- Email: [email protected]
Fatigue in aircraft structures
Probabilistic fracture mechanics
295
Corrosion fatigue of pressure vessel steels in water reactor coolants is a complex problem, but one in which considerable progress has been made in understanding the circumstances under which large environmental effects can occur, and how they can be avoided. 5 It has been found necessary to develop a time-based approach to the analysis of corrosion fatigue data rather than the traditional crack growth per cycle method. 6 This is quite a radical departure and is the subject of much current investigation. Undoubtedly, distinctions will have to be made between those circumstances under which environmental enhancement of the rate of crack growth occurs and those under which it does not. Then, with the aid of E D E A C and F A T D A C , these separated data groups can be analysed, and proper statistically based descriptions derived. In conclusion, it is clear that significant progress has been made in developing databases on fatigue crack growth in pressure boundary steels for water reactors, and in their statistical analysis. For the case of corrosion fatigue crack growth, further work remains to be done to determine the mean and variance of the data when analysed according to new time-based models of crack growth. These are important steps towards improved probabilistic analysis of fatigue crack growth in pressurized nuclear components.
References I. Marshall, W. An assessment of the integrity of PWR pressure vessels, UKAEA, 1982. 2. Rungta, R., Mindlin, H. and Gilman, J. D. Applications of EPRI data base on environmentally assisted cracking in the nuclear industry, Mat. Perjbrm., November 1986, pp. 43 52. 3. Eason, E. D., Andrew, S. P. and Warmbrodt, S. B. FATDAC--an interactive fatigue and corrosion fatigue data analysis code, A S M E Winter Meeting, Miami, November 1985. 4. James, L. A. and Jones, D. P. Fatigue crack growth correlations for austenitic stainless steels in air, ASME PVP No. 99 .from 'Predictive Capabilities in Environmentally Assisted Cracking', 1985. 5. Cullen, W. H. Proc. 2nd lAEA Specialists' Meeting on 'Subcritical Crack Growth', Sendai, Japan, 1985; NUREG/CP-O067, 1986. 6. Gilman, J. D. Application model for predicting corrosion fatigue crack growth in reactor pressure vessel steels in LWR environments, ASME Winter Annual Meeting, Miami, November 1985. F A T I G U E IN A I R C R A F T S T R U C T U R E S J. B. Young, Cranfield Institute o f Technology Fatigue loading on aircraft structures is produced mainly by gusts, manoeuvres, landing and ground loads. Fatigue life estimation is based on
296
T. Jutla
measurements of loads due to these effects that have been accumulated for over the past 30 years. Scatter factors are used to reduce the probability of failure to 1 in 1000. To achieve such a success requires a substantial expense on structural maintenance and repair. Proposals are in hand that will change the design process with the aim of reducing these costs. The sources of fatigue loading may be classified both according to their relative frequencies and to their association with either the number of flights or the duration or distance flown. The most significant ones are shown below, some affecting only limited areas of aircraft. Loads related to number of flights
(a)
Once or twice per flight: Steady wing (or rotor) lift with tail balance High lift devices, airbrakes, flaps Cabin pressurization Engine thrust and thrust reversal Undercarriage operations Landing reaction and spin-up Catapult and arrest Thermal effect of supersonic flight
(b) Many times per flight: Manoeuvres Operational Trim Speed control and airbrakes Undercarriage loads due to drag, runway roughness, braking, steering Vibrations and jet noise during run-up Loads related to flight duration or distance
Atmospheric turbulence Stabilizing control Pressure fluctuations due to jet efflux Slipstream, etc. (e.g. acoustic effects) Mechanical vibrations Helicopter rotor and rotor-induced loads The Engineering Sciences Data Unit (ESDU) presents data on gusts. For military aircraft, USAF MIL-A-008866A is a useful source of data for
Probabilistic ,fracture mechanics
297
manoeuvres. Both ESDU and MIL-A-008866A provide data on fatigue loading due to landing and ground loads.
Life evaluation and 'scatter' factors Scatter factors need to be applied to account for: (a) (b) (c)
Frequency of occurrence when this is based on an average. Material variability. Component variability. This is often taken as 3.33 on life, which has a log-normal distribution (based on a C.V. = 10% and three standard derivatives below mean, which is equivalent to Prob = 99"85%). (d) Loading: the most damaging part of the flight is usually the g r o u n d air-ground cycle (lowest to highest peaks in flight), and a factor of 1-5 is used to account for individual aircraft variation. Therefore, the safe fatigue life is obtained by dividing the mean fatigue life by a factor of 5 (i.e. 3"33 x 1"5).
Damage tolerance approach The philosophy in this approach is that consideration is given to crack growth from flaws which may be present in the structure as manufactured. Such flaws may arise from inherent metallurgical or manufacturing imperfections. Fatigue life is determined by the time taken for a crack to grow from initial to critical crack size, allowance being made for scatter in crack growth rates and loading.
295
Corrosion fatigue of pressure vessel steels in water reactor coolants is a complex problem, but one in which considerable progress has been made in understanding the circumstances under which large environmental effects can occur, and how they can be avoided. 5 It has been found necessary to develop a time-based approach to the analysis of corrosion fatigue data rather than the traditional crack growth per cycle method. 6 This is quite a radical departure and is the subject of much current investigation. Undoubtedly, distinctions will have to be made between those circumstances under which environmental enhancement of the rate of crack growth occurs and those under which it does not. Then, with the aid of E D E A C and F A T D A C , these separated data groups can be analysed, and proper statistically based descriptions derived. In conclusion, it is clear that significant progress has been made in developing databases on fatigue crack growth in pressure boundary steels for water reactors, and in their statistical analysis. For the case of corrosion fatigue crack growth, further work remains to be done to determine the mean and variance of the data when analysed according to new time-based models of crack growth. These are important steps towards improved probabilistic analysis of fatigue crack growth in pressurized nuclear components.
References I. Marshall, W. An assessment of the integrity of PWR pressure vessels, UKAEA, 1982. 2. Rungta, R., Mindlin, H. and Gilman, J. D. Applications of EPRI data base on environmentally assisted cracking in the nuclear industry, Mat. Perjbrm., November 1986, pp. 43 52. 3. Eason, E. D., Andrew, S. P. and Warmbrodt, S. B. FATDAC--an interactive fatigue and corrosion fatigue data analysis code, A S M E Winter Meeting, Miami, November 1985. 4. James, L. A. and Jones, D. P. Fatigue crack growth correlations for austenitic stainless steels in air, ASME PVP No. 99 .from 'Predictive Capabilities in Environmentally Assisted Cracking', 1985. 5. Cullen, W. H. Proc. 2nd lAEA Specialists' Meeting on 'Subcritical Crack Growth', Sendai, Japan, 1985; NUREG/CP-O067, 1986. 6. Gilman, J. D. Application model for predicting corrosion fatigue crack growth in reactor pressure vessel steels in LWR environments, ASME Winter Annual Meeting, Miami, November 1985. F A T I G U E IN A I R C R A F T S T R U C T U R E S J. B. Young, Cranfield Institute o f Technology Fatigue loading on aircraft structures is produced mainly by gusts, manoeuvres, landing and ground loads. Fatigue life estimation is based on
296
T. Jutla
measurements of loads due to these effects that have been accumulated for over the past 30 years. Scatter factors are used to reduce the probability of failure to 1 in 1000. To achieve such a success requires a substantial expense on structural maintenance and repair. Proposals are in hand that will change the design process with the aim of reducing these costs. The sources of fatigue loading may be classified both according to their relative frequencies and to their association with either the number of flights or the duration or distance flown. The most significant ones are shown below, some affecting only limited areas of aircraft. Loads related to number of flights
(a)
Once or twice per flight: Steady wing (or rotor) lift with tail balance High lift devices, airbrakes, flaps Cabin pressurization Engine thrust and thrust reversal Undercarriage operations Landing reaction and spin-up Catapult and arrest Thermal effect of supersonic flight
(b) Many times per flight: Manoeuvres Operational Trim Speed control and airbrakes Undercarriage loads due to drag, runway roughness, braking, steering Vibrations and jet noise during run-up Loads related to flight duration or distance
Atmospheric turbulence Stabilizing control Pressure fluctuations due to jet efflux Slipstream, etc. (e.g. acoustic effects) Mechanical vibrations Helicopter rotor and rotor-induced loads The Engineering Sciences Data Unit (ESDU) presents data on gusts. For military aircraft, USAF MIL-A-008866A is a useful source of data for
Probabilistic ,fracture mechanics
297
manoeuvres. Both ESDU and MIL-A-008866A provide data on fatigue loading due to landing and ground loads.
Life evaluation and 'scatter' factors Scatter factors need to be applied to account for: (a) (b) (c)
Frequency of occurrence when this is based on an average. Material variability. Component variability. This is often taken as 3.33 on life, which has a log-normal distribution (based on a C.V. = 10% and three standard derivatives below mean, which is equivalent to Prob = 99"85%). (d) Loading: the most damaging part of the flight is usually the g r o u n d air-ground cycle (lowest to highest peaks in flight), and a factor of 1-5 is used to account for individual aircraft variation. Therefore, the safe fatigue life is obtained by dividing the mean fatigue life by a factor of 5 (i.e. 3"33 x 1"5).
Damage tolerance approach The philosophy in this approach is that consideration is given to crack growth from flaws which may be present in the structure as manufactured. Such flaws may arise from inherent metallurgical or manufacturing imperfections. Fatigue life is determined by the time taken for a crack to grow from initial to critical crack size, allowance being made for scatter in crack growth rates and loading.