US3215511A - Gas turbine nozzle vane and like articles - Google Patents

Gas turbine nozzle vane and like articles Download PDF

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US3215511A
US3215511A US183939A US18393962A US3215511A US 3215511 A US3215511 A US 3215511A US 183939 A US183939 A US 183939A US 18393962 A US18393962 A US 18393962A US 3215511 A US3215511 A US 3215511A
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vane
percent
edge
gas turbine
nozzle vane
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US183939A
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Charles G Chisholm
Glenn A Fritzlen
Edward M Leach
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Union Carbide Corp
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Union Carbide Corp
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Priority to GB1096063A priority patent/GB1036446A/en
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • F01D5/284—Selection of ceramic materials
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T29/00—Metal working
    • Y10T29/49—Method of mechanical manufacture
    • Y10T29/49316—Impeller making
    • Y10T29/4932—Turbomachine making
    • Y10T29/49321—Assembling individual fluid flow interacting members, e.g., blades, vanes, buckets, on rotary support member
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T428/00—Stock material or miscellaneous articles
    • Y10T428/12—All metal or with adjacent metals
    • Y10T428/12014—All metal or with adjacent metals having metal particles
    • Y10T428/12028—Composite; i.e., plural, adjacent, spatially distinct metal components [e.g., layers, etc.]
    • Y10T428/12063—Nonparticulate metal component
    • Y10T428/12139—Nonmetal particles in particulate component
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T428/00—Stock material or miscellaneous articles
    • Y10T428/12—All metal or with adjacent metals
    • Y10T428/12451—Macroscopically anomalous interface between layers

Definitions

  • the present invention relates to gas turbine nozzle vanes and like articles and more particularly to a novel composite gas turbine nozzle vane having both improved thermal shock resistance and improved resistance to deformation under high temperature operating conditions.
  • the nozzle vanes of the apparatus are continuously subjected to a very severe and complex environment, e.g. high temperatures, severe thermal gradients, erosion, and forces which tend to deform the nozzle vanes, particularly the trailing portions thereof.
  • the aforesaid environmental conditions are developed mainly by the high velocity, high temperature etiluent of the turbine which continuously impinges on the leading edges of the nozzle vanes and, in its continuous passage, exerts strenuous forces against the thinner trailing edges of the vanes.
  • the heat which is conducted from the very high temperature regions at the leading portions of the vanes to the thinner trailing portions renders these thinner portions less capable of withstanding the strenuous forces exerted by the turbine efiluent.
  • FIGURE 1 shows a cross section of an embodiment of a composite gas turbine nozzle vane of the present invention.
  • FIGURE 2 shows the upper surface of the nozzle vane of FIGURE 1.
  • FIGURE 3 is a cross section of a further embodiment of the present invention and FIGURE 4 shows a still further embodiment of the present invention.
  • a gas turbine nozzle vane or like article in accordance with the present invention is a composite article of manufacture comprising a metallic body member and a heat-resistant non-metallic member fixedly joined thereto and forming at least part of an edge surface of the nozzle vane.
  • FIGURES 1 and 2 show in FIGURES 1 and 2 a gas turbine nozzle vane 1 having a metal body member 3 and a leading edge member 5 which extends over-the portion of the leading edge of the vane at which "ice the highest temperatures are developed during turbine operation.
  • Body member 3 is shown in the drawing to have a hollow center which can be utilized in providing cooling for the vane.
  • hollow construction is not essential to the present invention and solid vanes can also be used.
  • the body member is formed of a suitable high temperature metal or alloy such as the well-known nickelor cobalt-base super-alloys while the edge material is formed of different material such as a heat-resistant ceramic or metal ceramic material. Heat-resistant metal compounds, and heat-resistant bonded mixtures of metal compounds can also be used for the edge material.
  • the body members and edges are separately formed and can be fabricated by investment casting, extrusion, forging, slip casting or other known techniques. When the separate body and edge members are formed to the proper dimensions, they are fixedly connected to provide an article having a surface which is suitable for a gas turbine nozzle vane or like article. This can be accomplished for example, by the keying joint illustrated in the drawing. It is to be understood, of course, that while the particular manner of mechanically connecting the component parts of the composite turbine blade is not critical, the connection should be such as to provide sufficient strength for the intended use.
  • the composite gas turbine vane as above described when in operation, is exposed to the flow of turbine effluent, indicated in the drawing by the arrows, with the result that extremely high temperatures are developed at the leading edge.
  • a thermal barrier which is formed at the joint between the separately formed edge member and vane, the conduc tion of heat from the non-metallic leading edge to the body member and relatively thin trailing portion of the vane is substantially reduced.
  • whatever heat is developed in the body member of the vane tends to be more uniformy distributed so that the temperature gradient in the body member is much less severe than would be expected.
  • the nozzle vane of the present invention in effect, has greatly improved resistance to deformation and thermal shock and can thus be used for longer periods at operating temperatures substantially higher than the maximum permissible with a gas turbine vane of integral construction having the same composition as either the edge or body member of the composite vane.
  • a separate edge member 7 can be similar ly provided for the trailing portion of the nozzle vane. In this way, the temperature gradients in the body member are further moderated and conduction of heat from the forward portion of the vane to the trailing edge is further reduced. As a result, higher operating temperatures and longer operating times can be used.
  • the component parts of the nozzle vane are formed of different materials which are especially suitable for use under the operating conditions encountered at the locations of the respective component parts.
  • the composite structure of the present invention in effect divides the nozzle vane into a very high temperature forward region and a lowertemperature trailing portion which is subjected to severe deforming forces, different materials having different properties to accommodate the distinct operating environments can be employed with advantage.
  • the edge member was formed from a metal ceramic ma- 3 terial (23 percent alumina; '77 percent chromium) and the vane was formed from a nickel-base alloy (12.5 percent chromium, 4.5 percent molybdenum, 6 percent aluminum, 2 percent columbium plus tantalum and the 4 Example I
  • a nozzle vane as shown in FIGURES l and 2 was constructed having a metal ceramic leading edge (23 percent alumina; 77 percent chromium) and a nickel-base alloy balance essentially nickel with minor amounts of carbon, 5 boron, zirconium and titanium).
  • the metal ceramic edge member was formed by slip In the manufacture of the composite vanes of this casting and subsequently positioned in a wax pattern die invention from different materials, any suitable technique and incorporated as an Insert the f Pattsrn of a Vane can be employed for preparing the component parts for first'stage nozzle 9 9 gas turbme' such as for example, slip casting extrusion and Tne wax pattern containing the metal ceramic insert forming.
  • any suitable method for joining the was.processed by the lost Wax techmque to component parts can be used provided that a thermal fashion a mold After the usual additional mold procbarrier is maintained between the parts and the mechani- P nlcel'base anoy was Cast In the mold- After cal Strength of the joint is suflicient to Withstand the solidification and cooling of the metal, the resultant artiforces encountered in operation.
  • a particularly effective C16 was refnovad from the molfi and F
  • the ajmcle way to accomplish this is to first prepare the edge mem- 20 thus Obtameq was fomposlte tufbme Vane having 3 her from the desired heat-resistant material in a form metal ceramlc edge lowed to a mckel'base alloy body such as shown in the drawing and then incorporate the edge member as an insert in a Wax pattern designed
  • the nozzle vane described in Example I was tested for the body member
  • the Welbknown 10st or under very severe conditions as described in the followother technique is then followed to provide a mold in mg Example which the metal body member is cast.
  • the resulting Example H article is a composite turbine blade having an edge surface member securely fixed to the vane.
  • Example I The composite nozzle vane described in Example I was In addition to the materials previously mentioned, a subjected to a continuous test of ten consecutive threewide variety of substances Can be effec lvely empl ye hour cycles. Each of the first six temperature cycles was in the formation of the component parts of the turbine as f ll blade of the present invention.
  • the following table sets o forth a selection of materials which can be advantageous- 30 mlmltes Soak Q 2000 followed y ly employed; however, this tabulation is not to be con- 100 Counts 0f 1 minute hot (20000 d sidered as limitative. 30 seconds cold (quenched to below 400 F.)
  • Nickel-base alloy 12.5 percent Cr, 4.5 percent Mo, 6 percent 1281, 3 percent total of Cb, Ta, B, 0, Ti,
  • metal ceramic materials in the order in which r they appear in the table are disclosed in US. Patents 2,698,990, 2,656,596 and 2,783,530, respectively; titanium diboride is disclosed in U.S. Patent 3,003,885. Cemented titanium carbide is sold commercially as Kentanium, a trademark of Kennametal Inc.
  • a nozzle vane of the present invention formed of any of the above-specified or similar edge materials with any of the specified or similar vane materials will provide an article characterized by superior thermal shock resistance and improved resistance to deforming forces at its trailing portion.
  • certain combinations of material will provide a much better over-all turbine vane than others. Therefore, in the practice of the present invention, the conditions to be encountered at the various portions of the turbine vane are analyzed and the material best suited for operation under a particular condition is used in the construction of the vane component for that portion of the blade.
  • leading or trailing edge components of the turbine vane may extend over the entire edge surface or only at a critical portion or portions of the edge surface.
  • the forward edge member 5" extended over the entire leading edge and into the shroud section of the vane 9.
  • the edge member 5 was positioned in the center of the leading edge covering about 75 percent of the edge section.
  • the trailing portion it is recommended that the trailing edge components constitute the entire trailing edge of the vane and also a portion of the shroud section of the vane, because of the thinner cross section available to resist the complex stresses usually concentrated at that portion of the turbine vane.
  • the present invention provides a novel turbine vane having increased resistance to thermal shock and increased resistance to deformation at its trailing edge.
  • a gas turbine nozzle vane and the like comprising a metallic body member and a pre-formed non-metallic forward edge member closely engaged with the body member through a keyed mechanical interlock joint said body member being formed of a material selected from the group consisting of nickel and cobalt base alloys and said edge member being formed of a heat resistant matenial selected from the group consisting of ceramics and non-metallic metal compounds.
  • An article in accordance with claim 1 wherein the selected ceramic material has a composition of about 15 percent alumina, about 25 percent chromium and about 60 percent tungsten.
  • An article in accordance with claim 1 wherein the selected ceramic material has a composition of about 19 percent alumina, about 59 percent chromium, about 20 percent molybdenum, and about 2 percent titanium.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Ceramic Engineering (AREA)
  • Materials Engineering (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Description

1965 c. G. CHISHOLM ETAL 3, 5,5
GAS TURBINE NOZZLE VANE AND LIKE ARTICLES Filed March 30, 1962 9 4 l v I INVENTORS LJ CHARLES c. CHISHOLM GLENN A. FRITZLEN EDWARD M. LEACH ATTORNEY United States Patent 3,215,511 GAS TURBINE NOZZLE VANE AND LIKE ARTICLES Charles G. Chisholm and Glenn A. Fritzlen, Kokomo, and Edward M. Leach, Bunker Hill, Ind, assignors to Union Carbide Corporation, a corporation of New York Filed Mar. 30, 1952, Ser. No. 183,939 7 Claims. (Cl. 29-483) The present invention relates to gas turbine nozzle vanes and like articles and more particularly to a novel composite gas turbine nozzle vane having both improved thermal shock resistance and improved resistance to deformation under high temperature operating conditions.
In the operation of gas turbines and the like, the nozzle vanes of the apparatus are continuously subjected to a very severe and complex environment, e.g. high temperatures, severe thermal gradients, erosion, and forces which tend to deform the nozzle vanes, particularly the trailing portions thereof.
The aforesaid environmental conditions are developed mainly by the high velocity, high temperature etiluent of the turbine which continuously impinges on the leading edges of the nozzle vanes and, in its continuous passage, exerts strenuous forces against the thinner trailing edges of the vanes.
The continuous impinging contact of the eflluent against the leading edges of the vanes tends to'erode these edges and, further, produces severe temperature gradients in the vanes by the devlopment of localized extremely high temperatures at the leading edges and in contiguous sections of the trailing portions of the vanes.
In addition to the severe thermal shock to which the vanes are subjected due to the aforedscribed conditions, the heat which is conducted from the very high temperature regions at the leading portions of the vanes to the thinner trailing portions renders these thinner portions less capable of withstanding the strenuous forces exerted by the turbine efiluent.
Consequently, due to the severe temperature gradients and other thermal conditions encountered by the nozzle vanes presently employed in gas turbines and like apparatus, the permissible operational periods and operating temperatures of these apparatus are considerably limited.
It is therefore an object of the present invention to provide a gas turbine nozzle vane having both improved resistance to thermal shock and increased resistance to the deforming forces exerted by the turbine efliuent.
Other objects will be apparent from the following description and claims taken in conjunction with the drawing in which:
FIGURE 1 shows a cross section of an embodiment of a composite gas turbine nozzle vane of the present invention.
FIGURE 2 shows the upper surface of the nozzle vane of FIGURE 1.
FIGURE 3 is a cross section of a further embodiment of the present invention and FIGURE 4 shows a still further embodiment of the present invention.
A gas turbine nozzle vane or like article in accordance with the present invention is a composite article of manufacture comprising a metallic body member and a heat-resistant non-metallic member fixedly joined thereto and forming at least part of an edge surface of the nozzle vane.
The present invention will be more clearly understood by referring to the drawing which shows in FIGURES 1 and 2 a gas turbine nozzle vane 1 having a metal body member 3 and a leading edge member 5 which extends over-the portion of the leading edge of the vane at which "ice the highest temperatures are developed during turbine operation. Body member 3 is shown in the drawing to have a hollow center which can be utilized in providing cooling for the vane. However, hollow construction is not essential to the present invention and solid vanes can also be used.
In the present invention, the body member is formed of a suitable high temperature metal or alloy such as the well-known nickelor cobalt-base super-alloys while the edge material is formed of different material such as a heat-resistant ceramic or metal ceramic material. Heat-resistant metal compounds, and heat-resistant bonded mixtures of metal compounds can also be used for the edge material. The body members and edges are separately formed and can be fabricated by investment casting, extrusion, forging, slip casting or other known techniques. When the separate body and edge members are formed to the proper dimensions, they are fixedly connected to provide an article having a surface which is suitable for a gas turbine nozzle vane or like article. This can be accomplished for example, by the keying joint illustrated in the drawing. It is to be understood, of course, that while the particular manner of mechanically connecting the component parts of the composite turbine blade is not critical, the connection should be such as to provide sufficient strength for the intended use.
The composite gas turbine vane as above described, when in operation, is exposed to the flow of turbine effluent, indicated in the drawing by the arrows, with the result that extremely high temperatures are developed at the leading edge. However, due to what may be termed a thermal barrier which is formed at the joint between the separately formed edge member and vane, the conduc tion of heat from the non-metallic leading edge to the body member and relatively thin trailing portion of the vane is substantially reduced. Also, whatever heat is developed in the body member of the vane tends to be more uniformy distributed so that the temperature gradient in the body member is much less severe than would be expected.
Consequently, by virtue of its composite construction, the nozzle vane of the present invention, in effect, has greatly improved resistance to deformation and thermal shock and can thus be used for longer periods at operating temperatures substantially higher than the maximum permissible with a gas turbine vane of integral construction having the same composition as either the edge or body member of the composite vane.
In a further embodiment of the invention, illustrated in FIGURE 3, a separate edge member 7 can be similar ly provided for the trailing portion of the nozzle vane. In this way, the temperature gradients in the body member are further moderated and conduction of heat from the forward portion of the vane to the trailing edge is further reduced. As a result, higher operating temperatures and longer operating times can be used.
In the practice of the present invention, the component parts of the nozzle vane are formed of different materials which are especially suitable for use under the operating conditions encountered at the locations of the respective component parts.
For example, since the composite structure of the present invention in effect divides the nozzle vane into a very high temperature forward region and a lowertemperature trailing portion which is subjected to severe deforming forces, different materials having different properties to accommodate the distinct operating environments can be employed with advantage. In a specific nozzle vane in accordance with the present invention, and in a form similar to that shown in FIGURE 1, the edge member was formed from a metal ceramic ma- 3 terial (23 percent alumina; '77 percent chromium) and the vane was formed from a nickel-base alloy (12.5 percent chromium, 4.5 percent molybdenum, 6 percent aluminum, 2 percent columbium plus tantalum and the 4 Example I A nozzle vane as shown in FIGURES l and 2 was constructed having a metal ceramic leading edge (23 percent alumina; 77 percent chromium) and a nickel-base alloy balance essentially nickel with minor amounts of carbon, 5 boron, zirconium and titanium). It was found that under vane member (12-5 parcjent Chromlum P f molyb' severe thermal Shock testing at temperatures up to denum, 6 percent aluminum, 2 percentoolumbium plus about 21000 both the metal ceramic leading edge tantalum and the balance essentially nickel with minor of the Vane and the body membm. were substantially amounts of carbon, boron, zirconium and titanium). unafiacte 10 The metal ceramic edge member was formed by slip In the manufacture of the composite vanes of this casting and subsequently positioned in a wax pattern die invention from different materials, any suitable technique and incorporated as an Insert the f Pattsrn of a Vane can be employed for preparing the component parts for first'stage nozzle 9 9 gas turbme' such as for example, slip casting extrusion and Tne wax pattern containing the metal ceramic insert forming. Also, any suitable method for joining the was.processed by the lost Wax techmque to component parts can be used provided that a thermal fashion a mold After the usual additional mold procbarrier is maintained between the parts and the mechani- P nlcel'base anoy was Cast In the mold- After cal Strength of the joint is suflicient to Withstand the solidification and cooling of the metal, the resultant artiforces encountered in operation. A particularly effective C16 was refnovad from the molfi and F The ajmcle way to accomplish this is to first prepare the edge mem- 20 thus Obtameq was fomposlte tufbme Vane having 3 her from the desired heat-resistant material in a form metal ceramlc edge lowed to a mckel'base alloy body such as shown in the drawing and then incorporate the edge member as an insert in a Wax pattern designed The nozzle vane described in Example I was tested for the body member The Welbknown 10st or under very severe conditions as described in the followother technique is then followed to provide a mold in mg Example which the metal body member is cast. The resulting Example H article is a composite turbine blade having an edge surface member securely fixed to the vane. The composite nozzle vane described in Example I was In addition to the materials previously mentioned, a subjected to a continuous test of ten consecutive threewide variety of substances Can be effec lvely empl ye hour cycles. Each of the first six temperature cycles was in the formation of the component parts of the turbine as f ll blade of the present invention. The following table sets o forth a selection of materials which can be advantageous- 30 mlmltes Soak Q 2000 followed y ly employed; however, this tabulation is not to be con- 100 Counts 0f 1 minute hot (20000 d sidered as limitative. 30 seconds cold (quenched to below 400 F.)
TABLE Type Material Component Specific Example Nominal Composition, Weight/Percent Alloy Body member" Nickel-base alloy 12.5 percent Cr, 4.5 percent Mo, 6 percent 1281, 3 percent total of Cb, Ta, B, 0, Ti,
1" Do- Cobalt-base alloy.-. 21.5 percent Cr, 10 percent W, 7.5 percent Ta, 4 percent total of Si, C, Fe, Ni. Ceramic. Alumina A1203.
D0- Bcryllia About 99 percent BeO. Metal Ceramic Metal Ceramic 23 percent alumina, 77 percent chromium. Do .do 25 percent Cr, 60 percent W, 15 percent alumina. Do do .do. 59 percent Cr, 20 percent Mo, 19 percent alumina, 2 percent Ti. Metal Compound Titanium diboride 90 to 99-]- percent T182 plus modifying ingredients and impurities. Do Silicon nitride Principally Si3N Bonded mixture Cemented titanium Titanium Carbide bonded with about carbide. percent Ni or C0.
The metal ceramic materials, in the order in which r they appear in the table are disclosed in US. Patents 2,698,990, 2,656,596 and 2,783,530, respectively; titanium diboride is disclosed in U.S. Patent 3,003,885. Cemented titanium carbide is sold commercially as Kentanium, a trademark of Kennametal Inc.
In general, a nozzle vane of the present invention formed of any of the above-specified or similar edge materials with any of the specified or similar vane materials will provide an article characterized by superior thermal shock resistance and improved resistance to deforming forces at its trailing portion. However, it is clear that for particular applications certain combinations of material will provide a much better over-all turbine vane than others. Therefore, in the practice of the present invention, the conditions to be encountered at the various portions of the turbine vane are analyzed and the material best suited for operation under a particular condition is used in the construction of the vane component for that portion of the blade.
The following examples are provided to further illustrate the present invention.
Each of the four subsequent cycles was as follows:
30 minutes soak at 2100 F. followed by counts of 1 minute hot (2100 F.) and 30 seconds cold (quenched to below 400 F.)
The above-described test is very severe and it is ord-inarily expected that most super alloy nozzle vanes will withstand only 2 to 4 cycles at 1800 F. to 2000 F.
It was found that after testing as in Example II, the composite nozzle vane of the invention was satisfactory in all respects and that all portions of the vane, including the nickel-base alloy trailing edge were substantially unaffected by the severe conditions encountered in the test. The fact that the metallic trailing portion was unaffected was surprising since under the same test conditions, the trailing edge of an integral nickel-base alloy turbine vane of the same composition would have failed.
As to modifications of the present invention, it was found that the leading or trailing edge components of the turbine vane may extend over the entire edge surface or only at a critical portion or portions of the edge surface. Several modifications were successfully made and found to be suitable for service. In one configuration illustrated in FIGURE 4, the forward edge member 5" extended over the entire leading edge and into the shroud section of the vane 9. In another, as shown in FIGURE 2, the edge member 5 was positioned in the center of the leading edge covering about 75 percent of the edge section. As regards the trailing portion, it is recommended that the trailing edge components constitute the entire trailing edge of the vane and also a portion of the shroud section of the vane, because of the thinner cross section available to resist the complex stresses usually concentrated at that portion of the turbine vane.
From the above description, it can be seen that the present invention provides a novel turbine vane having increased resistance to thermal shock and increased resistance to deformation at its trailing edge.
While the above description has been directed to particular materials for use in the construction of the composite nozzle vanes of this invention, it is to be understood that other known metals and non-metallic materials which have been used or suggested for use in the manufacture of nozzle vanes and the like can also be employed with advantage in accordance with the present invention.
What is claimed is:
1. A gas turbine nozzle vane and the like comprising a metallic body member and a pre-formed non-metallic forward edge member closely engaged with the body member through a keyed mechanical interlock joint said body member being formed of a material selected from the group consisting of nickel and cobalt base alloys and said edge member being formed of a heat resistant matenial selected from the group consisting of ceramics and non-metallic metal compounds.
2. An article in accordance with claim 1 wherein the selected non-metallic metal compound is titanium diboride.
3. An article in accordance with claim 1 wherein the selected non-metallic metal compound is silicon nitride.
4. An article in accordance with claim 1 wherein the selected ceramic material has a composition of about 23 percent alumina and about 77 percent chromium.
5. An article in accordance with claim 1 wherein the selected ceramic material has a composition of about 15 percent alumina, about 25 percent chromium and about 60 percent tungsten.
6. An article in accordance with claim 1 wherein the selected ceramic material has a composition of about 19 percent alumina, about 59 percent chromium, about 20 percent molybdenum, and about 2 percent titanium.
7. An article in accordance with claim 1 wherein the selected ceramic material is cemented titanium carbide bonded with about percent of a material selected from the group consisting of nickel and cobalt.
References Cited by the Examiner UNITED STATES PATENTS 2,431,660 11/47 Gaudenzi. 2,769,611 1 1/ 56 Kramer. 2,774,678 12/56 Rodman 11753 2,946,681 7/60 Probst. 3,068,556 12/62 Kramer. 3,148,954 9/64 Haas.
FOREIGN PATENTS 1,118,53 6 11/ 6'1 Germ-any.
856,680 12/60 Great Britain.
379,097 3 40 Italy.
DAVID L. RECK, Primary Examiner.
HYLAND BIZOT, Examiner.

Claims (1)

1. A GAS TURBINE NOZZLE VANE AND THE LIKE COMPRISING A METALLIC BODY MEMBER AND A PRE-FORMED NON-METALLIC FORWARD EDGE MEMBER CLOSELY ENGAGED WITH THE BODY MEMBER THROUGH A KEYED MECHANICAL INTERLOCK JOINT SAID BODY MEMBER BEING FORMED OF A MATERIAL SELECTED FROM THE GROUP CONSISTING OF NICKEL AND COBALT BASE ALLOYS AND SAID EDGE MEMBER BEING FORMED OF A HEAT RESISTANT MATERIAL SELECTED FROM THE GROUP CONSISTING OF CERAMICS AND NON-METALLIC METAL COMPOUNDS.
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Cited By (70)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3304056A (en) * 1965-03-19 1967-02-14 Hitachi Ltd Turbine blades
US3315941A (en) * 1965-04-27 1967-04-25 Rolls Royce Aerofoil blade for use in a hot fluid stream
US3394918A (en) * 1966-04-13 1968-07-30 Howmet Corp Bimetallic airfoils
US3486833A (en) * 1967-05-05 1969-12-30 Gen Motors Corp High temperature composite gas turbine engine components
US3660882A (en) * 1969-04-28 1972-05-09 Boehler & Co Ag Geb Process for the production of turbine blades
US3844728A (en) * 1968-03-20 1974-10-29 United Aircraft Corp Gas contacting element leading edge and trailing edge insert
US3844727A (en) * 1968-03-20 1974-10-29 United Aircraft Corp Cast composite structure with metallic rods
US4022540A (en) * 1975-10-02 1977-05-10 General Electric Company Frangible airfoil structure
US4241110A (en) * 1978-07-20 1980-12-23 Mitsubishi Jukogyo Kabushiki Kaisha Method of manufacturing rotor blade
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US20110219782A1 (en) * 2010-03-10 2011-09-15 Rolls-Royce Deutschland Ltd & Co Kg Aerodynamically shaped supporting and/or fairing element in the bypass duct of a gas-turbine engine
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US10662779B2 (en) 2016-11-17 2020-05-26 Raytheon Technologies Corporation Gas turbine engine component with degradation cooling scheme
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US10711624B2 (en) 2016-11-17 2020-07-14 Raytheon Technologies Corporation Airfoil with geometrically segmented coating section
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US10767487B2 (en) 2016-11-17 2020-09-08 Raytheon Technologies Corporation Airfoil with panel having flow guide
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US3304056A (en) * 1965-03-19 1967-02-14 Hitachi Ltd Turbine blades
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US3486833A (en) * 1967-05-05 1969-12-30 Gen Motors Corp High temperature composite gas turbine engine components
US3844728A (en) * 1968-03-20 1974-10-29 United Aircraft Corp Gas contacting element leading edge and trailing edge insert
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US4241110A (en) * 1978-07-20 1980-12-23 Mitsubishi Jukogyo Kabushiki Kaisha Method of manufacturing rotor blade
FR2478734A1 (en) * 1980-03-19 1981-09-25 Gen Electric PROCESS FOR REPAIRING GAS TURBINE ENGINE BLADES AND ELEMENT FOR SUCH REPAIR
US4512719A (en) * 1981-07-24 1985-04-23 Motoren-Un Turbinen-Union Munchen Gmbh Hot gas wetted turbine blade
JPS59160004A (en) * 1983-03-01 1984-09-10 Agency Of Ind Science & Technol Stationary blade for gas turbine
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JPS59137878U (en) * 1983-03-07 1984-09-14 三菱農機株式会社 Crawler traveling device
JPS59138784U (en) * 1983-03-07 1984-09-17 トヨタ自動車株式会社 Surplus time notification device
JPS59174475U (en) * 1983-05-11 1984-11-21 株式会社大林組 Piping support fittings
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US4728262A (en) * 1986-01-22 1988-03-01 Textron Inc. Erosion resistant propellers
EP0292086B1 (en) * 1987-05-21 1991-07-31 Hudson Products Corporation Fan blade for an axial flow fan and method of forming same
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US5351395A (en) * 1992-12-30 1994-10-04 General Electric Company Process for producing turbine bucket with water droplet erosion protection
US5358379A (en) * 1993-10-27 1994-10-25 Westinghouse Electric Corporation Gas turbine vane
US5895205A (en) * 1995-08-07 1999-04-20 General Electric Co. Method for repairing partitions of a turbine diaphragm
EP0806546A1 (en) * 1996-05-02 1997-11-12 Asea Brown Boveri Ag Thermally stressed turbomachine vane with a ceramic insert in the leading edge
US5782607A (en) * 1996-12-11 1998-07-21 United Technologies Corporation Replaceable ceramic blade insert
EP0995880A3 (en) * 1998-10-19 2002-01-23 Alstom Turbine blade
US20020197152A1 (en) * 2001-06-26 2002-12-26 Jackson Melvin Robert Airfoils with improved oxidation resistance and manufacture and repair thereof
US6609894B2 (en) * 2001-06-26 2003-08-26 General Electric Company Airfoils with improved oxidation resistance and manufacture and repair thereof
US20060198734A1 (en) * 2002-12-18 2006-09-07 Alessandro Coppola Manufacturing method for obtaining high-temperature components for gas turbines and components thus obtained
US20060285973A1 (en) * 2005-06-17 2006-12-21 Siemens Westinghouse Power Corporation Trailing edge attachment for composite airfoil
WO2007001511A1 (en) * 2005-06-17 2007-01-04 Siemens Power Generation, Inc. Trailing edge attachment for composite airfoil
US7393183B2 (en) * 2005-06-17 2008-07-01 Siemens Power Generation, Inc. Trailing edge attachment for composite airfoil
EP2687679A1 (en) * 2005-06-17 2014-01-22 Siemens Energy, Inc. Trailing edge attachment for composite airfoil
US20070140859A1 (en) * 2005-12-21 2007-06-21 Karl Schreiber Leading edge configuration for compressor blades of gas turbine engines
US7744346B2 (en) * 2005-12-21 2010-06-29 Rolls-Royce Deutschland Ltd & Co Kg Leading edge configuration for compressor blades of gas turbine engines
US20070240845A1 (en) * 2006-04-18 2007-10-18 Graham Stephen D Investment cast article and method of production thereof
US20080159856A1 (en) * 2006-12-29 2008-07-03 Thomas Ory Moniz Guide vane and method of fabricating the same
US20120121395A1 (en) * 2009-04-23 2012-05-17 Volvo Aero Corporation Method for fabricating a gas turbine engine component and a gas turbine engine component
US20110219782A1 (en) * 2010-03-10 2011-09-15 Rolls-Royce Deutschland Ltd & Co Kg Aerodynamically shaped supporting and/or fairing element in the bypass duct of a gas-turbine engine
US20120234967A1 (en) * 2011-03-16 2012-09-20 Christou Kyriakos C Low-heat-transfer interface between metal parts
US9012824B2 (en) * 2011-03-16 2015-04-21 Raytheon Company Low-heat-transfer interface between metal parts
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US20130089431A1 (en) * 2011-10-07 2013-04-11 General Electric Company Airfoil for turbine system
US9835033B2 (en) 2012-03-26 2017-12-05 United Technologies Corporation Hybrid airfoil for a gas turbine engine
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US20130251536A1 (en) * 2012-03-26 2013-09-26 Sergey Mironets Hybrid airfoil for a gas turbine engine
US9011087B2 (en) * 2012-03-26 2015-04-21 United Technologies Corporation Hybrid airfoil for a gas turbine engine
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GB2504833A (en) * 2012-06-11 2014-02-12 Snecma A method of making a turbine blade with a trailing edge less than 1 mm thick
GB2504833B (en) * 2012-06-11 2016-03-30 Snecma A casting method for obtaining a part including a slender portion
US9962763B2 (en) 2012-06-11 2018-05-08 Snecma Casting method for obtaining a part including a tapering portion
EP2778059A1 (en) * 2013-03-14 2014-09-17 Bell Helicopter Textron Inc. Amorphous metal rotor blade abrasion strip
US20140271214A1 (en) * 2013-03-14 2014-09-18 Bell Helicopter Textron Inc. Amorphous metal rotor blade abrasion strip
US20160167269A1 (en) * 2013-07-29 2016-06-16 Safran Method of fabricating a composite material blade having an integrated metal leading edge for a gas turbine aeroengine
US10899051B2 (en) * 2013-07-29 2021-01-26 Safran Method of fabricating a composite material blade having an integrated metal leading edge for a gas turbine aeroengine
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