Metal matrix composites for aircraft propulsion systems

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00 oo ~I (NASA-TM-X-71685) METAL MATRIX COMPOSITES N75-19245' FOR AIRCRAFT PROPULSION SYSTEMS (NASA) 24 p HC $3.25 CSCL 21E Unclas G3/07 13396

METAL MATRIX COMPOSITES

FOR

AIRCRAFT PROPULSION SYSTEMS

by Robert A. Signorelli Lewis Research Center Cleveland, Ohio 44135 TECHNICAL PAPER to be presented at International Conference on Composite Materials Geneva, Switzerland, April 7-11, 1975 Boston, Massachusettes, April 14-18, 1975 METAL MATRIX COMPOSITES FOR AIRCRAFT PROPULSION SYSTEMS Robert A. Signorelli NASA-Lewis Research Center Cleveland, Ohio I NTRODUCTI ON Studies of advanced aircraft propulsion systems have indicated that performance gains and operating costs are possible through the application of metal matrix composites.

Compressor fan blades and turbine blades have been identified as components with high payoff potential as a result of these studies. This paper will present the current status of development of five candidate materials for such applications.

Boron fiber/aluminum, boron fiber/titanium, and silicon carbide fiber/titanium composites are considered for lightweight compressor fan blades. Directionally solidified eutectic superalloy and tungsten wire/ superalloy composites are considered for application to turbine blades for use temperatures to 1100 C (2000 F).

POTENTIAL COMPOSITE BLADE BENEFITS AND PROBLEMS As with most new materials, composites are in competition with standard or conventional materials for component applications.

Designers usually design a component using proven current materials and substitute the new material only after studies have been undertaken to develop experience and confidence. While composite hardware programs have been largely substitutional, engine system studies have been performed to indicate the potential advantages of composites when incorporated in the design stage. Fan and turbine blades have been identified in these studies as components with high payoff potential. The studies, summarized in ref. 1, have indicated advantages as shown in figure 1. This figure shows the benefits accrued by an advanced engine configured to use the higher strength and stiffness potential of composite fan blades and the higher creep- rupture strength and higher use temperature potential of superalloy-matrix composite turbine blades. Comparison is made with an equivalent engine using current materials.

The redesigned engine Involves fewer fan and compressor stages, a reduced number of turbine stages, a reduced number of bearings and a shortened overall length. A smaller and lighter engine is therefore possible through the use of composites.

However, there are some technical problems to over- come before composite component reliability is estab- lished and such an engine can be built. The metal matrix composite candidates for application to fan and turbine blades as discussed in this paper are shown in table I.

The table also shows the major problem areas under study to ready these materials for engine component application.

Fan Blades B/Al Impact Resistance More research and development effort has been de- voted to boron/aluminum (B/Al) for fan blade application than any other metal matrix composite. This composite has been developed to a state of readiness very near that required for successful full scale engine operation of a fan blade set in ground engine tests. The ground engine test, which would include evaluation of resist- ance to foreign object damage, would serve to prepare the material for flight demonstration.

The performance and properties demonstrated by B/Al in studies conducted thus far have been adequate for fan blade application with one exception. Resistance to large object impact, such as birds, has varied greatly.

Resistance to erosion from sand and rain has been ade- quate, as has been impact resistance to gravel and ice balls. Large object impact (such as birds) usually caused fracture of large portions of the airfoil. Poor impact performance is not unexpected based on comparison of pendulum impact strength values shown in figure 2 for titanium (ref.

2) and typical values for R/Al composites used in the initial studies (ref. 3). Since the compos- ite impact values were less than half of the value for titanium, R/Al blade failure from impact is not surprising. Improved impact resistance is thereforr required to apply R/Al to blades in the forward section of turbofan engines.

Studies to understand and improve impact resistance have been conducted at several laboratories (refs. 3-8).

The results discussed below are from programs conducted at the Lewis Research Center and at TR! under NIASA funding. The variables included in the programs are shown in table II. The choice of variables was influ- enced by efforts to minimize the embrittling effect of the low strain-to-failure boron fiber on the otherwise relatively ductile, tough aluminum alloy.

,atrix alloys were chosen to further increase ductility and failure strain. Large diameter fibers were selected to increase the interfiber distance in the composite, thereby reducing the volume of matrix constrained by the elastic fibers. Fabrication processing was varied to obtain densification of the composite and a high degree of bonding between ply layers. However, bonding temperatures were maintained as lov, as possible to minimize reaction at the fiber-matrix interface since reaction has been shown to reduce the mechanical properties of R/AI.

A significant improvement in impact resistance of B/Al was obtained using these approaches. Sore of the improvements are illustrated in figure 3. flatrix alloys 5052 and 1100 were selected because they have lower yield strengths and larger tensile strain-to-failure values.

Increasing the ductility of the matrix by se- lecting a different alloy increased the impact strength of the composite.

NJotched Charpy values increased from 18 joules (13 ft-lhs) with 5052 Al alloy matrix to joules (48 ft-lbs) with 1100 Al matrix. Increasing fiber diameter resulted in a further significant improvement in impact to 96 joules (71 ft-lbs).

The impact resistance was also strongly influenced by processing conditions.

The bonding temperature which contributed to high impact strength varied with the matrix alloy used as well as the fabrication process.

It does appear that the lowest bonding terperature, at which good bonding can be achieved, is generally the most desirable.

Previous work (ref. 9) has shown that fatigue resistance can be degraded by increasing bonding temperature from 450 to 475 C. The fatigue strength reduction was related to an increase in the reaction ORIGINAL PAGE IS OF POOR QUALITY zone at the B/Al Interface. An increase in the reaction zone might also be expected to reduce fiber strain to failure and thereby the energy absorption capability.

Decreased impact strength can also result from too low bonding temperature. Too low a temperature can result in inadequate bonding. Pelamination between monotape plies or at the fiber/matrix interface can then occur prematurely at relatively low applied stress levels.

Bonding temperatures between 455 and 482 C (850 - 900 F) appear to give high impact values.

The marked impact improvement in impact resistance of R/Al is Illustrated in figure 4 by comparison with a typical titanium fan blade material and previous B/A1 composite impact values. A ten-fold increase in laboratory pendulum impact values is very encouraging but not necessarily indicative of satisfactory foreign object damage resistance at the high velocity of fan blade operating conditions. However, high velocity ballistic impact tests on static test specimens of Improved R/Al have indicated similar trends of increased impact resistance (ref. 8).

Improved impact resistant R/Al is now being applied to fan and compressor blades and preliminary results are encouraging. The recent successes with B/Al composites suggest that FOD resistance, the remaining major technical Impediment to successful R/Al fan blade flight demonstration may be overcome.

B/Al Fan Blade Cost Considerations One of the more difficult tasks involved in ap- plying composite materials is to estimate realistic cost benefits. The labor-intensive research and development fabrication techniques commonly used for composites, combined with limited production experience, form a questionable basis for cost projections. However, R/Al has been studied for about ten years and a number of detailed cost projections have been made (ref. 1). The cost projections shown in table III are based upon boron filament and tape costs that have been projected by representatives of the producers and require no new technology. The $250/,kg price is fairly close to the current large order price. At that B/Al cost, the fan blade cost is equivalent to that for conventional forged titanium blades.

The cost benefits that may be achieved by lower cost components are often the primary justification or mandatory prerequisite to applications of new technology in the harsh economic climate that ORIGINAL PAGE IS OF POOR QUALITY prevails today. As shown in table III, it is antici- that reductions in the cost of R/Al will permit pated blade costs substantially lower than that of forged t I tan urm.

R/TI and SIC/Ti Titanium matrix composites also have been developed for application to fan blades. Filament reinforcement material would seem to be a of the standard fan blade logical candidate for intensive development as a fan However the severe fiber degradation encountered blade.

during the high temperature exposure required for fabri- cation has discouraged early research efforts (refs.

10-11). Although the level of titanium matrix composite research has been limited, some progress has been made table IV (ref.

as indicated by the properties listed in 1). The high shear strength and transverse properties, combined with the relatively good retention of strength to 538 C (1000 F), make the material a good candidate for elevated temperature blade application in turbojet engines. The erosion resistance of titanilm to ice, sand, and rain for fan blade applications has been well established by years of service.

Another important advantage of titanium matrix composites for elevated temperature applications is the thermal expansion match of matrix and fiber. Thermal fatigue failure can be the limitirg failure mode for the type of cyclic operation that is typical for fan blades. The "typical" operating cycle for aircraft engines in commercial service in the U. S. is estimated to be one hour with about 3000 hours of operation each year. The 3000 heating and cooling cycles per year can cause failure of turbine blades and could conceivably cause problems in the case of high temperature fan and compressor blades.

The high shear and transverse properties of titanium matrix composites may also make angleply orientation unnecessary thereby simplifying fabrication and reducing cost.

lowever, this class of materials is not without problems. In addition to the fiber degradation caused by high fabrication temperatures, the density of titanium composites is higher than that of R/Al.

The specific modulus values of both types of composites are about equivalent, but the specific tensile strength values are lower than those of R/AI. Very limited impact data have been published; however the notched Charpy impact values reported (ref. 12) were 7 joules (5.2 ft-lbs) or lower. Notched unreinforced titanium PAGE IS ORIGINAL 5 OF POOR, QUALITYi impact values are about 20 joules (15 ft-lbs). Titanium an impact improvement matrix composites may benefit from study similar to that conducted with B/Al. Since unnotched titanium specimens have demonstrated pendulum Indi- impact strengths over 135 joules (100 ft-lbs), as cated in ref. 8, it might also be expected that titanium composites would have the potential for higher impact strengths than those obtained thus far.

Titanium composite fan blades have been fabricated with both boron and SIC fibers and cost analysis pro- jections for large scale production (ref. 13) indicate that they can be cost competitive with titanium alloy blades.

Turbine Blades Directionally Solidified Eutectics engine cycle studies have indicated Gas turbine that significant benefits In cost, size, weight, and performance are possible through the use of advanced are composite technology. Higher cycle temperatures beneficial and superalloy turbine blades typically are cooled to permit higher turbine Inlet temperatures than would be possible otherwise. Cycle studies also indi- cate that higher combustion temperatures are countered need to decrease exhaust pollutants such as by the nitrous oxide and the need to reduce engine noise gener- ation. Both noise and pollution emissions are generally of the ways more severe with higher temperatures. One that composite turbine blade materials can be used more efficiently is to decrease the amount of cooling now required in the higher temperature forward turbine rear stage blades, which are stages. In addition, the now cooled a small amount, might be operated without cooling entirely. The higher temperature capability of could make such an overall reduction In composites perform- cooling air possible. Another way to increase engine weight would be by increasing ance and decrease the rotational speed of the engine. The higher strength of the composite blades could permit such an Increase, which in turn could decrease engine size and weight as in indicated in figure 1. The increased payoff obtained line with the above discussion would not increase noise or pollution since neither the combustion temperature the jet exhaust velocity would be increased.

nor (In-situ Directionally solidified (d. s.) eutectics composites) are being developed for the first generation eutectic systems have turbine blades. Two composite Micrographs for development (refs. 14-17).

been selected are shown in figure 5.

of rod and lamellae reinforcement these composites is formed reinforcement in both of The are directionally solidified in-situ when the alloys planar front stringent conditions to achieve under the the reinforcement aligned in solidification with direction.

growth d. s. eutectic alloys Rupture properties of three for one of the figure 6, along with data are shown in VIA alloy.

conventional superalloys, TRW-NASA strongest at are superior to the superalloy All of the eutectics F). There are F) and 1090 C (2000 both 980 C (1800 two nickel-base dispersion differences between the minor shown, which are currently strengthened eutectics for turbine blade considered leading candidates (ref. 17).

applications stable for long times at While d. s. eutectics are gradients, thermal cycling or thermal high temperatures, under normal service both of which are encountered can cause in gas turbine engine operation, conditions of This has caused degradation structural instability.

at compositions. However, properties for some composite Nb reinforced of TaC and NI least some compositions 3 resistance to thermal eutectics have demonstrated cycling (ref. 18).

and to increase shear strength Studies are underway exposed Oxidation protection for transverse ductility.

as is the case for conven- surfaces is also required, in turbine blade applications.

tional superalloys used defect free for the long Coatings do not remain for turbine 3000 - 8000 hours) expected lives (typically to provide a margin of oxl- blades and it is important Therefore, in the uncoated composite.

dation resistance to provide adequate oxidation the relative ability factors in may be one of the deciding protection be chosen composition will eventually determining which for blade use.

in developing d. s.

Blade cost is also a factor production of for blade application. While eutectics (refs. 15- blade shapes has been demonstrated complex 2 - 10 cm/hr and careful growth rates are generally 16), are need to assure safe control and inspection quality of d. s. eutectics is generally blades. The ductility they are intended to than that of the superalloys lower ductility the acceptance of reduced replace. However, trend.

designers Is a continuing by gas turbine engine because were regarded with skepticism Cast superalloys forged and variable properties when of lower ductility The blade material.

superalloys were the standard SIC and being conducted to use effort currently sizable a good Indication in turbines is other brittle ceramics to are still some problems trend. While there of this of d. s.

working In the field overcome, researchers these that turbine blades of are confident eutectics turbine engines in the will be operated In gas materials foreseeable future.

Refractory Wire Superalloys have composites also wire/superalloy Refractory application to turbine subject of study for been the at a number of have been Investigated blades. They composite - 24). This type of (refs. 19 laboratories above at temperatures potential for application has the currently being developed, those of the d. s. eutectics development and application in figure 7. The as shown could permit composites of refractory wire/superalloy without as 1150 C (2100 F) use temperatures as high 1200 C and as high as barrier coated fibers diffusion strength and barriers. Rupture (2300 F) with diffusion with laboratory have been demonstrated impact resistance data (refs. 22 and 2).

F) data at 1090 C (2000 rupture Density-normalized refractory The strengths of the are shown In figure 8.

greater are up to four times composites wire/superalloy twice those superalloys and than those of conventional Increases of Further strength the d. s. eutectics.

for to increase are possible refractory wire/superalloys for diffusion strength advantage. The potential this as hlih as calculated to be coated wire has been barrier values for the density-normalized four to six times from at 1090 C (2000 F).

d. s. eutectics fabricated have been Refractory wire/superalloys strengths and miniature Izod impact that have Charpy those obtained with super- that compare favorably with few data However, blade temperatures.

alloys at normal to oxidation, to indicate resistance have been obtained Further- mechanical fatigue.

and thermal and erosion, have been most programs evaluated in the specimens more, are not metallurgy methods which using powder fabricated production. Diffusion suited to low cost blade ideally tapes of refractory wire/superalloy bonding of monolayer adds production and also be preferred for volume would R to the potential for improved properties. Improved ductility can be incorporated by using wrought superalloy foils instead of hot pressed powders. Also, fully densified superalloy foil may reduce fiber/matrix reaction.

Monolayer tape fabrication of refractory wire/ superalloy composites has been developed using the process shown schematically in figure 9, taken from ref.

25. Roth powder cloth and alloy foil have been used successfully. The high cost of obtaining small quanti- ties of foil of a number of research alloy compositions was avoided by using powder cloth. However, it Is envisioned that alloy foil will be more efficient for much volume production. Press bonding has been used for of the research effort, becuase of the ease of producing small research quantities with varying compositions.

However, for large volume production, the effort to develop the proper conditions of pressure, roll speed, and temperature can be justified to achieve rapid, high volume, low cost production of monolayer tape. The processing of turbine blades fabricated using this method should be similar in labor and processing and therefore the costs should approach the fan blade costs described earlier in the paper.

The micrograph in figure 10 shows the excellent quality of densificatlon and the lack of fiber/matrix reaction obtained with press bonded monotape. Similar quality monotape was achieved with a limited amount of roll bonded tape. Even more significant was the limited depth of reaction which occurred at the fiber/matrix interface with test specimens fabricated by secondary diffusion bonding of monolayer tape. The micrographs in figure 11 show the limited depth of reaction (0.001 cm) obtained after 200 hours at 1190 C (2000 F). This reaction depth is about that obtained with as-fabricated powder metallurgy specimens. Previous studies (refs. 21, 22, and 26) have shown that composite properties corre- late with reaction depth. Since increased reaction depth results in reduced properties, the limited reaction should result in retention of high rupture strengths.

The data currently most needed are concerned with the performance of refractory wire/superalloy composites in cyclic thermal fatigue tests. Very few tests have conducted in the past and results have been mixed.

been The concern stems from the large thermal mismatch of nickel superalloys and tungsten wire. Thus, the thermal fatigue resistance of refractory wire/superalloy compos- ites must be demonstrated for them to be considered candidates for use as turbine blades in aircraft turbine engines. Land based turbines and other applications that do not 'have short operating cycles would require less resistance to thermal fatigue.

Although limited in amount, the thermal fatigue data obtained recently for refractory wire/superalloy composites are very encouraging. The photos in figure 12 show a specimen tested for 1000 cycles from 85 to 2200 F in a Naval Air Systems Command program conducted at TRW. The specimens were direct-resistance-heated in a Gilmore Universal testing machine. No matrix or fiber cracking was apparent in the W-1 ThO /FeCrA1Y composite.

These preliminary results suggest that thermal fatigue failure of refractory wire/superalloy composites may not be a serious problem.

CONCLUDING REMARKS Significant progress has been made In the develop- ment of composites for application to aircraft propul- sion systems. The progress is particularly noteworthy because it has been achieved in an enviroment of cost competitiveness in which acceptance of new technology is difficult.

Boron/aluminum has demonstrated a marked improve- ment in impact resistance. These results give hope that FOP resistance, the remaining technical impediment to flight demonstration of B/Al blades can be overcome.

Efficient, automated blade fabrication procedures have been identified which have the potential to produce fan blades at equivalent or lower cost than forged titanium blades.

Titanium-matrix composites with boron or silicon carbide reinforcement have demonstrated good properties at room and elevated temperatures. These properties indicate that titanium-matrix composites also should be candidates for fan blade applications for elevated temperatures.

Directionally solidified eutectics are being considered for turbine blade applications at blade material temperatures at least 50 C (100 F) above those currently used for conventional superalloys.

Compositions have been identified and mechanical property data have been obtained to indicate a good potential with d. s. eutectics for achieving turbine blade use temperature increases within these goals.

IS PAGE ORIGINAL QUALITY OF POOR composites are wire reinforced superalloy Tungsten developed for application at blade use being F) above those of temperatures of at least 100 C (200 the first generation of directionally solidified of refractory wire/ eutectics. Fabrication processing bonding superalloy composites from secondary diffusion been developed. Test specimens of monolayer tape has reduced fiber/matrix made from monolayer tape have shown stress-rupture properties compared reaction and improved metallurgy to previous composites made by a powder Preliminary thermal fatigue results indicate method.

may not be a serious problem.

that this failure mode Is bright for application of metal The outlook systems, as matrix composites to aircraft propulsion aerospace systems.

well as to other REFERENCES Status and Prospects.

1. Metal Matrix Composites: Board, Publication National Materials Advisory NMAB-313, Dec. 1974.

2. ASM Metals Handbook. ASH, Vol. 1, 1961.

3. Krelder, K. G.; and Prewo, K. M.: Boron-Reinforced 4, Metallic Aluminum. Composite Materials, Vol.

Matrix Composites. Academic Press, 1974, pp.

400-480.

Swanson, G. O.: Toughness of 4. Hancock, J. R.; and Filamentary Boron/Aluminum Composites. Composite and Design, Second Conference.

Materials: Testing ASTM STP 497, ASTH, 1972, pp. 299-310.

Energy of Boron- 5. Prewo, K. M.: The Charpy Impact Vol. 6, 1972, pp. 442.

Aluminum. J. Comp. Matls., 6. Olster, E. F.; and Jones, R. C.: Toughening Mechanisms in Continuous Filament Unidirectionally Composite Materials: Testing Reinforced Composites.

Design, Second Conference. ASTM STP 497, ASTM, and 1972, pp 189-205.

R.; and Allred, R. E.: Fracture Toughness 7. Hoover, W.

with Weak Fiber-Matrix of Borsic-Al Composites 1972.

Bonds. Sandia Laboratories, SC-DC-714467, Impact 8. Melnyk, P.; and Toth, i. J.: Development of Resistant Boron/Aluminum Composites for Turbojet Engine Fan Blades. NASA CR-134770, Nov. 1974.

9. Hancock, J. R.; and Shaw, G. G.: Effect of Filament- Matrix Interdiffusion on the Fatigue Resistance of Boron/Aluminum Composites.

Composite Materials: Testing and Design, Third Conference. ASTM STP 546, ASTM, 1974, pp 497-506.

10. Klein, M. J.; Reid, M. L.; and Metcalfe, A. G.: Compatibility Studies for Viable Titanium Matrix Composites. AFML TR 69-242, Oct.

1969.

11. Collins, R. R.; Brentnall, W. D.; and Toth, I. J.: Properties and Fracture Modes of Borsic/Titanium.

AFML TR 73-43, 1973.

12. Prewo, K. M.; and Kreider, K. G.: The Deformation and Fracture of Borsic Reinforced Titanium Matrix Composites. Titanium Science and Technology, Vol.

4. Plenum Publishing Corp., 1974, pp. 2333-45.

13. Toth, I. J.: Metal Matrix Composite Blade Fabrication Methods.

Proc. 19th SAMPE Symposium, New Industries and Applications for Advanced Materials Technology.

SAMPE, April 1974, pp.

406-16.

14.

Weeton, J. W.: Fiber Reinforced Superalloys, Ceramics and Refractory Metals and Directionally Solidified Eutectics (Heat Resistant Composites).

Composites - State of the Art. AINIE, 1974.

15. Ashbrook, R. L.: Directionally Solidified Composite Systems Under Evaluation.

NASA TM X-71514, April 1974.

16. Thompson, E. R.; and Lemkey, F. D.: Directionally Solidified Eutectic Superalloys.

Composite Materials, Vol. 4, Metallic Matrix Composites.

Academic Press, 1974, pp. 102-156.

17. Harrison, R. W.: Private Communication, Aircraft Engine Group, General Electric Co., Cincinnati, Ohio, 1974.

18. Lawley, A. L.: The Mechanical Behavior and Structur- al Stability of Eutectic and Fiber Reinforced Composites. Proc.

19th SAMPE Symposium, New Industries and Applications for Advanced Materials Technology. SAMPE, April 1974, pp.

429-36.

19. Dean, A. V.: The Reinforcement of Nickel-Base Alloys with High Strength Tungsten Wires. Rept. R-206, April 1965.

Nat. Gas Turbine Estab., England, 20. Baskey, R. H.: Fiber Reinforcement of Metallic and Nonmetallic Composites. ASD TDR 63-619, July 1963.

21. Petrasek, D. W.; Signorelli, R. A.; and JWeeton, J.

W.: Refractory Metal-Fiber - Nickel-Base Alloy Composites for Use at High Temperatures. NASA TN D-4787, 1968.

22. Petrasek, D. W.; and Signorelli, R. A.: Preliminary Evaluation of Tungsten Alloy Fiber/Nickel Base Alloy TN Composites for Turbojet Engine Applications. NASA D-5575, 1970.

23. Ahmed, I.; Bananno, J. M.; Loomis, K. E.; and Hefferman, W. J.: Metal Matrix Composites for High Applications. Rep. WVT-7155 (AD 734304), Temperature Watervliet Arsenal, Oct. 1971.

24. Signorelli, R. A.: Review of Status and Potential of Tungsten Wire/Superalloy Composites for Advanced Gas Turbine Engine Blades. NASA TM X-2599, 1972.

25. Brentnall, W. P.; and Toth, I. J.: Fabrication of Wire Reinforced Nickel-Base Alloy Tungsten Composites. NASA CR-134664, Oct. 1974.

26. Petrasek, D. W.; and Signorelli, R. A.: Stress- Rupture Strength and Microstructural Stability of Tungsten-Hafnium-Carbon Wire Reinforced Superalloy Composites. NASA TN D-7773, Oct. 1974.

ENGINE COMPONENTS FOR COMPOSITE - PROBLEM AREAS TABLE I.

PROBLEM AREAS MAJOR METAL MATRIX COMPOSITE COMPONENT IMPACT RESISTANCE FAN BLADES BORONIALUMINUM BLADE COST CHARACTERIZATION PROPERTY BORONITITANIUM RESISTANCE IMPACT SILICON CARBIDEITITANIUM OPTIMIZATION

fPROPERTY

TURBINE BLADES SOIDIFIED SERVICE SIMULATION TESTS DIRECTIONALLY EUTECTICS IcoSTS FABRICATION PROCESSING WIREISUPERALLOY REFRACTORY ITHERMAL FATIGUE TESTS Co TABLE II. - VARIABLES FOR BIAl IMPACT IMPROVEMENT STUDY FIBER DIAMETER MATRIX ALLOY DUCTILITY FABRICATION PROCESSING FIBER PLY ORIENTATION - FAN BLADE COST TABLE III.

BORONIALUMINUM TITANIUM B/Al TAPE AT $2501KG ($1151LB) $1501BLADE $1501BLADE AT $551KG ($25/LB) BIAl TAPE $1001BLADE COSTS BASED ON: FIRST STAGE BLADE 25 CM LENGTI 000 BLADES PER YEAR FOR 6.25 CM CHORD 10 5 YEARS OF BORONITITANIUM AND SILICON TABLE IV. - TYPICAL PROPERTIES CARBIDE/TITANIUM (Ref. 1).

Shear strength Elastic modulus Material Test Tensile Longitudinal Transverse strength tem- Longitudinal Transverse system 2 2 msi MNIm2 ksi MNmrn ksi GN/1m msi GN/m pera- MNmrn ksi ture.

o C 26 --- --- 455 66 241 35 179 Rm. 1241 180 BITi viO 100 pmB 75 248 36 200 29 496 SiC coated Rm. 1310 190 517 165 24 --- --- 345 50 207 30 BITi 538 827 120 vlo (10000 F) 100 pmB 75 262 38 207 30 448 Rm. 1207 175 517 SiCITi 221 32 172 25 262 538 1034 150 345 50 50 vlo 100 pm SiC BASELINE INE ENG 1 3-STAGES 10-STAGES 4-STAGES 2 I / / TURBINE BEARING BEARING FAN COMPRESSOR \ COMPARTMENT 0 2-STAGES 5-STAGES 2-STAGES 1 COMPOSITE ENGINE -- SHORTER LENGTH - ] Figure 1. - Benefits from composite turbine blades.

N 1- 50 VIO 50 VIO 63 V0I TITANIUM ALLOY 100 pmB 100 pmB 145 pmB AI IN 6061 AI REF. 2 IN 6061 Al IN2024 REF.

alloy and early Figure 2. - Impact strength of titanium composites.

BIAI 18 JOULES (13 FT-LB) JOULES (47 FT-LB) 50 VIO 0.14 MM B IN 5052 A1 (UNIDIRECTIONAL) 50 V/O 0.14 MM B IN 1100 Al (UNIDIRECTIONAL) (UNIDIRECTIONAU 96 JOULES (71 FT-LB) 50 V/O 0.2 MM B IN 1100 AI (UNIDIRECTIONAU Figure 3. - Improved BIAl impact resistance.

60- 80- 45- 60 C, o - UNIDIRECTIONAL TITANIUM IMPROVED BIAl ALLOY UNIDIRECTIONAL BIAl 4. - Improved BIAI impact resistance.

Figure dQ - Cr Ln cL Z o le CD 400 - 980P C (18000 F) ° F 10900 C (2000 F) 45- 300- 30 - 200 V)) nO I- L15 - 0- VIA TaC Ni Nb TaC Ni-Co-Cr-Al SUPER- Co-Ni-Cr Ni-Cr-A ALLOY Figure 6. - 100 hour rupture strength of D.S. eutectics (ref. 15L FSUPERALLOYS COMPQSITES S 2.TUNGSTEN-WIRE- S 0 EUTECTICS \SUPERALLOY COMPOSITES 40 L . - BLADE CONDITIONS I I.5 20 - 1 900 950 1000 1050 1100 1150 1200 1250 1300 TEMPERATURE, OC I I I I 2100 2300 1700 1900 TEMPERATURE, OF CS-64485 1000-hour life.

Figure 7. - Potential blade use temperatures for AT 10900 C (20000 F) 100x10 3 2. 5x10 ' . 80 a2.0 1.5 RANGE , 60 - 7 T 40 - 1.0 L .5 0 0 SUPER- EUTECIICS 218 W Th02 W-H-C ALLOYS REINFORCEMENT MATERIAL - %-REFRACTORY- 70-VOL WIRE - SUPERALLOY CS-64490 COMPOSITES of rupture properties 8. - 1000-hour stress Figure - superalloy composites.

refractory-wire ALLOY FOIL MAT FILAMENT CLOTH POWDER FILAMENT SMAT ROLLINOF MONOLAYER ALOY FOIL / STACK-UP POWDER CLOTH FILAMENT - POWDER

CLOTH

- HEAT rIN ENCAPSULATION 16" DIA. WATERBURY DIE HEATED ROLL - FARRELL PRESUREAT L PRHA FURNACE FURNACE ENCAPSULASHUTTLE \ - ENCAPSULATED MECHANISM MONOTAPE - ASSEMBLY BONDING PRESS BONDING ROLL the manufacture techniques for bonding of diffusion Flow diagram Figure 9. - W fiber-Ni alloy matrix monotapes.

of (Ref.25.).

wire/superalloy monotape.

Figure 10. - Tungsten IS PAGE ORIGINAL QUALITY OF POOR N GO eK II -I TOTAL REACTION ZONE 0.001 CM (0.0004") I I ( ) 10900 C 20000 F STRESS RUPTURE TEST. (REF.25).

200 HOUR - Figure ll. - Fiber-matrix reaction of tungsten wire/superalloy monotape.

X10 22000 F).

1000 CYCLES 30 - 12000 C (85 - superallov composite.

of thermally cycled tungsten wire reinforced Figure 12. - Photomicrographs of Irving Machlin).

(Photos courtesy NASA-Lewis-