510 *Aug 31, 1977Mar 4, 1980Societe Nationale D'etude Et De Construction De Moteurs D'aviation (S.N.E.C.M.A.)Axial flow compressor rotor drum
US4195967 *Oct 21, 1977Apr 1, 1980Messerschmitt-Boelkow-Blohm GmbhRotor free of flapping hinges and free of lead-lag hinges
US4500379 *Aug 25, 1983Feb 19, 1985Olin CorporationMethod of making a composite fiber reinforced plastic frame
US4580944 *May 15, 1984Apr 8, 1986United Technologies CorporationAerodynamic flexible fairing
US4747900 *Jun 5, 1985May 31, 1988Rolls-Royce PlcMethod of manufacture of compressor rotor assembly
US4786347 *Jun 5, 1985Nov 22, 1988Rolls-Royce PlcMethod of manufacturing an annular bladed member having an integral shroud
US4826645 *Jun 5, 1985May 2, 1989Rolls-Royce LimitedMethod of making an integral bladed member
US4907736 *May 18, 1988Mar 13, 1990Airfoil Textron Inc.Method of forming articles
US5013216 *Nov 29, 1989May 7, 1991Airfoil Textron Inc.Composite blade perform with divergent root
US5018271 *Sep 9, 1988May 28, 1991Airfoil Textron Inc.Method of making a composite blade with divergent root
US5176501 *Dec 17, 1990Jan 5, 1993The University Of British ColumbiaPropeller with an elastic sleeve
US5362344 *Feb 3, 1993Nov 8, 1994Avco CorporationDucted support housing assembly
US7938627Nov 11, 2005May 10, 2011Board Of Trustees Of Michigan State UniversityWoven turbomachine impeller
US8449258May 28, 2013Board Of Trustees Of Michigan State UniversityTurbomachine impeller
US8506254Apr 18, 2011Aug 13, 2013Board Of Trustees Of Michigan State UniversityElectromagnetic machine with a fiber rotor
US20070297905 *Nov 11, 2005Dec 27, 2007Norbert MullerWoven Turbomachine Impeller
US20140064956 *Aug 15, 2013Mar 6, 2014Rolls-Royce PlcGuide vane assembly
WO2007013892A2 *Nov 11, 2005Feb 1, 2007Board Of Trustees Of Michigan State UniversityComposite turbomachine impeller and method of manufacture
WO2007013892A3 *Nov 11, 2005Sep 27, 2007Univ Michigan StateComposite turbomachine impeller and method of manufacture
U.S. Classification156/175, 416/241.00A, 416/241.00R, 29/419.1, 29/889.7, 156/433, 156/180, 264/136, 416/230, 156/182
International ClassificationB29C53/56, B29C53/82
Cooperative ClassificationB29C53/564, B29C53/824, B29L2031/08
European ClassificationB29C53/82B3, B29C53/56C