Back to Basics

The Role of Nondestructive Testing in
Aircraft Damage Tolerance

by Dale L.Ball^*

 

Guest Technical Editor: This "Back to Basics" article provides valuable insight into the role of NDT in damage tolerance methods for aircraft life management. To be accepted, new NDT solutions, including the embedded sensors for structural health monitoring addressed in this special issue, must offer either cost reduction or improved safety.

The damage tolerance philosophy developed jointly by the US Air Force and a number of US aircraft manufacturers in the late 1960s and early 1970s has been tremendously successful in mitigating structural failures due to fatigue in both military and commercial aircraft designed and built since that time. As described by the US Department of Defense (1996a), this philosophy is applied over the entire life cycle of the airframe, beginning with the design phase and continuing throughout the operational deployment phase. In each life cycle phase, nondestructive testing (NDT) plays a critical role in implementing damage tolerance requirements and imposing damage tolerance control. As we shall see below, any new developments in NDT capability, reliability and cost can have a significant effect on the cost and effectiveness of maintaining aircraft structural integrity.

 

The Damage Tolerance Approach to Structural Integrity
US Department of Defense damage tolerance guidelines (1996b) state that cracklike discontinuities must be assumed to exist in all critical elements of a structure at the time that that structure enters service. For a given structural element, the discontinuity is assumed to be in the most critical (highly stressed) area and in the most critical orientation. The guidelines go on to state that these discontinuities will not grow to a size sufficient to cause failure within a specified interval of time in a specified operational environment. Measures can be taken in at least three broad areas in order to meet the very stringent requirements of the guidelines: care must be taken during detail design to keep operating stresses low and to minimize or avoid sharp stress raisers; materials must be selected based on their ability to resist cracking; and NDT procedures must be developed and consistently applied, both during initial certification and for subsequent periodic tests. This third step is a crucial part of structural integrity assurance; it establishes a critical parameter in damage tolerance analysis - the assumed initial crack size. From a damage tolerance certification point of view, the smaller the assumed initial crack size, the longer the fatigue crack growth life will be. However, this value can be no smaller than the minimum size that can be reliably detected by NDT.

---

Improvements in NDT capability can have significant, beneficial effects on the cost of fracture control.

---

Discontinuity Types and Distributions
As shown in Figure 1, in any given structural component, up to three types of discontinuities, each with its own size or number distribution, will exist (Gallagher et al., 1984). The first type consists of intrinsic material discontinuities. These discontinuities are the result of the material production process (alloying, heat treating, forming and so on) and include porosity, microcracks, inclusions and surface pits. One of the ways that damage tolerance principles can be applied during the design phase is through the development and selection of materials for which this discontinuity distribution is minimized. These discontinuities are generally below the typical NDT detection capability and the structural element should be sized such that the upper bound of the distribution a_o will not grow appreciably during design service usage. This maximum intrinsic discontinuity size effectively defines the initial discontinuity size for continuing damage analysis as described below. It may also serve as the basis for the initial sizing of multisite damage cracks, required for widespread fatigue damage analysis.

Figure 1 - Discontinuity distributions and limiting sizes used in fracture control (Walker et al., 1979). 

 

The second group of discontinuities consists of those that are introduced during fabrication (machining, assembly, finishing and so on). These include machining marks, scratches or any type of damage that could produce a cracklike discontinuity. Some portion of these discontinuities will be large enough to be detected by NDT. They will also be large enough to grow appreciably in the design operating environment. So again, damage tolerance requirements are imposed, this time during the manufacturing phase, by performing detailed nondestructive testing of the critical regions of all fracture critical parts. A great deal of attention is given to establishing the maximum size of a discontinuity that could escape detection during fabrication and thus exist in a structural component at the time that the vehicle enters service. This discontinuity size is designated a_i and serves as the initial crack size for the fatigue crack growth analyses which are used to estimate the service life or required test intervals for the component.

Discontinuities that are formed in service make up the third group. These include cracks formed by fatigue, corrosion and impact damage. These discontinuities are also large enough to grow in service and must be guarded against. Damage tolerance requirements are imposed during the operational phase by requiring that cracks (both fabrication and service induced) do not grow to a critical size within a specified period of time (either the design service life or the required test interval). The largest discontinuity that could remain undetected after an inservice test is designated a_s. This value is generally larger than a_i because the test is done on an assembled structure; to one extent or another, fasteners, surrounding structure, coatings and so on will degrade the efficacy of any given NDT technique.

 

Summary of Requirements
Obviously, single values cannot be defined for a_i and a_s; they are dependent on a large number of factors, some of which are not well controlled. These factors range from structure type and accessibility of critical location to test method (ultrasonic, eddy current, dye penetrant and so on), to the environment in which the test must take place to the experience level of the inspector. Nonetheless, a value must be chosen for use in damage tolerance certification analyses. This problem has been resolved by defining a_i and a_s on the basis of probability of detection and confidence limits. The probability of detection is a function of crack length; its value (for a given crack length) is defined as the proportion of correct measurements by a given NDT system with a representative operator for a number of structural elements in a defined environment. This determination requires that a statistical analysis of a large number of controlled tests be performed.

Definition of the probability of detection does not take into account some of the variability that will be experienced on the factory floor or in the field. As a result, another statistically based measure, which accounts for these uncertainties (uncertainties that are not intrinsic to the test technique itself, such as the test environment or the ability of the inspector), is defined. This measure is referred to as the confidence limit. While it is clearly not possible to have 100% probability of detection with 100% confidence, it is desirable to set these requirements as high as is economically feasible in order to improve the odds of safe operation during the required service life or test interval. The values that were eventually agreed upon during the development of the damage tolerance requirements were 90% probability of detection with a 95% confidence limit.

Specifying an a_i or a_s value based on a required probability of detection/confidence limit is still too cumbersome, however, for design or fleet management use. Instead, the aircraft structural specification (US Department of Defense, 1996b) segregates all fracture controlled structures into three categories and then provides specific initial crack sizes for typical geometric details in each category. In all cases, the a_i or a_s assigned to any given structural category or detail type will satisfy the 90/95 probability of detection/confidence limit requirement. The first of the three structural categories is designated as "slow crack growth." This includes structures that are unable to be tested or that can only be tested at the depot with component removal. The second category is "fail safe, multiple load path," which refers to structures having multiple elements carrying a primary load but which are joined directly to each other, thus making it possible for cracks to grow from one element into the next. The last category is referred to as "fail safe crack arrest." This too is a multiple load path structure, but includes intermediate structural elements that prevent cracks from growing from one primary load carrying member into an adjacent one. A summary of the required initial crack sizes for primary cracks is given in Table 1. Similar guidelines are given in the JSGS-87221 Handbook (1996b) for secondary cracks and for inservice tests.

Table 1 Initial size assumptions for primary discontinuities
Category	Critical Detail	Intitial Discontinuity Size Assumption*
Slow crack growth and fail safe primary element	holes, cutouts and so on	for thickness ≤ 1.3 mm (0.05 in.); through thickness  with ci = 1.3 mm (0.05 in.)
		for thickness > 1.3 mm (0.05 in.); quarter circular corner with ci = 1.3 mm (0.05 in.)
	other	for thickness ≤ 3.2 mm (0.125 in.); through thickness  with 2ci = 6.4 mm (0.25 in.)
		for thickness > 3.2 mm (0.125 in.); semicircular thickness  with ci = 3.2 mm (0.125 in.)
	components with embedded discontinuities	to be determined based on NDT capability
	welded componentst	to be determined based on NDT capability
Fail safe multimode path (adjacent structure)	holes, cutouts and so on	for thickness ≤ 1.3 mm (0.05 in.); through thickness  with ci = 0.05 + Δc±
		for thickness > 1.3 mm (0.05 in.); quarter circular corner with ci = 0.05 + Δc±
	other	for thickness ≤ 3.2 mm (0.125 in.); through thickness  with 2ci = 0.25 + 2Δc±
		for thickness > 3.2 mm (0.125 in.); semicircular surface  with ci = 0.125 + Δc±
Fail safe crack arrest (adjacent structure)	skin stringer or skin frame	ci extends across two panels (bays); central stringer (frame) broken
	skin stringer or skin frame with tear straps	ci extends between tear straps (bays); central stringer (frame) broken
	holes, cutouts and so on other	quarter circular corner with ci=0.005 + Δc± semicircular surface  with ci = 0.01 + Δc±
* discontinuity oriented in most critical direction t discontinuity assumed in both weld material and in heat affected zone of parent material ± Δc is the growth of the indicated discontinuity between time that part enters service and time of primary element failure

 

Secondary cracks are cracks that grow under the same load and in the same region as a primary crack; they must be considered in the event that termination of the growth of the primary crack does not coincide with total element failure. The assumed initial size for secondary cracks is based on the maximum intrinsic material discontinuity size a_o discussed above. Historically, the analysis of secondary cracking has been performed in the context of "continuing damage." That is, the analysis is performed subsequent to, and is decoupled from, the primary crack growth analysis. In recent years, however, improved understanding of a group of phenomena collectively known as widespread fatigue damage has shown that the presence of secondary cracking can seriously degrade residual strength capability (Advisory Group for Aerospace Research and Development, 1995). This premature residual strength loss can only be captured analytically if the secondary cracks are considered in the primary (lead) crack analysis.

 

Effect of Improved NDT Capability
For military aircraft production and operations, there are very substantial costs associated with damage tolerance (fracture) control. During the design phase, significantly more analysis is done on fracture critical parts than on noncritical ones. During manufacturing quality control, traceability requirements are significantly more stringent and during operational usage tests are often more frequent. Given the role that NDT plays in establishing the initial discontinuity sizes used for damage tolerance analyses, there is a very real link between test capability and cost of acquisition and operation of any given aircraft which must meet damage tolerance requirements.

For a given critical location in a given structural element for a specified loading environment, a fatigue crack growth analysis will be used to determine both the initial test interval and the subsequent inservice test intervals. As shown in Figure 2, if the total calculated fatigue crack growth life from a_i is N_a, then the first required test must take place at or before t = N_a/2. The factor of 1/2 arises because there is a standard, mandated safety factor of two on life. Note that if N_a/2 is greater than the design service life, then no testing will be required. All subsequent test intervals are based on the crack growth analysis from the larger starting crack size a_s. As a result, the subsequent test intervals N_b/2 are shorter.

The first effect assessment that we will consider is in the area of in service testing. These tests are the most difficult because they are conducted in the depot, at best, and at worst, in the field. As mentioned above, access to many critical areas can be limited or even nonexistent. Using the type of analysis shown in Figure 2, we may project what effect advancements in inservice NDT capability might have on the test schedule for a fracture controlled part.

Figure 2 - Definition of initial and inservice test intervals based on as fabricated and inservice initial crack sizes.

 

We begin with a baseline analysis in which a_s > a_i, for example, a_i = 1.3 mm (0.05 in.) and a_s = 2.5 mm (0.1 in.); the resulting test schedule is shown in Figure 3. In this case, the component must be tested four times in order to achieve the required design life. If, however, advancements in NDT capability were to be made which would allow a_s to approach a_i, at least one test could be eliminated (see Figure 4). Assuming that cost per test remains flat over time, this would translate into a 25% cost savings.

Figure 3 - Test schedule for damage tolerance critical part with a_s>a_i. 

Figure 4 - Test schedule for damage tolerance critical part with a_s=a_i. 

Figure 5 - Test schedule for damage tolerance critical part with ai<<as.

 

 

Another area of NDT capability that can have a significant effect on damage tolerance certification is that done as a part of quality assurance in the factory. For fracture critical parts, this capability defines the maximum crack size that will go undetected in as fabricated structure a_i. If we again use the analysis shown in Figure 3 as a baseline, we can assess the effect that improvements in in-factory NDT capability will have. The results shown in Figure 5 show the calculated test schedule for the case where a_i has been reduced to 0.2 mm (0.01 in.). In this example, a rather dramatic reduction in test requirements and in associated fleet maintenance cost is projected. This scenario is in fact accommodated in the JSGS-87221 aircraft structures specification (US Department of Defense, 1996b) - if the aircraft manufacturer can demonstrate the ability to detect discontinuities which are smaller than those specified in Table 1 with 95/90 probability of detection/confident limit, then it is permitted to use those lower values for damage tolerance certification analyses. This accommodation is typically only made at a specific critical location or small sets of related locations.

 

Conclusion
Nondestructive testing is a cornerstone of the aircraft damage tolerance certification and control processes. While it is clear that the success of these processes in preventing aircraft structural failure is a testament, in part, to just how capable today's NDT technologies are, it is also clear that there are vast opportunities for improvement. A very useful discussion of critical needs for management of aging US Air Force aircraft, along with candidate NDT technologies and recommendations for research, was given in a report on aging US Air Force aircraft by the National Research Council (1997). Improvements in NDT capability can have significant, beneficial effects on the cost of fracture control. What is even more important than the economic benefit, though, is the reduced risk of structural failure, which for most aircraft translates directly into reduced risk of loss of life.

 

References
Advisory Group for Aerospace Research and Development, Report AGARD-CP-568, Widespread Fatigue Damage in Military Aircraft, Proceedings of the 80th Meeting of AGARD Structures and Materials Panel, Neuilly-Sur-Seine, France, 1995.

Gallagher, J.P., F.J. Giessler, A.P. Berens and R.M. Engle, Jr., AFWAL-TR-82-3073, USAF Damage Tolerant Design Handbook, Wright-Patterson AFB, Ohio, Air Force Wright Aeronautical Labs, May 1984.

National Research Council, NMAB-488-2, Aging of U.S. Air Force Aircraft, Washington, DC, National Materials Advisory Board, National Research Council, Committee on Aging of US Air Force Aircraft, 1997.

US Department of Defense, MIL-HDBK-1530, General Guidelines for Aircraft Structural Integrity Program, Washington, DC, US Department of Defense, 1996a.

US Department of Defense, JSGS-87221, Aircraft Structures, General Specification for and Handbook, Washington, DC, US Department of Defense, 1996b.

Walker, E.K., J.C. Ekvall and J.E. Rhodes, "Design for Continuing Structural Integrity," Journal of Engineering Materials and Technology, Transactions of ASME, Vol. 102, 1980, pp. 32-39.

 

* Lockheed Martin Aeronautics Company, 1 Lockheed Blvd., PO Box 748, MZ 8862, Ft. Worth, TX 76101; (817) 935-5902; e-mail <dale.l.ball@lmco.com>.

 

Copyright © 2003 by the American Society for Nondestructive Testing, Inc. All rights reserved.