ASM Metals HandBook Vol. 17 - Nondestructive Evaluation and Quality Control

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Fig. 4 Specimen crack depth distribution for two case studies. (a) Case study No. 1. (b) Case study No. 2

This case study, which demonstrates the criticality of planning the reliability demonstration program, showed only the

importance of specimen flaw size to the final outcome of the reliability study.

In the following articles in this Section, fracture control philosophy is discussed, reliability demonstrations that have been

completed are reviewed, and the analysis of NDE data is examined. The final article in this Section, "Models for

Predicting NDE Reliability," provides information on advanced modeling studies for predicting reliability for the

inspection of a particular component. The objective of this Section is to provide the necessary information and resources

to conduct an effective NDE reliability program.

Reference cited in this section

A.P. Berens and P.W. Hovey, "Flaw Detection Reliability Criteria Volume I--

Methods and Results, Final

Report," Report AFWAL-TR-84-4022 Volume I, United States Air Force Material Laboratory, Wright-

Patterson Air Force Base, 1984

Introduction to Quantitative Nondestructive Evaluation

Vicki E. Panhuise, Allied-Signal Aerospace Company, Garrett Engine Division

References

1. "Ultrasonic Inspection of Product over 0.5 in. (13 mm) Thick," Aerospace Material Sp

ecification

2630A, Society of Automotive Engineers, 1980 (original release 1960)

"Standard Practice for Fabricating and Checking Aluminum Alloy Ultrasonic Standard Reference

Blocks," E 127, Annual Book of ASTM Standards, American Society for Testing and Materials

"Standard Recommended Practice for Fabrication and Control of Steel Reference Blocks Used in

Ultrasonic Inspection," E 428, Annual Book of ASTM Standards,

American Society for Testing and

Materials

4. "Recommended Practice, Personnel Qualification and Certification in NDT," SNT-TC-

IA, American

Society for Nondestructive Testing

W. Rummel, Recommended Practice for a Demonstration of Nondestructive Evaluation (NDE)

Reliability on Aircraft Production Parts, Mater. Eval.,Vol 40 (No. 9), Aug 1982, p 922

6. W.H. Lewis, et al., "Reliability of Nondestructive Inspections--Final Report," Report SA-

ALC/MME

76-6-38-1, United States Air Force, Air Logistics Center, Kelly Air Force Base, 1978

7. A.P. Berens and P.W. Hovey, "Flaw Detection Reliability Criteria Volume I--

Methods and Results,

Final Report," Report AFWAL-TR-84-

4022 Volume I, United States Air Force Material Laboratory,

Wright-Patterson Air Force Base, 1984

Fracture Control Philosophy

William D. Cowie, United States Air Force, Aeronautical System Division, Propulsion Directorate

Introduction

FRACTURE CONTROL PHILOSOPHIES are being used in the design, development, and life management of United

States Air Force (USAF) turbine engine and airframe components. This article describes the fracture control program for

turbine engine components. The section "Applications (Case Studies)" of the article "Applications of NDE Reliability to

Systems" in this Volume provides an overview of fracture control programs for both airframe and turbine engine

components.

The establishment of a fracture control philosophy and the implementation of a fracture control program have been

integral components of the USAF turbine engine development process since 1978. They have been applied to new engine

programs as part of the USAF Engine Structural Integrity Program (ENSIP) described in military standard MIL-STD-

1783 (issued formally in 1984). This military standard was reviewed and approved by the Aerospace Industries

Association of America in 1982.

Fracture control philosophy has also been applied to existing inventory USAF engines through structural durability and

damage tolerance assessments. In all, fracture control programs have been applied or are being applied to the F-100, TF-

34, F100-PW-220, F100-PW-229, F110-GE-100, F110-GE-129, F101-GE-102, F109-GA-100, F-119, F120, and T406

engines and have resulted in the implementation of enhanced nondestructive evaluation (NDE) methods (for example,

eddy current inspection) at manufacturing and at field/depot. These inspections have been successful in detecting early

cracking and in accelerating corrective actions. Several developmental efforts in the last 5 years have identified

fluorescent penetrant inspection process improvements that must be implemented within industry and Air Force depots to

improve flaw detection reliability. The need to quantify detection reliability for imbedded defects is also identified.

The engine development process has been evolutionary in terms of the application of upgraded requirements. The new

process of fracture control, sometimes referred to as damage tolerance, is contained in ENSIP, and it involves material

selection as well as design and life management. Recent experience clearly demonstrates that the damage tolerance

requirement is cost effective when assessed on a life cycle basis.

Fracture Control Philosophy

William D. Cowie, United States Air Force, Aeronautical System Division, Propulsion Directorate

Overview of ENSIP

In the past 16 years, a large number of structural problems have occurred in USAF gas turbine engines. Many of these

were safety problems that resulted in loss of aircraft, and an even greater number affected durability, causing a high level

of maintenance and modification costs. All of these problems have adversely affected fleet readiness. The Engine

Structural Integrity Program was intended to reduce these problems substantially and was developed based on the

following specific lessons:

• It is unrealistic (and can be dangerous) to assume defect-free structure in safety-of-flight components

• Critical parts (and part details) and potential failure modes must

be identified early and appropriate

control measures implemented

• Internal thermal and vibratory environments must be identified early in the engine development

• Predicted analytical stresses must be verified by test for complex components

• Materials and processes must be adequately characterized (particularly, the fracture properties)

• Design stress spectra, component test spectra, and full-

scale engine test spectra must be based on the

anticipated service usage of the engine, that is, accelerated mission-related testing

• Potential engine/airframe structural interactions must be defined and accounted for

•

Management procedures (such as individual engine tracking procedures and realistic inspection and

maintenance requirements) must be defined and enforced

The Engine Structural Integrity Program was established by the Air Force to provide an organized and disciplined

approach to the structural design, analysis, development, production, and life management of gas turbine engines, with the

goal of ensuring engine structural safety, increased service readiness, and reduced life cycle costs. The five major tasks

associated with ENSIP are the development of design information; design analysis and component and material

characterization; component and core engine testing; ground and flight engine testing; and production quality control and

engine life management. Each major task is subdivided into a number of subtasks (Table 1) that guide the development

process.

Table 1 Tasks of the engine structural integrity program

Task I: Design information

Development plans

ENSIP master

Durability and damage control

Material and process characterization

Corrosion prevention and control

Inspection and diagnostics

Operational requirements

Design service life and usage requirements

Design criteria

Task II: Design analysis and material characterization and development tests

Design duty cycle

Material characterization

Design development tests

Structural/thermal analysis

Installed engine inspectability

Manufacturing and quality control

Task III: Component and core engine tests

Component tests

Strength

Vibration

Damage tolerance

Durability

Core engine tests

Thermal survey

Vibration strain and flutter boundary survey

Task IV: Ground and flight engine tests

Ground engine tests

Strength

Damage tolerance

Accelerated mission test

Thermal survey

Vibration strain and flutter boundary survey

Flight engine tests

Fan strain survey

Thermal survey

Installed vibration

Deterioration

Task V: Engine life management

Updated analyses

Structural maintenance plan

Operational usage survey

Individual engine tracking

Durability and damage tolerance control actions (production)

The Engine Structural Maintenance Plan represents the output of the ENSIP program. This plan identifies and

defines individual part life limits, the necessary inspection periods for each fracture-critical part, and the inspection

procedure. The basic components of the Engine Structural Maintenance Plan are as follows:

•

Structural safety is obtained in ENSIP by requiring a structure with a damage tolerance that is capable

of accommodating flaws induced either in manufacture or service

•

Durability design requirements stipulate that the economic life of the engine must exceed the specified

design service life of the aircraft when flown to the design usage spectra

• Maintainability criteria

require that old parts fit and function with new parts, that repair life be defined,

and that inspectability and structural diagnostics be designed into the engine and its components

• A materials and process characterization plan controls materials develo

pment through key engine

development points

• Environmental definition requirements specify the thermal, dynamic, and steady-

state stress; the stress

spectra; and the component sensitivities

• A comprehensive ground test policy is utilized to ensure complian

ce with safety, durability, and

maintainability requirements

• A usage and tracking policy is used to form the basis of an engine life management program

Fracture Control Philosophy

William D. Cowie, United States Air Force, Aeronautical System Division, Propulsion Directorate

Overview of ENSIP

In the past 16 years, a large number of structural problems have occurred in USAF gas turbine engines. Many of these

were safety problems that resulted in loss of aircraft, and an even greater number affected durability, causing a high level

of maintenance and modification costs. All of these problems have adversely affected fleet readiness. The Engine

Structural Integrity Program was intended to reduce these problems substantially and was developed based on the

following specific lessons:

• It is unrealistic (and can be dangerous) to assume defect-free structure in safety-of-flight components

•

Critical parts (and part details) and potential failure modes must be identified early and appropriate

control measures implemented

• Internal thermal and vibratory environments must be identified early in the engine development

• Predicted analytical stresses must be verified by test for complex components

• Materials and processes must be adequately characterized (particularly, the fracture properties)

• Design stress spectra, component test spectra, and full-

scale engine test spectra must be based on the

anticipated service usage of the engine, that is, accelerated mission-related testing

• Potential engine/airframe structural interactions must be defined and accounted for

•

Management procedures (such as individual engine tracking procedures and realistic inspection and

maintenance requirements) must be defined and enforced

The Engine Structural Integrity Program was established by the Air Force to provide an organized and disciplined

approach to the structural design, analysis, development, production, and life management of gas turbine engines, with the

goal of ensuring engine structural safety, increased service readiness, and reduced life cycle costs. The five major tasks

associated with ENSIP are the development of design information; design analysis and component and material

characterization; component and core engine testing; ground and flight engine testing; and production quality control and

engine life management. Each major task is subdivided into a number of subtasks (Table 1) that guide the development

process.

Table 1 Tasks of the engine structural integrity program

Task I: Design information

Development plans

ENSIP master

Durability and damage control

Material and process characterization

Corrosion prevention and control

Inspection and diagnostics

Operational requirements

Design service life and usage requirements

Design criteria

Task II: Design analysis and material characterization and development tests

Design duty cycle

Material characterization

Design development tests

Structural/thermal analysis

Installed engine inspectability

Manufacturing and quality control

Task III: Component and core engine tests

Component tests

Strength

Vibration

Damage tolerance

Durability

Core engine tests

Thermal survey

Vibration strain and flutter boundary survey

Task IV: Ground and flight engine tests

Ground engine tests

Strength

Damage tolerance

Accelerated mission test

Thermal survey

Vibration strain and flutter boundary survey

Flight engine tests

Fan strain survey

Thermal survey

Installed vibration

Task V: Engine life management

Updated analyses

Structural maintenance plan

Operational usage survey

Individual engine tracking

Durability and damage tolerance control actions (production)

The Engine Structural Maintenance Plan represents the output of the ENSIP program. This plan identifies and

defines individual part life limits, the necessary inspection periods for each fracture-critical part, and the inspection

procedure. The basic components of the Engine Structural Maintenance Plan are as follows:

•

Structural safety is obtained in ENSIP by requiring a structure with a damage tolerance that is capable

of accommodating flaws induced either in manufacture or service

• Durabi

lity design requirements stipulate that the economic life of the engine must exceed the specified

design service life of the aircraft when flown to the design usage spectra

• Maintainability criteria require that old parts fit and function with new parts, t

hat repair life be defined,

and that inspectability and structural diagnostics be designed into the engine and its components

•

A materials and process characterization plan controls materials development through key engine

development points

• Environmental definition requirements specify the thermal, dynamic, and steady-

state stress; the stress

spectra; and the component sensitivities

•

A comprehensive ground test policy is utilized to ensure compliance with safety, durability, and

maintainability requirements

• A usage and tracking policy is used to form the basis of an engine life management program

Fracture Control Philosophy

William D. Cowie, United States Air Force, Aeronautical System Division, Propulsion Directorate

ENSIP and Fracture Control Philosophy Policy

Damage tolerance is defined as the ability of the engine to resist failure due to the presence of flaws, cracks, or other

damage for a specified period of usage. The damage tolerance or fracture control philosophy used in ENSIP is shown in

Fig. 4. Components are designed for crack growth so that the safety limit exceeds two times the required inspection

interval. The safety limit or residual life is the time for assumed initial flaws to grow and cause failure. Because the

requirement is to inspect at one-half the safety limit, the design goal for the safety limit is two times the required design

life (that is, no inspections). The minimum design requirement for the safety limit is two times the planned depot visit

interval. An important aspect of the damage tolerance requirement is that it applies only to fracture-critical components.

Fig. 4 Damage tolerance approach to life management of cyclic-

limited engine components. The safety limit or

residual life is the time for the initial flaw to grow and cause failure. The size of the initial flaw, a

, is based on

the inspection method or material defect distribution (for imbedded defects).

Fracture-critical components are defined as those components whose failure will result in probable loss of the

aircraft due to noncontainment or, for single-engine aircraft, power loss that presents sustained flight because of direct

part failure or by causing other progressive part failures. Damage tolerance requirements are applied only to fracture-

critical components (that is, components that must maintain their integrity during flight) and not, in general, to durability-

critical components (that is, components that affect maintenance schedules). As expected, component classification is

affected by aircraft engine configuration (single engine or multiengine). Component classification is established early and

is identified in the contract.

Initial Flaw Size. Initial flaws are assumed to exist in fracture-critical components. Experience has shown that

premature cracking (that is, crack initiation prior to the LCF limit) occurs at high-stress areas and where components

initially contained both material- and manufacturing-related quality variations (voids, inclusions, machining marks,

scratches, sharp cracks, and so on). The fracture control or damage tolerance requirement assumes a sharp crack as the

initial flaw when characterizing these abnormal initial conditions. The assumed initial imbedded flaw sizes are based on

the intrinsic material defect distribution or the NDE methods to be used during manufacture. The assumed surface flaw

size also depends on the NDE capability. An inspection reliability of 90% probability of detection (POD) at the lower-

bound 95% confidence level (CL) is required for the assumed initial flaw sizes.

The assumed initial flaw size to account for intrinsic material defect distribution should encompass 99.99% of the defect

population if a scatter factor of two is used to establish the inspection interval, or 99.9% if a scatter factor of one is used.

If embedded defects cannot be inspected in service, the 99.99 percentile (or the 99.9 percentile) is used to satisfy the

design life requirement.

An initial flaw size not less than 0.75 mm (0.030 in.) in length (for surface flaws) or 0.4 × 0.4 mm (0.015 × 0.015 in.) in

size (corner cracks) for nonconcentrated stress areas (bores, webs, and so on) is required. Initial flaw sizes for other

surface locations (holes, fillets, scallops, and so on) will be consistent with the demonstrated capability (90% POD/95%

CL) of the inspection systems proposed for use. It is recommended that the initial design and sizing of components be

based on 0.75 mm (0.030 in.) long surface flaws or 0.4 × 0.4 mm (0.015 × 0.015 in.) corner cracks at all locations. This

design recommendation is based on the initial flaw size that can be detected with fluorescent penetrant inspection. This

includes fully automated fluorescent penetrant inspection systems that are being developed to meet the 0.75 mm (0.030

in.) and 0.4 × 0.4 mm (0.015 × 0.015 in.) inspection criteria.

These flaw sizes are intended to represent the maximum size of the damage that can be present in a critical location after

manufacture and/or inspection. The specification of these flaw sizes is based on the demonstrated flaw detection

capability of the nondestructive inspection (NDI) method. During design of the components, the assumed initial flaw size

that is appropriate for various NDI methods is:

• 0.75 mm (0.030 in.) surface length where the NDI method is fluorescent penetrant inspection

• 0.25 mm (0.010 in.) surface length where the NDI method is eddy current or ultrasonic inspection

• 1.3 mm

(0.002 in.

) area for imbedded defects utilizing ultrasonic inspection

• 5 mm (0.200 in.) surface length and imbedded sphere = 0.2 × thickness for weldments

•

When initial flaw sizes are based on material defect distribution, selected size shall encompass 99.99%

of the distribution

• Demonstration that assumed flaw sizes can be reliably detected with a 90% POD and a 95% CL

The capabilities of the NDI method must be demonstrated by the contractor. The design of NDE reliability experiments is

discussed in the article "NDE Reliability Data Analysis" in this Volume.

Residual strength is defined as the load-carrying capability of a component at any time during the service exposure

period, considering the presence of damage and accounting for the growth of damage as a function of exposure time. The

requirement is to provide limit load residual strength capability throughout the service life of the component. In other

words, the minimum residual strength for each component (and location) must be equal to the maximum stress that occurs

within the applicable stress spectra based on the design duty cycle. Normal or expected overspeed due to control system

tolerance and engine deterioration is included in the residual strength requirement, but fail-safe conditions, such as burst

margin, are excluded. The residual strength requirement is illustrated in Fig. 5.

Inspection Intervals. It is highly desirable to have no

damage tolerance inspections required during the design

lifetime of the engine. This in-service noninspectable

classification requires that components be designed such

that the residual life or safety limit be twice the design life.

Designing components as in-service noninspectable is a

requirement for those components or locations that cannot

be inspected during the depot maintenance cycle.

However, the weight penalty incurred to achieve a safety

limit/residual life/damage growth interval twice the design

life may be prohibitive on some components/locations.

Therefore, in-service inspections will be allowed on some

components subject to justification. The basis for the

justification is characterization of the costs as a function of

the requirements as established by trade studies. Cost is

usually expressed in terms of weight or life cycle cost, and

the requirement in terms of safety limit/residual

life/damage growth interval.

The depot or base-level inspection interval for damage

tolerance considerations should be compatible with the

overall engine maintenance plan. Once again, it is highly

desirable that the inspection interval be equal to the design

service life of the parts in the hot gas path (that is, the hot-

part design service life, which is equal to one-half the

design lifetime of the engine) because this is the expected

minimum depot or maintenance interval for the engine or

module. It is required that the minimum damage tolerance

inspection interval be contained in the contract

specification.

Flaw Growth. It is required that the assumed initial flaw sizes will not grow to critical size and cause failure of a

component due to the application of the required residual strength load in two times the inspection interval. The flaw

growth interval is set equal to two times the inspection interval to provide a margin for a variability that exists in the total

process (that is, inspection reliability, material properties, usage, stress predictions, and so on). Factors other than two

should be used when individual assessments of the variables that affect crack growth can be made (for example, to

account for observed scatter in crack growth during testing).

Fig. 5 Diagram of the residual strength requirement

It is important that the effects of vibratory stress on unstable crack growth be accounted for in establishing the safety

limit. Experience shows that the threshold crack size can be significantly less than the critical crack size associated with

the material fracture toughness, depending on the material, the major stress cycle, and the vibration stress. As shown in

Fig. 6, the conventional Goodman diagram may not disclose the true sensitivity of initial defects to vibratory stresses. The

threshold crack size must be established at each individual sustained-power condition (idle, cruise, intermediate) using the

appropriate values of steady stress and vibratory stress. The smallest threshold crack size will be used as a limiting value

in calculating the safety limit if it is less than the critical crack size associated with the material fracture toughness.

Fig. 6 Interac

tion of vibratory stress and initial flaws. (a) Large vibratory stress required to initiate crack. (b)

Low vibratory stress will propagate cracks. The crack growth threshold, A

, represents the threshold of vibratory

motion that will cause the growth of a given crack size.

Fracture Control Philosophy

William D. Cowie, United States Air Force, Aeronautical System Division, Propulsion Directorate

Fracture Control Verification

Verification that the fracture control policy is met is accomplished by the development and implementation of a Damage

Tolerance Control Plan, by analysis and test, and by the implementation of reliable inspection methods during

manufacture and field/depot maintenance.

A Damage Tolerance Control Plan is prepared that identifies and schedules each of the tasks and interfaces in the

functional areas of design, materials selection, tests, manufacturing control, and inspection. Specific tasks that are

addressed in the Damage Tolerance Control Plan are:

• Trade studies for design concepts/material/weight/performance/cost

• Analysis

• Development and qualification tests

• Fracture-critical parts list

• Zoning of drawings

• Basic materials fracture data

• Material properties controls

• Traceability

• NDI requirements