US3666453A

US3666453A - Titanium-base alloys

Info

Publication number: US3666453A
Application number: US71331A
Authority: US
Inventors: Richard Ernest Goosey
Original assignee: Imperial Metal Industries Kynoch Ltd
Current assignee: Imperial Metal Industries Kynoch Ltd
Priority date: 1969-09-24
Filing date: 1970-09-11
Publication date: 1972-05-30
Anticipated expiration: 1989-05-30
Also published as: FR2062537A5; GB1297152A; DE2045816A1; JPS4941246B1

Abstract

CREEP RESISTANT ITIANIUM ALLOYS CONTAINING ALUMINIUM, ZIRCONIUM, TIN, SILICON, HAFNIUM AND OPTIONALLY MOLYBDENUM AND BETA STABILISERS.

Description

United States Patent 01 lice 3,666,453 Patented May 30, 1972 3,666,453 TITANIUM-BASE ALLOYS Richard Ernest Goosey, Aldridge, England, assignor to Imperial Metal Industries (Kynoch) Limited, Birmingham, England No Drawing. Filed Sept. 11, 1970, Ser. No. 71,331 Claims priority, application Great Britain, Sept. 24, 1969,

47,049/ 69 Int. Cl. C22c 15/00 U.S. Cl. 75175.5 9 Claims ABSTRACT OF THE DISCLOSURE Creep resistant titanium alloys containing aluminium, zirconium, tin, silicon, hafnium and optionally molybdenum and beta stabilisers.

BACKGROUND OF THE INVENTION (1) Field of the invention This invention relates to titanium-base alloys intended for use in service conditions demanding high creep resistance such as in aircraft engines.

(2) Description of the prior art Titanium alloys for use in the present generation of advanced aircraft engines now have to meet a creep property requirement of less than 0.1% total plastic strain during 100 hours exposure to a stress of 31 hbar. (20 tonf./in. at a temperature of 520 C. At the same time these alloys must possess adequate room temperature tensile strength, i.e. at least 100 hbar. (65 tout/in?) combined with ductility of at least 15% reduction of area and 8% elongation measured on a gauge length of D where D is the diameter of the test specimen. Furthermore, these values must be retained after exposure to service temperatures, since in practice alloys used in aircraft engine components are subjected to thermal cycles between normal atmospheric temperature and service temperatures of at least 500 C. At the same time, creep resistance of a very high order is required and, since shaping is involved, forgeability must be good to avoid any possibility of cracking. Welding is frequently used in fabrication of parts made from such alloys and there must be no embrittlement as a result of local heating to welding temperatures and a high degree of weldability is required.

One alloy which meets these requirements after suitable heat-treatment is one which is known by the designation IMI Titanium 685 which has the nominal composition 6% aluminum, 5% zirconium, 0.5% molybdenum, 0.25%

silicon, rest titanium. This alloy not only meets the mechanical requirements, but is also we-ldable without inducing loss of ductility in the welded area, is resistant to embrittlement by hydrogen pick-up, is readily forgeable and is thermally stable.

The above-mentioned alloy was itself an improvement on another earlier alloy IMI Titanium 684 which advanced the utility of titanium by raising the service temperature to about 520 C. and had a composition of 6% aluminium, 5% zirconium, 1% tungsten, 0.3% silicon. Whilst this alloy possessed excellent high temperature properties, it was found that it was not always sufficiently resistant to embrittlement on exposure to high temperatures. Embrittlement in this context means loss of ductility measured at room temperature before and after exposure to high temperatures.

With so many properties required simultaneously, it is diflicult to achieve large advances in any one property without adversely aifecting other properties. It is not surprising, therefore, that many proposals have been made in the patent literature relating to various combinations of aluminium, tin, zirconium, molybdenum, tungsten, niobium, vanadium and silicon. Each of these proposals was directed to a combination of good high temperature properties but as far as we are aware none have succeeded in raising the service temperature of the alloy whilst at the same time maintaining the other properties discussed herein.

It is becoming apparent to those skilled in the art that titanium base alloys containing some or all of the abovementioned elements are approaching a service temperature limit and further improvements in the service temperature are likely to be small particularly as the behaviour of the above-mentioned elements is well understood.

By introducing an element in the composition different from those mentioned above we have been able to make a further improvement in the class of alloys discussed.

SUMMARY OF THE INVENTION According to the present invention, a titanium-base alloy consists of one or both of the alpha stabilisers aluminium and tin in an amount equal to an aluminium equivalent of 4.5-6.5% by weight, aluminium being present in an amount not less than 2.5% and tin, when present, not exceeding 12%; zirconium, 27%, silicon ODS-0.5%, hafnium 0.55.0%, balance titanium apart from impurities. Optionally the alloy may also contain 0-2% molybdenum.

The term aluminium equivalent refers to the wellknown replacement of aluminium by tin in the proportion of 3% tin for each 1% of aluminium replaced wherein the mechanical properties remain substantially unaffected by such replacement.

DESCRIPTION OF THE PREFERRED EMBODIMENT The alloys of the invention are based on the discovery that a small addition of hafnium in the presence of silicon improves creep resistance and stability Without detriment to the other properties necessary in an aircraft engine a1- loy. Moreover, the improvement in creep resistance and stability enables the alloy to be used in service up to 540 C., an increase of 20 C. on the temperature hitherto regarded as the upper limit of service temperatures of the present titanium alloys available to the aircraft engine industry.

It is believed that the hafnium has the effect of strengthening the alpha phase at elevated temperatures. When hafnium 'is added 'to' alloys lacking silicon, such as titanium, 5% aluminium, 3% tin and 2% zirconium or titanium, 3% aluminium, 6% tin and 2% zirconium, there is little improvement in tensile properties with hafnium contents up to 5% but such alloys have been found to possess a high degree of metallurgical stability. Instability resulting resulting from metallurigcal changes occurring during such exposure is usually monitored by changes in room temperature tensile properties measured before and after exposure. The results of such tests carried out on titaniumaluminium-tin-zirconium bases containing various amounts of hafnium are shown in Table III and indicate that hafnium bestows upon the alloys a high degree of metallurgical stability up to hours at 540 C. The addition of hafnium does not improve the creep strength of such alloys at 540 C. which is inferior to that of known alloys. The presence of both silicon and hafnium, however, produces a considerable improvement as is shown in Table I in which at a test temperature of 540 C., the total plastic strain of an alloy containing silicon is shown to be twice as great as that of a similar alloy containing both silicon and hafnium.

Hafnium is a dense metal and when the intended use of the alloy is in aircraft, the strength/Weight ratio is particularly important. Having regard to this ratio and to the optimum combination of tensile and creep properties, the preferred range is 0.5-3% hafnium.

Molybdenum is beneficial in producing a worthwhile increase in tensile strength without adversely aifecting creep properties. The preferred amount is 0.5% and examples of the improvement in strength are shown in Tables I and II. For maximum improvement in tensile strength and a high degree of weldability the preferred range of molybdenum is 0.25-0.75 However, a further increase in tensile strength without detriment to creep properties for applications in which weldability is not important can be obtained with molybdenum contents up to 2%.

The heat-treated alloys of the invention have better creep resistance at a temperature of 540 C. than the previously known alloys and the properties of these alloys are compared in Table II. Tensile properties before and after creep testing are similar in the alloys of the invention and the prior art alloys indicating that all are of similar strength and are thermally stable, the alloys of the invention being superior in this respect. The diiference in properties is to be seen in the total plastic strain at 540 C. in which the prior art alloys have a much greater strain than the alloys of the invention and are, in fact, considerably in 4 excess of the maximum permitted strain of 0.1% in 100 hours;

Results in the tables are given where appropriate in SI units, the tensile strength values being in hectobars (1 hbar.=l.02 kg./mrn. =0.6-5 tonf./in. The gauge length of test pieces used for elongation determination is 5 times the diameter (El 5D).

With regard to the alpha stabilising elements aluminium, tin and zirconium, the functions of these are well known having been the basic constituents for many known alloys. The limits specified are dictated by the need to provide a strong, ductile alloy which is resistant to embrittlement by hydrogen pick-up, is readily forgeable and weldable without loss of ductility. Although tin is a preferred constituent, it can be omitted, provided that the aluminium content is increased by 1% for each 3% tin. The tin-free alloy has the advantage of lower density.

The alloys of the invention heat-treated for 1 hour at 1050 C., air cooled, reheated for 24 hours at 550 C. and air cooled, provide for the first time in a titanium-base alloy less than 0.1% plastic strain on exposure to a stress of 31 hbar. (20 tout/in?) at a temperature of 540 C. and are particularly suitable for the manufacture of aircraft engine components as the alloys can be readily forged and welded without detriment to ductility.

TABLE I Percent total 0.2% PS Tensile strength plastic El 5D, RA, Addition to 'Il-5 Al-3 Sir-3 Zr strain hbar. (tnnL/infi) hbar. (tout/in?) percent percent 0.25 Si 0. 382 88. 4 (57. 3) (63. 9) 11% 23 0.25 Si plus 2 HI 0. 182 95. 0 (61. 5) 99 3 (64. 4) 13 22 0.25 Si plus 2 lit-0.5 Mo 0. 192 95. 5 (61. 8) 107 6 (69. 7) 9 15 Creep tested for 300 h. at 540 0. under a stress of 23 hbar. (15 tontllinfi).

Norm-All tests on 12 mm. square bar, heat-treated 1 h. 1,050 0., air-cool plus 24 h. 550 0., air-cool.

TABLE II Percent total plastic strain Tensile properties before and alter creep in 100 h. at testing at 540 0. 540 C. under a stress of 0.2% PS ar. hbar. TS libar. E15D, RA Composition (tent/in?) (tonL/infi) (tent/in?) percent percent T14 Al-fi Sn-3 Zr-2 Hit-0.25 Si Unexposed 95. 1 105. 8 954 19 Ti-4 .Al-6 Sn-3 Zr-2 Elf-0.5 Mo0.26 Si Unexposed 93. 7 108.0 11 20 96. 4 10s. 1 12 1M (62. 4) (70. l)

Ti-fi Al-S Zr-0.5 Mo0.25 Si Unexposed 91. 5 105. 0 10 22 Ti-1 Al-G Sn-B Zr-l bib-0.25 Si Unexposed 94. 1 105. 5 12 22 *Fraetured in threaded grips.

NorE.-Post creep duetilitles determined without further surface preparation of test piece. All material beta heat-treated 1 h. 1050" 0., air cooled, aged 24 h. 550 C. (12 mm. square bar).

TABLE III Elong. on 5 X 0.2% proof stress Tensile strength dlam- Reduction Exposure eter, in area, Nominal composition (non-stressed) hbar. (tout/in!) hbar. (tonL/lnfi) percent percent 'Il-5 Al8 Sn-2 Zr-l EL. 85. 1 (55.1) 91. 7 (59.3) 12 22 300 h. 540 0.- 86. 6 (56. 93. (60. 5) 1,000 h. 540 C. 83. 5 (54. 0) 91. 0 (58. 9) 9 21 2 Hi 84. 2 (54. 5) 93. 5 (60. 5) ll 21 300 h. 540 0-- 85.7 (55. 95.0 (61.5) 9 20 1,000 h. 540 C- 83. 7 (54. 2) 92. 0 (59.5) 12 20 5 Hi 87. 5 (56. 7) 98. 0 (63. 4) 11 23 300 h. 540 0.- 88. 2 (57.3) 100.5 (65.0) 12 21 1,000 h. 540 C. 89. 0 (57. 7) 98. 7 (63. 9) 12 17 Ti-3 Al6 Sn2 Zr1 Hf 79. 0 (51. 1) 88. 2 (57. 1) 12 27 300 h. 540 0.- 80.0 (51.8) 87.9 (56.8) 13 27 2 Hi 80. 4 (52.0) 91. 0 (58. 9) 12 22 300 h. 540 0.- 81. 5 (52.8) 91. 7 (59. 3) 12% 5 Hi 80.5 (52. 1) 91. 2 (59. 1) 12 24 300 h. 540 0-- 81.7 (52. 9) 91.5 (59. 3) 12% 23 I claim:-

1. A titanium-base alloy consisting of one or both of the alpha stabilisers aluminium and tin in an amount equal to an aluminium equivalent of 4.56.5% by Weight, aluminium being present in an amount not less than 2.5% and tin when present, not exceeding 12%; zirconium 2- 7%, silicon ODS-0.5%, hafnium 0.55.0%, and option ally, 02% molybdenum, balance titanium apart from impurities.

'2. A titanium base alloy as claimed in claim 1 containing 0.25-0.75% molybdenum.

3. A titanium-base alloy as claimed in claim 2 containing 0.5% molybdenum.

4. A titanium-base alloy as claimed in claim 1 containing 0.5-3.0% hafnium.

S. A titanium-base alloy consisting of 4% aluminium, 6% tin, 3% zirconium, 2% hafnium, 0.25% silicon, balance titanium, apart from impurities.

6. A titanium-base alloy consisting of 5% aluminium, 3% tin, 3% zirconium, 2% hafnium, 0.25% silicon, balance titanium, apart from impurities.

7. A titanium-base alloy consisting of 5% aluminium, 3% tin, 3% zirconium, 2% hafnium, 0.5% molybdenum, 0.25 silicon, balance titanium, apart from impurities.

8. A titanium-base alloy consisting of 4% aluminium,

6% tin, 3% zirconium, 2% hafnium, 0.5% molybdenum,

0.25% silicon, balance titanium, apart from impurities.

9. An aircraft engine component made from a titanium-base alloy consisting of one or both of the alpha stabilisers aluminium and tin in an amount equal to an aluminium equivalent of 4.5-6.5 -by weight, aluminium being present in an amount not less than 2.5% and tin when present, not exceeding 12%; zirconium 2-7%, silicon 0.050.5%, hafnium 0.55.0%, and optionally, 0-2% molybdenum, balance titanium, apart from impurities.

References Cited UNITED STATES PATENTS