Aerospace Materials: Chapter 3. Materials For Supersonic Civil Transport Aircraft

• 1. Chapter 3 Materials for supersonic civil transport aircraft Yann Barbaux and Jacques Cinquin Introduction The consequences of the drastic economical and technical requirements for future supersonic civil transport [1] on the materials selection for the dierent parts of the aircraft structure have been detailed and discussed elsewhere [2]. As a result, Aerospatiale, BAe and DASA decided to increase their eort on materials studies and to launch specic research programmes on aluminium alloys and carbon bre reinforced polymers (CFRPs). Major research programmes were initiated on aluminium alloys in 1992 [3] and on organic matrix composites in 1994 [4]. Aluminium alloys The work programme of recent research was divided into two main tasks, corresponding to the study of the two factors assumed to inuence directly the creep resistance and the thermal stability of metals: . Task 1: selection of the main precipitation system . Task 2: optimization of the chemical composition and of the process para- meters. The critical analysis of existing data resulted in the selection of 33 chemical compositions, from the four alloy systems given in table 3.1. These alloys were direct chill (DC) cast and rolled down to 14 mm thick plates and 1.6 mm thick sheets on laboratory equipment at DERA, British Aluminium and Pechiney. They were then tested for creep, thermal stability and corro- sion. Based on the results obtained on these alloys, a selection of 14 dierentCopyright 2001 IOP Publishing Ltd

2. Table 3.1. Selected alloy systems. Alloy system Main precipitation system Al-Cu (2001 type)0 (Al2 Cu) Al-Cu-Mg (2024 or 2618 type) S 0 (Al2 CuMg) Al-Mg-Si-Cu (6013 type)0 (Al5 Cu2 Mg8 Si7 ) 3. 0 (Mg2 Si) Al-Li-Cu-Mg0 (Al3 Li)T1 (Al2 CuLi) compositions from the S 0 and 0 4. 0 precipitation systems was made on which the eect of minor alloying element and thermo-mechanical process variations was studied. Results obtained in this project [5] were very satisfactory: all the alloys tested presented a creep behaviour and a fracture toughness much improved as compared with CM003 alloy (enhanced 2618), which was, at that time, the best reference in terms of creep/damage tolerance compromise. This is illustrated by gure 3.1 and table 3.2, which present respectively creep results in accelerated conditions and fracture toughnesses on compact tension specimens of three of the alloys (labelled A1, C1 and D6), in comparison with creep results from CM003 and fracture toughness results from 2024. Creep life times were extended by a factor of up to seven under dierent 5CREEP ELONGATION (%) STRESS: 250 MPa TEMPERATURE: 175C 4A1C1D6CM003 3 2 1 0 050 100 150 200250TIME (HOURS) Figure 3.1. Creep curves at 1758C/250 MPa.Copyright 2001 IOP Publishing Ltd 5. Table 3.2. Fracture toughness from R-curves on 400 mm widecompact tension specimens. AlloyKc (MPa/m2 )Kc0 (MPa/m2 ) A1 T6 (mod. 2650)125 90 C1 T6 (mod. 6056)160110 D6 T6 (Al-Cu-Mg-Ag)158110 2024 T3130 90 creep test conditions (including 1508C/250 MPa) compared with CM003, and fracture toughness values from R-curves were equal to or better than the damage tolerant 2024 T3 reference. Industrial sheets and plates from the two most promising compositions, a modied version of 2650 Al-2%Cu- Mg alloy and an optimized version of 6056 alloy, entered an exhaustive evaluation programme in 1997, and the results conrm the improvement in creep behaviour over CM003, although the benet is reduced compared with the laboratory tests.In parallel with the development of improved alloys, Aerospatiale has also started studies on the interactions between creep and fatigue on notched coupon specimens and on specimens representative of technological details such as pocket recess or assemblies. These studies are based on the develop- ment of two parallel methods: . a modelling approach combining thermo-elasto-plastic nite elements and physical/metallurgical prediction of creep damage . an experimental approach with the development of specic test equipment capable of reproducing close to real exposures on technological specimens. The results obtained show a slight detrimental eect of 5000 and 10 000 hours of creep exposure at 1308C on the fatigue behaviour of notched specimens in 2650 alloy. Carbon bre reinforced polymers It has already been published [2] that composite materials with carbon bres and polymeric matrices are candidates to achieve the required weight savings on future supersonic civil transport. The main requirements are acceptable properties regarding subsonic ight specications (i.e. damage tolerance), and thermal stability in supersonic ight conditions. Dierent types of matrices are under investigation for Mach 2.05 applications with IM or HR bres, including second generation epoxy, cyanate-based, thermoplastic and bismaleimides, as shown in table 3.3. The work programme was divided into two main research areas:Copyright 2001 IOP Publishing Ltd 6. Table 3.3. Candidate polymer matrices for carbon bre reinforced polymers.State of knowledge Second generation epoxyWell-known processExpected service temperature: 1208C Cyanate based systemsProcess similar to epoxy systemNew products on the marketExpected service temperature: 1508C Thermo-plastic Potential hot forming processExpected service temperature: 1808C BismaleimidesProcessing generally with post-curingExpected service temperature: 1808CLow damage tolerance compared with 2nd generation epoxy . the inuence of long-term thermal ageing on carbon bre reinforced poly- mer physical and mechanical properties . the long-term behaviour of carbon bre reinforced polymers under complex thermo-mechanical loading. The inuence of thermal ageing has been studied by isothermal ageing up to 4000 hours at dierent temperatures from 1008C to 1808C, and by thermal cycling over the ranges 508C to 1208C or 1808C for up to 1000 cycles. The cumulative time at the maximum temperature for 1000 thermal cycles is equivalent to 4000 hours under isothermal conditions. Dierent properties have been investigated after these thermal ageing exposures, such as lled hole compression, compression after impact, glass transition temperature and microstructure. The inuence of thermal ageing on the mechanical properties can be related to the degree of curing of the matrix, and also to the chemical type of the matrix. Figures 3.2 and 3.3 clearly show a post-cure eect on the second generation epoxy, not fully transformed during the initial curing. For the cyanate system, there is a real mechanical property degradation when isothermal ageing is performed above 1608C. Figures 3.3 and 3.4 show that the mechanical property degradation, or the post-cure eect, appears during the rst 1000 hours of ageing. These rst tests were performed at a higher temperature than the service tem- perature, corresponding to Mach 2.05, in order to obtain in a short time the rst indications of thermal ageing response for the dierent families of matrix.Another important point is the inuence of thermal ageing on the damage tolerance properties of the carbon bre reinforced polymers. The results, shown in gure 3.5, have been obtained with compression after impact (CAI) tests performed on bismaleimide composites. Dierent para- meters have been investigated such as the duration of ageing up to 4000Copyright 2001 IOP Publishing Ltd 7. Figure 3.2. Filled hole compression after 2000 hours of isothermal ageing. hours, the thermal cycling eect, the position of the impact (before or after thermal ageing), and the temperature of ageing, 1208C or 1808C. These results indicate that the position of the impact before or after the thermal ageing is an important parameter. If the maximum temperature is 1208C, Figure 3.3. Glass transition temperature versus time of ageing at 1808C.Copyright 2001 IOP Publishing Ltd 8. Figure 3.4. Filled hole compression at 1808C after isothermal ageing at 1808C. no degradation of properties is observed up to 4000 hours of ageing or 1000 cycles. In these conditions, no oxidation or microcracks are observed in the composite materials. If the temperature of ageing is 1808C, under isothermal conditions we do observe oxidation on the exposed edges of the samples, as shown in gure 3.6(a). If the ageing is done under thermal cycling conditions Figure 3.5. Compression after impact test performed after thermal ageing on bismaleimide composite.Copyright 2001 IOP Publishing Ltd 9. (a) (b) Figure 3.6. Micrographic observation of a bismaleimide composite. (a) 4000 hours at 1808C under isothermal conditions. (b) 1000 cycles with Tmax 1808C and Tmin 508C. with Tmax 1808C and Tmin 508C, we do observe microcracks inside the composite material, as shown in gure 3.6(b).Additionally to the study of the eect of thermal ageing, Aerospatiale developed specic creep test procedures on 4584S specimens to test the creep behaviour of various composite materials. Figure 3.7 presents the creep behaviour of the dierent candidate composites for supersonic aircraft.Aerospatiale is also beginning a research programme to assess the long term behaviour of carbon bre reinforced polymers under complex thermo- mechanical loading. The typical ight spectrum for the future supersonic civil transport in the hypothesis of Mach 2.05 is presented in gure 3.8. This ight spectrum induces strong creepfatiguethermal cycling interactions. In order to test the degrading eects of coldhot thermal cycling and low frequency fatigue cycling compared with classical creep testing, on the residual proper- ties of the composites after exposure, three specic accelerated thermo- mechanical cycling conditions were dened, as shown in gure 3.9. Each cycling type corresponds to 10 000 hours at 1208C. Compared with typical ight conditions, the maximum temperature has been increased by 208C and the maximum stress has been doubled for test acceleration.A specic testing apparatus has been developed to perform the three cycling spectra (gure 3.9). These cycling spectra have been applied to three composite systems with the same bre: one bismaleimide, one cyanateCopyright 2001 IOP Publishing Ltd 10. Figure 3.7. Creep testing on 4584S laminates at 1208C, 70 MPa. and one epoxy. First results are available on the cyanate matrix composite. Table 3.4 presents residual properties after cycling exposure on quasi- isotropic open hole tension (OHT) and lled hole compression (FHC) specimens. These results tend to show that coldhot thermal cycling deter- mines the composite compression properties. The duration of exposure is limited compared with what has to be justied (at least 60 000 hours). This means that long-term tests have to be carried out and special care has to be paid to the development of reliable models, able to predict long-term behaviour from short-term accelerated tests. Figure 3.8. Typical ight spectrum of supersonic aircraft.Copyright 2001 IOP Publishing Ltd 11. Temp (C) 120 Stress Time 0,3 S010000 hrsTimeType 1 TempCold-Hot thermal (C)cycle120x 2500 Temp0,5 h4 hrs0,5 h(C)120Time - 55StressStress Time0,3 S00,3 S0 x 2500 4 hrsTime TimeType 2 Type 3 Figure 3.9. Creep facilities with thermal cycling chamber (558C 2008C). Table 3.4. Eect of cycling on the residual proper-ties of cyanate matrix composites.Loss of stress(% of initial stress)FHC OHT Type 1 (creep)3.9 2 Type 2 (thermal cycle) 10.4 2.9 Type 3 (fatigue)5.7 1.9Copyright 2001 IOP Publishing Ltd 12. Summary The pre-design studies conducted at Aerospatiale, BAe and DASA indicate that, because of the drastic economical and technical requirements dened for future supersonic civil transport aircraft, an important share of the struc- ture of this aircraft will have to be made out of polymeric matrix composites and advanced lightweight aluminium alloys.For the aluminium alloys, studies were oriented in two directions: . the development of improved alloys . the analysis of the behaviour of aluminium components in creepfatigue interaction conditions. Concerning the rst topic, recent research has resulted in the development of new low density alloys, derived from 2650 and 6056 families, with much better creep resistance than the 2618A Concorde alloy, combined with a frac- ture toughness better than the reference 2024 T3 subsonic alloy. For the carbon bre reinforced polymers, the test procedure for evaluating the damaging of these materials in real cycling conditions has been established, and pilot test equipment has been designed and built. Key points that will need to be studied in more detail have been identied, such as the inuence of thermal ageing on the damage tolerance properties of the composite materials and the creepfatiguethermal ageing interactions. References [1] Swadling S J 1993 J. de Physique IV 3 1130 [2] Barbaux Y et al 1994 Proceedings of EAC 94 Toulouse 2527 October, pp 433439 [3] Patri G and Frison G 1994 Revue Aerospatiale 108 [4] Barbaux Y and Polmear I J 1995 Proceedings of PICAST 2-AAC 6 Melbourne 2023 March 2 pp 515520 [5] Barbaux Y et al 1995 Proceedings of EUROMAT 95 Padova, SeptemberCopyright 2001 IOP Publishing Ltd