Durability and Damage Tolerance Certification for Additive

Durability and Damage

Tolerance Certification for

Additive Manufacturing (AM)

of Aircraft Structural Parts

19 March 2019

Chuck Babish

U.S. Air Force

2019 Aircraft Structural Integrity Symposium

Melbourne, Australia; 18 – 20 March 2019

2

Outline

• Overview

• Aircraft Structural Integrity Program (ASIP)

Material & Process (M&P) requirements

• Challenge for Additive Manufacturing (AM) –

accurate prediction of Durability and Damage

Tolerance (DADT)– DADT Analysis (DADTA) foundation

– Surrogate damage

– DADT control

– Parts selection

• Probabilistic risk analysis for AM parts

• Summary

3

Overview

• Aircraft Structural Integrity Program (ASIP) MIL-STD-1530

establishes the framework for DADT CertificationTask I Task II Task III Task IV Task V

Design InformationDesign Analyses &

Development TestingFull-Scale Testing

Certification & Force

Management DevelopmentForce Management Execution

1. ASIP Master Plan 1. Materials and Structural

Allowables1. Static Tests 1. Structural Certification 1.L/ESS Execution

2. Design Service Life & Design

Usage 2. Loads Analysis

2. First Flight Verification Ground

Tests

2. Strength Summary & Operating

Restrictions (SSOR)2. IAT Execution

3. Structural Design Criteria 3. Design Loads/Environment

Spectra 3. Flight Tests

3. Force Structural Maintenance

Plan (FSMP)3. DADTA Updates

4. Durability & Damage Tolerance

Control4. Stress and Strength Analysis 4. Durability Tests

4. Loads/ Environment Spectra

Survey (L/ESS) System

Development

4. L/ESS and IAT System Updates

5. Corrosion Prevention & Control 5. Durability Analysis 5. Damage Tolerance Tests 5. Individual Aircraft Tracking (IAT)

System Development 5. NDI Updates

6. Nondestructive Inspection 6. Damage Tolerance Analysis 6. Climatic Tests 6. Force Management Database

Development

6. Structural Risk Analysis

Updates

7. Selection of Materials,

Processes, Joining Methods &

Structural Concepts

7. Corrosion Assessment 7. Interpretation & Evaluation of

Test Findings7. Technical Orders

7. CPC Plan & Corrosion

Assessment Updates

8. Sonic Fatigue Analysis 8. Resolution of Test Findings 8. Analytical Condition Inspection

9. Vibration Analysis 9. FSMP Updates

10. Aeroelastic and

Aeroservoelastic Analysis10. Technical Orders Updates

11. Mass Properties Analysis 11. Repairs

12. Survivability Analysis 12. Structural Maintenance

Database Execution

13. Design Development Tests 13. Structural Certification

Updates

14. Structural Risk Analysis14. Economic Service Life

Analysis Updates

15. Economic Service Life

Analysis15. Others as Required

4

Overview

• ASIP MIL-STD-1530 includes

key requirements for

selection of M&P

– Contained in Task I Design

Information section

– Applicable throughout life

cycle whenever new M&P are

considered

• Intent of Task V Force

Management Execution

(sustainment phase) is to revisit

any prior Task as required

– Applicable to AM pursuits!

Task I

Design Information

1. ASIP Master Plan

2. Design Service Life & Design

Usage

3. Structural Design Criteria

4. Durability & Damage Tolerance

Control

5. Corrosion Prevention & Control

6. Nondestructive Inspection

7. Selection of Materials,

Processes, Joining Methods &

Structural Concepts

5

Outline

• Overview







– DADT control

– Parts selection


• Summary

6

ASIP Selection of M&P

(MIL-STD-1530D Paragraph 5.1.7)

• …“materials, processes, joining methods, and structural

concepts shall be selected to result in a structurally

efficient, cost-effective aircraft structure that meets the

strength, rigidity, durability, damage tolerance, and other

requirements...”

• “Prior to a commitment to new materials…an evaluation

of their stability, producibility, characterization of

mechanical and physical properties, predictability of

structural performance, and supportability shall be

performed...” (aka “Lincoln’s 5 Factors”)

• “The risk associated with the selection of the new

materials, processes…shall be estimated and risk

mitigation actions defined”

“Satisfy Lincoln’s 5 Factors”

7

ASIP M&P Requirements

Lincoln’s 5 Factors (1 of 3)

1. Stability - “Maturity of material, process, and joining

method selections shall be evaluated to determine if

consistent and repeatable quality and if predictable

costs can be achieved to meet system performance and

production requirements. Process parameters and

methods shall be established and controlled via

specifications, standards, and manufacturing

instructions”

2. Producibility - “Material, process, and joining method

selections shall be evaluated to determine if scale-up to

production sizes and rates can be achieved without

adversely affecting performance, costs, and quality. The

material, process, and joining method selections shall

consider inspectability during the manufacturing

process”

8



3. Characterization of mechanical and physical properties -

“Material, process, and joining method selections shall

be characterized to determine mechanical and physical

properties for the appropriate environments in the as-

fabricated condition using the manufacturing processes

and joining methods. Key mechanical properties include

but are not limited to: strength, elongation, fracture

toughness, damage growth rates, fatigue, stress

corrosion and damage growth rate thresholds. Key

physical properties include but are not limited to:

density, corrosion resistance, damage population,

surface reflectivity, thermal stability, coefficient of

thermal expansion, fire resistance, fluid resistance, and

surface roughness.”

9



4. Predictability of structural performance - “Material,

process, and joining method selections shall be

evaluated to determine if validated analysis methods

and/or empirical methods are established to enable

accurate prediction of structural performance (for

example, strength, rigidity, durability, damage

tolerance). If validated methods don’t exist at the time of

selection, risk mitigation actions shall be established.”

5. Supportability - “Material, process, and joining method

selections shall be evaluated to determine if cost-

effective inspection and repair methods are either

available or can be developed in a timely manner

considering the sustainment environment throughout

the entire life cycle…”

10

Outline

• Overview







– DADT control

– Parts selection


• Summary

11

Most Difficult Challenge for AM of

Aircraft Structural Parts

• Validated analysis method for accurate

prediction of durability and damage tolerance

(ASIP M&P selection factor 4) which requires:

1. Established DADTA foundation

2. Established surrogate damage

3. Established DADT control

4. Methodical parts selection

“Work the Hardest Problem Up Front”

12

1. DADTA Foundation

Introduction

• How many locations are potentially critical and

therefore require DADTA?

Subtractive Manufacturing Additive Manufacturing

2 to 3 Locations? 5 to 7 Locations?

13

1. DADTA Foundation

Requirements (1 of 2)

1.1 Develop validated DADTA method

– Results in reasonable correlation with test data

– Accounts for complex geometries, load interaction,

residual stresses, etc. as needed

1.2 Generate material data

– Material data (fatigue crack growth rates, fracture

toughness) developed using the same processes,

machines, build orientation, etc. for the production parts

– Sufficient replicates to demonstrate material

consistency recognizing AM technology is particularly

susceptible to producing parts with discontinuities that

can serve as sites for crack nucleation reinforcing the

need for statistical characterization

14

1. DADTA Foundation


1.3 Establish DADT criteria– Specific to the AM application for use in the DADTA

and to establish part-specific inspections

– For reference, current DADT criteria includes: • Average usage for damage tolerance stress spectra

• 90th percentile usage for durability stress spectra

• Nominal part dimensions

• Average crack growth rate

• Lower bound fracture toughness

• Typically no explicit residual stresses

• Specified initial crack sizes

• Factor of 2 on DADTA predictions to establish the initial

inspection requirement

• Note: If material data scatter is judged to be too high,

adjustments to 1 or more criteria are made

15

2. Surrogate Damage

Introduction (1 of 5)

• MIL-STD-1530 defines damage as “any flaw, defect,

crack, corrosion, disbond, delamination, discontinuity,

or other type that degrades, or has the potential to

degrade, the performance of the affected component”

• MIL-STD-1530 states that “damage tolerance criteria

shall be applied to all safety-of-flight structure” and

“include establishment of surrogate damage types,

sizes, orientations, locations, with consideration of all

phases of the life cycle to include: material processing,

shipping, handling, manufacturing, flight operations, and

maintenance”

Surrogate Damage has 4 Attributes

16

2. Surrogate Damage


• For DADT of structural parts made from wrought metallic

materials, surrogate damage type is a fatigue crack, in

the most critical orientation, in all critical locations (this

covers 3 of the 4 damage attributes)

• USAF developed Equivalent Initial Damage Size (EIDS)

distribution data to establish surrogate damage sizes for

use in the DADTA in the 1970s– MIL-STD-1530 states that the EIDS “is an analytical

characterization of the initial quality of the aircraft structure at

the time of manufacture, modification, or repair. The EIDS

distribution is derived by analytically determining the initial

damage size distribution that characterizes the measured

damage size distribution observed during test or in service”

• EIDS data used in structural risk analysis during past 20

years indicate this early work is still representative

17

2. Surrogate Damage


0.000 0.005 0.010 0.015 0.020 0.025 0.030

Pro

ba

bilit

y D

en

sit

y F

un

cti

on

Equivalent Initial Damage Size (inches)

Wrought Aluminum, Individual & Combined Data

All 1.1

2.1 3.1

4.1 4.2

4.3 4.4

4.5 5.1

6.1 7.1

Initial Crack Size Assumption for

Durability

("Upper Bound of Normal Manufacturing

Quality")518 Total Cracks:

7 Aircraft

11 Materials

18

2. Surrogate Damage


0.000 0.005 0.010 0.015 0.020 0.025 0.030 0.035 0.040 0.045 0.050 0.055 0.060

Pro

ba

bilit

y D

en

sit

y F

un

cti

on


Wrought Aluminum, Combined Data

Initial Crack Size

Assumption for Damage Tolerance

("Rogue Damage Size")

Initial Crack Size

Assumption for Durability

("Upper Bound of Normal

Manufacturing Quality")

~Best NDI Capability

Combined

518 Cracks

19

2. Surrogate Damage


1.0E-07

1.0E-06

1.0E-05

1.0E-04

1.0E-03

1.0E-02

1.0E-01

1.0E+00

0.000 0.010 0.020 0.030 0.040 0.050 0.060 0.070 0.080 0.090 0.100

Exceed

an

ce P

rob

ab

ilit

y



All 1.1

2.1 3.1

4.1 4.2

4.3 4.4

4.5 5.1

6.1 7.1

20

2. Surrogate Damage


2.1 Establish surrogate damage type(s)– Generate data that demonstrates the surrogate

damage type(s) covers the full range of potential

damage such as lack of fusion, porosity, cracking,

and lack of penetration in AM parts

2.2 Establish surrogate damage locations– Most critical location(s) for the part

2.3 Establish surrogate damage orientation(s)– Most critical orientation(s) for the part

21

2. Surrogate Damage


2.4 Establish surrogate damage sizes– Damage size for Normal Control (NC) parts based on a

probability of exceeding the EIDS of 1x10-1; but not

less than 0.01 inches

– Damage size for Durability Critical (DC) parts based on

a probability of exceeding the EIDS of 1x10-3; but not


– Damage size for Fracture Critical (FC) parts based on

a probability of exceeding the EIDS of 1x10-7; but not


Note: NC, DC, and FC Defined in Section 4

1.0E-07

1.0E-06

1.0E-05

1.0E-04

1.0E-03

1.0E-02

1.0E-01

1.0E+00

0.000 0.010 0.020 0.030 0.040 0.050 0.060 0.070 0.080 0.090 0.100

Ex

ce

ed

an

ce

Pro

ba

bilit

y



All 1.1

2.1 3.1

4.1 4.2

4.3 4.4

4.5 5.1

6.1 7.1

22

2. Surrogate Damage


NC

DC

FC

2 of 11 > 0.01”

Combined < 0.01”

6 of 11 > 0.01”

Combined ~0.02”

2 of 11 > 0.05”

Combined ~0.05”

23

3. DADT Control


3.1 Demonstrate process control– Address the stability and producibility factors

– Include residual stress management efforts such as

stress relief and additional manufacturing steps such

as heat-treating, Hot Isostatic Pressing (HIPing), and

machining to achieve surface finish requirements

3.2 Generate mechanical property data– Statistically-based design data from specimens using

the “production” processes and include worst-case

orientations and regions of the part

24

3. DADT Control


3.3 Develop, validate, verify, and implement Non

Destructive Investigation (NDI)– Enable effective Quality Assurance (QA)

– If an NDI-based initial damage size assumption is

used (not preferred), verify the NDI capability for all

damage types, damage orientations, and damage

locations within specific regions of each part

considering access, part thickness, and other

important factors (this is not a trivial task!!!)

25

3. DADT Control


3.4 Establish dimensional control and verification– After the AM process is completed and after all post-

processing is completed

3.5 Dissect “Production” parts and characterize

material– Perform mechanical testing of specimens made from

excised material to verify properties to include worst-

case orientations and regions of the part

– Perform metallographic and micrographic

examinations to determine if defects/damage are

within expectations and to evaluate/demonstrate NDI

performance

26

3. DADT Control


3.6 Conduct mechanical property testing of

production parts– Fabricate and test “Witness” coupons from each

production part to verify properties, and to provide

data and experience to support the use of the AM

process to more critical/complex parts

– Fabricate the coupons using the same AM parameters

as production parts to include: scan schedule,

interpass time, and orientation on build plate

27

3. DADT Control


3.7 Establish accept/reject criteria– Considering all of the DADT control results

– This includes establishing criteria to reject an AM part

based on the frequency of defects/damage, even if

none of the defects/damage exceed the initial damage

size assumption

28

4. Methodical Parts Selection


• MIL-STD-1530 establishes part classification & definitions

29



• Safety-of-flight– Fracture-Critical (FC) traceable part

• Either single load path or judged to require serialization and

traceability

– Fracture-Critical (FC) part• Not single load path nor judged to require serialization and

traceability

• Not safety-of-flight– Durability-Critical (DC) part

• Judged to require additional controls beyond those for normal-

controls parts

– Normal Controls (NC) part• Standard aerospace practices are sufficient in the design,

manufacturing, and maintenance of the part to ensure

structural integrity

30


Requirements

4.1 Start with NC parts only– Lowest risk if unanticipated issues occur

– Still requires durability crack growth analysis, all

supporting material data, surrogate damage, etc.

– Expectation: an aggregate of NC parts will provide data

and experience necessary to evaluate risk of

proceeding with DC aircraft structural parts due to:• Varying geometries & complexities

• Multiple lots of materials

• Multiple AM machines & operators

• Characterized material property variability

• Additional data for EIDS characterization

31


Requirements

4.2 Then move to DC parts– Next level of risk

– Use NC & DC parts experience to provide confidence

for pursuing FC parts

4.3 Then pursue FC parts– Requires damage tolerance analysis and different

surrogate damage sizes (as a minimum)

– Requires improved understanding of variability

– Requires improved understanding of supportability

– Requires additional NDI development & validation

– Requires robust accept/reject criteria

32

DADT Certification of AM Parts

Summary (1 of 2)

1. Established DADTA foundationa. Method

b. Data

c. Criteria

2. Established surrogate damagea. Type

b. Orientation

c. Location

d. Size

33

DADT Certification of AM Parts

Summary (2 of 2)

3. Established DADT controla. Process control

b. Mechanical property testing for design data

c. NDI

d. Dimensional control

e. Dissection of production parts

f. Mechanical property testing of production parts

g. Accept/reject criteria

4. Methodical parts selectiona. Normal control parts, then

b. Durability critical parts, then

c. Fracture critical parts

34

Outline

• Overview







– DADT control

– Parts selection


• Summary

35

Probabilistic Risk Analysis

Introduction

• Modified Lincoln method, see references:– Lincoln, J.W., ASD-TR-80-5035, “Method for Computation of

Structural Failure Probability for an Aircraft,” July 1980

– UDR-TR-2005-00240, “PRoF v3 PRobability Of Fracture Aging

Aircraft Risk Analysis Update,” December 2005

• Freudenthal method, see references:– Domyancic, L.; McFarland, J.; Cardinal, J.; and Burnside, H.,

“Review of Methods for Single Flight Probability of Failure

(SFPOF),” Southwest Research Institute, March 2011

– Brussat, T. R., “Recommended Methodology Updates to Improve

Single Flight Probability of Failure Estimation,” 2012 ASIP

Conference, San Antonio TX

– UDR-TR-2015-15, “PROF v3.2 Probability Of Failure – A-10 Fleet

Management Analysis Update,” July 2015

36


Inputs & Output

1.E-09

1.E-08

1.E-07

1.E-06

1.E-05

1.E-04

1.E-03

0 2000 4000 6000 8000 10000 12000

Sin

gle

Flig

ht P

robabili

ty o

f F

ailu

re

Flight Hours

SFPoF

Similar Details

Hours per Flight0.0

2.0

4.0

6.0

8.0

10.0

12.0

14.0

16.0

18.0

20.0

0 1 2 3 4 5 6

K/s

igm

a

Crack Size (in.)

Alpha Plot

BaselineK/σ vs

Crack

Size

1.E-07

1.E-06

1.E-05

1.E-04

1.E-03

1.E-02

1.E-01

1.E+00

0.00 5.00 10.00 15.00 20.00 25.00 30.00

Exce

ed

an

ce

Pro

ba

bili

ty

Stress

Max Stress per Flight

BaselineMax

Stress

per Flight

Initial

Crack

Size 0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

0.000 0.005 0.010 0.015 0.020

Cum

ula

tive

Pro

ba

bility

Crack Size (in.)

Initial Crack Size Distribution

Baseline

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1.0

0.00 0.20 0.40 0.60 0.80 1.00

Cu

mu

lative

Pro

ba

bili

ty

Crack Size (Inches)

Probability of Detection

Baseline

Probability

of

Detection

Inspection

Intervals

0

0.5

1

1.5

2

2.5

3

3.5

4

0 5000 10000 15000 20000 25000

Cra

ck S

ize

(in

.)

Flight Hours

Crack Growth Curve

Baseline

Crack

Growth

Curve

Fracture

Toughness0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1.0

20 25 30 35 40 45 50

Cu

mu

lative

Pro

ba

bili

ty

Fracture Toughness (Ksi * Inch ^ 0.5)

Fracture Toughness

Baseline

Repair

Crack

Size 0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

0.000 0.005 0.010 0.015 0.020 0.025 0.030

Cum

ula

tive P

rob

ab

ility

Crack Size (in.)

Repair Crack Size Distribution

Baseline

ඵ

37

Risk Analysis Inputs & Output

Related to DADT for AM Challenges

1.E-09

1.E-08

1.E-07

1.E-06

1.E-05

1.E-04

1.E-03

0 2000 4000 6000 8000 10000 12000

Sin

gle

Flig

ht P

robabili

ty o

f F

ailu

re

Flight Hours

SFPoF

Similar Details

Hours per Flight0.0

2.0

4.0

6.0

8.0

10.0

12.0

14.0

16.0

18.0

20.0

0 1 2 3 4 5 6

K/s

igm

a

Crack Size (in.)

Alpha Plot

BaselineK/σ vs

Crack

Size

1.E-07

1.E-06

1.E-05

1.E-04

1.E-03

1.E-02

1.E-01

1.E+00

0.00 5.00 10.00 15.00 20.00 25.00 30.00

Exce

ed

an

ce

Pro

ba

bili

ty

Stress

Max Stress per Flight

BaselineMax

Stress

per Flight

Initial

Crack

Size 0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

0.000 0.005 0.010 0.015 0.020

Cum

ula

tive

Pro

ba

bility

Crack Size (in.)

Initial Crack Size Distribution

Baseline

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1.0

0.00 0.20 0.40 0.60 0.80 1.00

Cu

mu

lative

Pro

ba

bili

ty

Crack Size (Inches)

Probability of Detection

Baseline

Probability

of

Detection

Inspection

Intervals

0

0.5

1

1.5

2

2.5

3

3.5

4

0 5000 10000 15000 20000 25000

Cra

ck S

ize

(in

.)

Flight Hours

Crack Growth Curve

Baseline

Crack

Growth

Curve

Fracture

Toughness0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1.0

20 25 30 35 40 45 50

Cu

mu

lative

Pro

ba

bili

ty

Fracture Toughness (Ksi * Inch ^ 0.5)

Fracture Toughness

Baseline

Repair

Crack

Size 0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

0.000 0.005 0.010 0.015 0.020 0.025 0.030

Cum

ula

tive P

rob

ab

ility

Crack Size (in.)

Repair Crack Size Distribution

Baseline

ඵ

DADTA

Found.

DADTA

Found.

DADTA

Found.

Surrogate

Damage

DADT

Control

38


Priorities• #1: Get the loads/environment and usage severity right

for each aircraft– Used in crack growth analysis

– Used in equivalent flight hours (EFH) calculations

– Used to generate max stress per flight distribution

• #2: Get the crack growth model right– Used to convert actual flight hours into EFH so that fleet and test

results can be combined

– Used to convert crack sizes at EFH into EIDS values to generate

flaw size distribution

– Used to project flaw size distribution in time

– Used to determine when failure criterion is reached

• #3: Get the EIDS distribution right and update it as

additional inspection results are obtained– Must consider that NDI is used to find/miss cracks and therefore

directly influences result

• #4: Get the NDI PoD function right

39

Outline

• Overview







– DADT control

– Parts selection


• Summary

40

Summary

• DADT certification of structural parts requires significant

test data to characterize the variability in DADT properties

and to establish the initial damage type and size

assumptions used in the DADTA

• The pursuit of AM for structural parts should begin with

NC parts only

• If the test data and experience from building hundreds of

various NC parts supports it, the pursuit of DC parts can

begin

• Only when the test data and experience from building

many hundreds of various NC and DC parts exists and is

favorable, should the pursuit of FC parts begin

• This approach is expected to be more efficient in the long

run than performing significant testing for part-specific

AM applications only

41

Documents

Durability and Damage Tolerance Certification for Additive