The strength of the aircraft
The principle of safe damages. Safety of aircraft is directly related to durability.
The design is called a safe operation, if required minimum inspection and repairs at a satisfactory core functions. Satisfactory performance is negligible probability of structural failure for a civil aircraft or acceptably low probability of failure for military aircraft. Safety of passengers and crew of civil aircraft is of paramount importance. Methods of analysis of structures, reliable in operation, designed primarily for civilian aircraft.
Modern plane is semi-monocoque type structure consisting of thin-walled sheets, supported by beams (farms) and the stringers to prevent buckling. The outer skin or wall forms an aerodynamic contour of the unit - the fuselage, wing, stabilizer. The stiffeners are attached to the inner surface of the skin and perceive concentrated loads. This design for many years served as the main object of aerodynamic research and distinguishes devices from conventional building designs.
The required service life of civil aircraft is determined on the basis of comprehensive economic considerations. He is 10-15 years. Designer primarily trying to ensure a long-term operation of the aircraft without cracking. To do this, he uses the method of calculation by which minimizes stress concentrations and tries to keep the voltage as low as possible, based on the requirements of the flight characteristics. For parts that are difficult to repair or replace, the designer can try to ensuredesired durability without cracking, equal to the lifetime of the aircraft. For many designs it is impossible. In addition, there is a risk of damage to structures serving transport, stumbling on the runway and decaying parts of the engine or the propeller. The designer must minimize the loss of strength resulting in occurrence of cracks or fatigue damage during operation of the aircraft. He solves this problem as follows:
picks up the material and determine the dimensions of parts to ensure adequate structural strength in cracks;
uses elements of safety (track variable loads and traffic, hindering the development of cracks);
selects a material having a low rate of fatigue cracks.
One of the modern means of improving the reliability of designs while increasing resource reduction of materials and improve economic efficiency - the design and definition of the duration of operation on the principle of safe damages. This takes into account the presence of structural elements of the initial metallurgical and technological defects and the formation of cracks in them the accumulation of operational damage.
Development and implementation of the principle of safe damages are possible only with the use of methods of fracture mechanics. Determination of stress-strain state of structural elements containing defects such as cracks, is the most responsible and difficult stage of strength calculation. In accordance with generally accepted notions of stress-strain state of the body with the crack is completely characterized by the values of the stress intensity factor. In its preliminary determination based almost all currently known criteria of brittle and quasi-brittle fracture, as well as dependencies, describing the growth of fatigue cracks.
The term "safe damage" refers to a structure designed so as to minimize the possibility of damage output plane because of the spread of undetected defects, cracks or other such defects. The production structures, which should be no damage, it is necessary to solve two major problems. These problems are to provide a controlled-growth defects, ie. e. safe operation with cracks, and in the forced containment of damage, so that should be provided with a residual life or the residual strength. In addition, the calculation of allowable damages does not preclude the need for careful analysis and calculation of fatigue.
The main point which is based on the concept of safe damage consists in the fact that there are always defects, even in new designs, and that they may remain undetected. Thus, the first condition for the admissibility of the defect is a condition that any element of the design, including all additional units to transmit the load, must allow safe operation in the presence of cracks.
Control Growth defects. The occurrence of fatigue cracks can be avoided by creating a structure at all points where the voltage is below a certain level would. However, reducing the stress leads to an increase in the design weight. Furthermore, cracks may occur not only from fatigue, but also for other reasons, for example due to accidental damage produced during operation or due to material defects. Therefore, in the real construction permit the presence of some small cracks in the structure at the time of shipment. The larger of these cracks may develop during operation.
The most important element of the safety principles of damage becomes a time period during which the crack can be detected. Due to various contingencies probability of detecting cracks when viewed unstable. Sometimes scarcely visible cracks found in the most remote zones of the construction and at the same time can be passed very largecracks elsewhere. For the case when was missed during the inspection "Boeing-747» crack length 1800 mm below the fairing in the pressurized cabin of an aircraft.
Therefore, structural elements for determining the bearing capacity of the airframe must be made fracture control program. An important element of the program is to develop a fracture control verification methods. For each element to be developed and proposed appropriate test methods. For certain parts of the elements may require the use of non-destructive testing of different sensitivity. Timing verification set on the basis of available information about the growth of a crack with specified initial size of the defect and the size of the detected defect, which depends on the sensitivity of flaw detection method employed. Terms of checks should be established on the basis of that, provided that the required safety factor undetected defect has not reached a critical size before the next test. Usually, intervals between regular inspections are assigned so that before reaching any critical crack size passed two checks.
Safety principles of damage to aircraft design necessitated greater use of non-destructive methods for monitoring the technical condition of functional systems. The possibilities of different methods of non-destructive testing for the detection of fatigue cracks. NDT methods are constantly being improved.
Fatigue, corrosion and crack-resistant. In practical operation of the sun are many cases of destruction of parts components and assemblies of material fatigue. Such degradation is the result of repetitive or variable loads. And for fatigue failures it requires significantly less than the maximum load than at static breaking. In flight and on the ground when driving many items and structural elements of sun exposed and variable loads, although the nominal voltage often low, the stress concentration, which usually does not reduce the static strength, can lead to fatiguedestruction. This is confirmed by the practice of exploitation of not only the sun but also ground vehicles. Indeed, you can almost always observe the fatigue failure and very rare - the destruction of the static loads.
Feature fatigue failure - the absence of deformation in the zone of destruction. Similar phenomena are observed even in materials such as mild steel, which is highly plasticity under static destruction. This is a dangerous feature of fatigue failure, since there are no signs preceding the destruction. Emerging signs of fatigue are usually very small and difficult to detect until they reach macroscopic size. Then they spread rapidly and in a short time there is complete destruction. Thus, timely detection of fatigue cracks - a difficult task. Most frequently the fatigue cracks are generated in the zone shape changes or defects in surfaces of the parts.
Such defects, as well as a small change in the working section of the details do not affect the static strength, as plastic deformation reduces the effect of stress concentration. At the same time, fatigue failure of parts plastic deformations tend to be small, thereby reducing the concentration of stress in the area and there is no account of the concentrationstress is essential, however, it is important in the design of components operating under varying loads, making them easier and safer against fatigue failure.
Thus, the factors influencing the fatigue resistance include: stress concentrators, dimensions of parts, the relative importance of both static and cyclic loads and corrosion, especially corrosion of friction, which is the result of repeated small movements of the two contacting surfaces.
Fatigue is usually caused by the destruction of many thousands or millions of load cycles. However, they can occur after tens or even hundreds of cycles.
All the elements, parts and units of the Armed Forces are exposed to dynamic loads when driving on the ground and in flight. Variable load various kinds, acting on structural components, machine parts and devices, are responsible for corresponding variable voltage, which eventually lead to fatigue fractures. The rate of processes of mechanical destruction of stressed parts and units, respectively, and the time to failure depends on the structure and material properties of the stress caused by the action of load, temperature, and other factors. However, the nature of the fatigue fracture is a kind of a form different from the brittle fracture.
Fatigue failure of parts usually begins near the metallurgical or technological defects, stress concentration zones, as well as the presence of technological defects in the products.
As is known, static failure is mainly determined by the probability of occurrence of a large load in flight, for example, by a gust of air as a result of which the Sun will operate the load exceeds the limits of the static strength of the structure, ie, the possibility of static destruction - it is essentially a question of probability of occurrence of a large load.
Fatigue failure under these assumptions - the result of the application of a sufficient number of load cycles, or a sufficient number of flights Sun at a certain distance.
The main difference between fatigue and static loading is as follows:
a major factor in fatigue strength for a given distribution of loads, even with the scatter of the data is the number of load changes or service life; for the static strength and the destruction - of the load;
nature of a probabilistic approach to fatigue loading is significantly different from that of a probabilistic approach to the static loading - for the specific operating conditions influence the likelihood of a single large load on the aircraft, for example, the gust exceeds the static destructive and does not depend on the time of use. This may occur at the beginning and end of life. The probability of fatigue failure is changed during operation, increase significantly by the end of its service life. At the same time designers and scientists believe that the assigned resource or limit the service life and the corresponding level of probability must be such that the frequency of recurrence of the destruction was a sufficiently small value that, if possible, it would be accepted. That value is the probability 10 9, and that taken as a basis for leading foreign and domestic aviation companies.
Aviation experts believe that the corrosion fatigue as well as damage to the same extent determines the service life of the aircraft structure. Most sources of corrosion - structural damage when loading the sun on the ground and scratched skin.
It is known that corrosion damage to the structure is entirely dependent on the operating conditions and the quality of the Armed Forces service.
The instructions, first of all, attention is drawn to the corrosion of the main structural elements of power. It is found that the corrosion is caused by a more internal than external factors. Thus, the cause of corrosion - liquid spilled in the area of the buffet (especially fruit juice) and toilets.
Areas of the fuselage structure, are particularly susceptible to corrosion and fatigue cracks (shaded).
The least dangerous in relation to the total fatigue (uniform) corrosion. But in actual use uniform corrosion in its pure form is rare and is usually supplemented with ulcerative lesions. The effect of such corrosion fatigue resistance.
It can be seen that depending on the area and depth of corrosion damage, fatigue life of an alloy D16T significantly reduced. The area of corrosion damage reduces the fatigue resistance of less than the diameter and depth of corrosion pits.
When using the process of accumulation of fatigue and corrosion damage alternate with partial overlapping each other. It is usually assumed that the corrosive lesions develop on parking, and fatigue - in flight. Corrosion damage is stress concentrators.
Terms and approaches used in the justification of resources within 103 l. h for 20-25 years of operation, determine the need to use while ensuring safety at the present stage, along with the principle of "safe-life" as a progressive principle of "Safe damage."
This last principle allows fatigue damage to the structural elements during the time interval between two consecutive inspections under the conditions that the interval is not too great, damage does not reach its limiting state, and does not lead to destruction of the structure as a whole.
Consequently, the strength criterion of the aircraft, claiming the inadmissibility of cracking, incorrect for the structure as a whole, as in a long-term operation of aircraft virtually impossible to avoid fatigue cracks in some of its elements. It is necessary to find a crack in time and prevent their further development for the maximum allowable size.
Thus, the strength resource of the aircraft should be based on the criterion of strength, taking into account the intensity of the origin and development of cracks for design in general, and in the elements that do not lead to a catastrophic outcome.
There is the concept on which it is believed that during the 30 min. 101 l. h should be safe, and then to 60 * 103 l. h - operation provided by the structural properties of survivability.
Recall that under the sun vitality or functional systems refers to the property that provides the proper performance of the functions specified in the flight (or flights) with individual faults or damage their elements or nodes. It is ensured by the provision, the specific design solutions, favoring rather slow development of damage and sufficient strength in the presence of a fault to be readily available for the detection of damage and objective control, if possible.
Experience shows that during prolonged wear of operation, fatigue and corrosion damage are the most massive failures.
Fatigue cracks lead to a decrease in strength of the structure and determine its strength reliability. Therefore, the design must be provided that the following conditions: the development and distribution of cracks in structural elements should be so slow that the residual static strength in the development of cracks to the size of its visual detection was sufficient for trouble-free operation of the sun without restrictions.
Consider some of the results of tests of samples of skin of the fuselage aircraft with pressurized cockpit. For a schematic diagram of a fatigue crack in the panels of the fuselage of the aircraft DC-10. The residual strength of the aircraft fuselage DC-10 investigated on panels size 4267 2642 x mm radius of curvature Zoe mm. Tests performed under combined loading simulating inertia load and the boost pressure in the passenger cabin. For this panel taken from the top of plating with having an initial crack equal 12 mm. As can be seen, the first stage of the test at a nominal pressure Pa to 0,65 15 000 cycles of crack growth is almost not observed. After the cut in the power elements and some increased internal pressure crack growth rate began to increase, not reaching, however dangerous values. When 46 000 cycles was the destruction of the central bulkhead, then the destruction of the two frames, which resulted in a sharp increase in the rate of development of cracks and destruction of other security elements. Complete destruction of the panel took place at the crack length 1157 mm and at a pressure greater than in 1,53 times the nominal pressure in the cabin.
Similar tests conducted on other panels with a set of security elements, have shown the ability to create designs of increased vitality and of the principle of "safe" of damage to the structure ensuring monitoring its condition at the MOT.
However, the most dangerous fatigue failure of structural elements of the fuselage. For example, cracks in the skin of the fuselage of the aircraft "Comet", appeared near the cutouts for windows, caused the two accidents of this type of aircraft.
The main reason for re-cracking load of the fuselage skin with pressurized cockpit aircraft "Comet" and design flaws. As is known, the aircraft skin undergoes repeated tension-compression load. They led to the development of cracks in the stress concentration. After performing plating crack completions of this type were not observed.
The design allows for increased survivability of certain dimensions of the damage that must meet the more general regulatory requirements. For example, the company "Douglas" believes that the residual strength of the structure of the passenger aircraft must be provided at the fracture wing length 400 mm disrupted middle stringer and in the fuselage for longitudinal crack length 1000 mm disrupted middle titanium stopper or transverse crack up to 400 mm destroyed the middle spar.
The company "Lockheed" determines the following possible damage to the fuselage: a crack in the skin may be long 300 mm destroyed in the middle of the frames or stringer; longitudinal crack in the skin - up to 500 mm; crack, running from the corner of a cut-out to 300 mm with the destruction of the frame or stringer.
The ICAO requirements specified that a minimum level of residual strength of the damaged structures must match the maximum operating load of 66,6% estimated for the calculation of the most important cases of loading.
GOST 27.002 83 defines durability as a property of the object continue to operate until a certain status in the installed system AMO. The limit condition can be caused by: fatal violation of safety requirements due to violation of the structural strength; unavoidable care units for the parameters of tolerance; unavoidable reduction in effectiveness; the need to perform major repairs in accordance with current regulatory and technical documentation.
As reliability and durability is laid in designing aircraft manufacturing ensured and maintained during operation. For AT longevity is determined by the conditions of safety and expediency of its further application on the basis of comparative effectiveness and the possibility of replacing a more perfect models. When designing items AT take into account possible load during the operation modes; choosing appropriate material for parts processing methods. For the elements operating in a friction material is chosen, the most wear under expected operating conditions, and so on. etc.
All this allows designers to not only create a workable design, but also to carry out the relevant calculations can ensure the required standards of durability designed equipment.
Durability as a property of the structure depends on numerous factors, which can be divided into strength, operational and organizational.
Strength include design, manufacturing, processing, load and temperature factors. Among them are stress concentrators in the elements of construction and residual stresses resulting from the imperfect technology and due to plastic deformation in the assembly of parts and repairs; properties of materials and their change during the operation, including an initial static strength; fatigue limit; the stress intensity factor for the type of separation and destruction of the shift.
Experts believe that using modern achievements of science, engineering and technology, we can ensure the longevity of the structure of the main parts of the aircraft to 40 • 103 l. h. Without cracking aircraft can bump 30 103 x x l. h. If we assume that the cost-effective life (or duration of operation) is 60 • 103 l. h, it is possible to provide a guaranteed about half of this period, the sun and the other half will be operated with damage tolerance parts and assemblies and their replacement during repairs.
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