**Airworthiness Assurance R&D Branch**

1996 Research Accomplishments

John M. Fabry, Branch Manager

US Department of Transportation

FAA William J. Hughes Technical Center

Atlantic City Airport, NJ 08405

Table of Contents

Title Page

Advanced Materials Research

Certification Methodology for Stiffener Terminations

Effects of Stiffener/Rib Separation on Damage Growth and Residual
Strength

Structural Integrity Evaluation of the Lear Fan 2100 Aircraft

The Effect of Preloading on Fatigue Damage in Composite Structures: Part
1

Comparative Evaluation of Failure Analysis Methods for Composite
Laminates

Crashworthiness

Survey and Analysis of Rotorcraft Flotation Systems

Transport Water Impact and Ditching Performance

Transport Water Impact, Part II

Propulsion Systems and Fuels

Spin Synchronous X-Ray Sinography for Nondestructive Imaging of Turbine
Engines under Load

Aging Aircraft

Characterization of Early Stages of Corrosion Fatigue in Aircraft Skin

Antisymmetric Buffet Loads on Horizontal Stabilizers

An Analysis of Ground-Flight Loads Measured on a B-727

Incorporating the Dripless Bubbler with a Scanner for Ultrasonic C-Scan
Imaging of Defects in Bonded Aircraft Structures

An Overview of the Engine Titanium Consortium

An Overview of the Contaminated Billet Study

An Element-Free Galerkin Method for Three-Dimensional Crack Growth
Modeling

The Effects of Beam Distortion On The Strength Of Ultrasonic Signals

Transducer Design for High Sensitivity Ultrasonic Inspection of Titanium

Eddy Current Probe Design for Edge Effect Reduction

Air-Coupled C-Scan of Boeing Repair Coupon

Modeling the Relationship between Microstructure and Backscattered Noise

Ultrasonic Microscopy of Synthetic Spherical TiN Inclusions in Forged
Ti-6Al-4V

Development of Ultrasonic Models to Guide the Design and Evaluation of
Inspection Systems for Titanium Rotating Components

Modeling Ultrasonic Inspectability in Forgings

Improved Methodology for Predicting Probability of Detection (POD) of
Synthetic Hard Alpha Inclusions in Titanium

Hand-Held, Flexible Eddy Current Probe for Inspection of Curving
Surfaces

Eddy Current Probe Design

Fatigue Crack Initiation and Growth in Riveted Specimens

A Study of Fatigue Crack Generation and Growth in Riveted Aluminum Clad
2024-T3 Specimens

Economical Eddy Current

Eddy Current Methods for Crack Detection

Eddy Current Methods for Corrosion Detection

Frequency-Shifted Sagnac Interferometer for Ultrasound Measurement

Self-Compensating Ultrasonic Technique

Self-Focusing Ultrasonic System

Quantitative Thermal Wave Imaging of Corrosion Thinning

Nondestructive Evaluation Parameters - Failure Property Relationships of
Adhesive Bonds

Feasibility Study of a Rotorcraft Health and Usage Monitoring System
(HUMS): Usage and Structural Life Monitoring Evaluation

Feasibility Study of a Rotorcraft Health and Usage Monitoring System
(HUMS): Results of Operator\'s Evaluation

Corrosion and Corrosion Fatigue of Airframe Materials

Stochastic Modeling of Antisymmetric Buffet Loads on Horizontal
Stabilizers in Massively Separated Flows

Fracture Testing Of Large-Scale Thin Sheet Aluminum Alloy

Development Of A D Sight Aircraft Inspection System: Phase II

Light Shaping Diffusers For Improved Visual Inspection Of Aircraft

Investigation of Fuselage Structure Subject to Widespread Fatigue Damage

Automated Inspection of Aircraft

Validation of the Magneto-Optic/Eddy-Current Imager

A Methodology For The Economic Assessment Of Nondestructive Evaluation
Techniques Used In Aircraft Inspection

Tire Test Correlation: Radial Versus Bias-Ply Tires

Flight Loads Data For A Boeing 737-400 In Commercial Operation

Axial Crack Propagation and Arrest in a Pressurized Fuselage

Implications of Corrosion Pillowing on the Structural Integrity of
Fuselage Lap Joints

Introduction of Bonded Composite Doublers to Commercial Aircraft

Corrosion of Aluminum Alloys in the Presence of Fire-Retardant Aircraft
Interior Materials

Maintenance Resource Management (MRM) Training and Crew Resource
Management (CRM) Training: Analysis, Prescriptions, and Proscriptions

Visual Inspection Research Project Report on Benchmark Inspections

Variation in Load Factor experience of Fokker F27 and F28 Operational
Acceleration Exceedance Data

Engineering Approach to Damage Tolerance Analysis of Fuselage Skin
Repairs

The Role of Fretting Fatigue on Aircraft Rivet Hole Cracking

Catastrophic Failure Prevention

Fiber-Reinforced Structures for Small Turbine Engine Containment

**Document Compiled by H. R. Marek**

**FAA William J. Hughes Technical Center**

**Airworthiness Assurance R&D Branch**

**Atlantic City Airport, NJ 08405**

# Advanced Materials Research

## Certification Methodology for Stiffener Terminations

C. H. Shah, H. P. Kan, and M. Mahler[^1]

**Research Objective**

Document the results of a combined experimental and analytical effort
that leads to the formulation of a recommended structural certification
approach for stiffener terminations.

**Approach**

An experimental program was conducted to determine the static strength
and fatigue life of a typical stiffener termination configuration for
commercial aircraft composite structure. The test specimen was analyzed
to predict both the global structural response and the local failure
initiation. Statistical analyses were also conducted to assess the
scatter in static strength and fatigue life.

The test specimen design was based on the results of comprehensive
technology assessment and a detailed stress analysis. An I‑section
stiffener cocured to a soft skin was selected based on the design
concept of the Boeing's B‑737 composite horizontal tail. The dimensions
used for the test specimens were determined from the finite element
analysis used to simulate the actual structural response.

**Accomplishment Description**

The mode of the initial failure was, as expected, stiffener separation
from the skin caused by interfacial stresses. Similar initial failures
were observed in the static compression test specimens with proper
constraints to prevent overall specimen buckling as well as local
buckling. The fatigue behavior of the test specimens was observed to be
typical of composites, which exhibit relatively high threshold and high
life scatter.

The effect of stiffener termination angle, when changed from baseline
75° to 45°, on static strength as well as fatigue life was determined to
be insignificant. The flange fasteners had no effect on the failure
initiation but retarded the growth of the delamination under fatigue
load.

An analysis was conducted to predict the overall specimen response as
well as the failure initiation. The finite element method was used to
determine the relationship between the applied load and the strain
response of the specimen. The predicted strains were compared with the
measured strains at various strain gage locations. Excellent correlation
was found between the measured and predicted strains at key locations. A
local elasticity model was used to predict the initiation of
stiffener/skin separation at the stiffener termination site. The
analytical results were conservative, compared to the test data for all
specimen configurations and loading conditions.

A statistical method was employed to assess the scatter in the static
strength and fatigue life data. The scatter parameters obtained from the
results of this program were comparable to those of composites and
bolted joints.

**Significance**

From the results of this program, a certification approach was
formulated which did not exist previously. In this approach the static
strength of composite structures with a properly designed stiffener
termination can be certified based on a B-­basis design allowable derived
from final failure data.

**Expected Results**

The experimental data generated in this program provide useful data that
is needed to define a certification procedure for composite structures
with stiffener terminations.

**Reference:** Shah, C.; Kan, H.; Mahler, M: Certification Methodology
for Stiffener Terminations, DOT/FAA/AR-95/10, April 1996.

**Point of Contact:** P. Shyprykevich, AAR-431, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-4967, FAX: (609) 485-4004, e-mail:
pshyprykevich@admin.tc.faa.gov

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## Effects of Stiffener/Rib Separation on Damage Growth and Residual Strength

H. P. Kan and M. Mahler[^2]

**Research Objective**

To define the type, extent, and location of discrete source damage which
would meet the requirements for composite structures in commercial
aircraft similar to those in Federal Aviation Regulation (FAR) 25.571
for metal structures.

**Approach**

Two existing composite aircraft structures were selected to demonstrate
the certification methodology developed. Both structures were designed
to satisfy the damage tolerance requirements. These structures are
basically soft wing skins with bonded stringers. This design feature
makes skin/stiffener separation a potential damage mode that threatens
the integrity of the structure. A transport wing was selected as a
representative structure for a large airplane, and a military aircraft
wing was selected as a representative structure for small airplane.

Three competing discrete source damage types were considered: impact
damage, delaminations, and skin/stiffener disbonds. The influence of
these damage types on the residual strength of the two existing
composite structures was analytically determined. The severity of impact
damage and delaminations were analytically compared with that of
skin/stiffener disbonds. Critical disbond sizes were determined as a
function of residual strength and compared to FAR 25.571 in terms of
design limit load for immediately obvious damage.

**Accomplishment Description**

For typically designed composite wing structures, a completely disbonded
stringer represented the most severe damage scenario among the damage
types considered. This type of damage mainly affects bonded or cocured
structures under predominantly compression loads. The local strength at
the damaged location, depending on the design details of the structure,
may be significantly lower than the residual strength due to impact
damage. Because of the large strength reduction, damage tolerance design
based on such a damage scenario would impose a significant weight
penalty to the structure, if the rib spacing is large.

**Significance**

Compliance with Federal Aviation Regulations FAR 25.571 for fail safety
in commercial aircraft is typically demonstrated by analysis and test of
stiffened metal panels with a two-bay crack extending through the
stiffener separating the two bays. A similar compliance methodology for
composite aircraft structures did not exist before this study. This
research will aid in the development of such a methodology.

**Expected Results**

Guidance for compliance to FAR 25.571 for composite aircraft structure
including type and location of damage to be evaluated by analysis and
verified by testing.

**Reference:** Kan, H. P.; Mahler, M.: Effects of Stiffener/Rib
Separation on Damage Growth and Residual Strength, DOT/FAA/AR-95/12, May
1996.

**Point of Contact:** P. Shyprykevich, AAR-431, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-4967, FAX: (609) 485-4004, e-mail:
pshyprykevich@admin.tc.faa.gov

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## Structural Integrity Evaluation of the Lear Fan 2100 Aircraft

H. P. Kan and T. A. Dyer[^3]

**Research Objective**

Conduct a systematic structural integrity evaluation of the damage prone
components of the Federal Aviation Administration (FAA)-owned Lear Fan
2100 aircraft using the methodology developed under the FAA/Navy
programs.

**Approach**

A Lear Fan 2100 aircraft was inspected using nondestructive techniques.
The inspection was conducted on aircraft serial number E009, the third
flying prototype manufactured. FAA personnel indicated that the aircraft
had experienced approximately 230 hours of flight time. The upper wing
skins and the upper fuselage skin, areas considered most likely to
suffer damage during manufacturing and maintenance operations, were
evaluated using ultrasonic and thermographic techniques. A total of 19
defects around the fastener heads in the upper wing skins were
identified by the MAUS (Mobile Automated Ultrasonic System developed by
McDonnell Douglas). In addition, one area of mild porosity in the wing
skin and one area of possible disbond between mating fuselage skins were
also detected. Analytical results showed that these defects are not
severe enough to impose a threat to the integrity of the wing structure.
After an extensive review of the structural configurations, flight
loads, and the full-scale test articles, a damage tolerance evaluation
was conducted for the upper wing skins. The capability of the structure
to tolerate impact damage and assembly induced defects was
systematically evaluated. The upper wing skin was divided into small
regions, based on the arrangement of the substructures and the
distribution of the skin thickness, for the damage tolerance
evaluations.

**Accomplishment Description**

Damage tolerance capability of the upper wing skins against
assembly/manufacturing induced damage was analytically evaluated. The
baseline damage scenario used was defined based on the results of a
recently completed FAA/Navy sponsored program, which generally produced
damage more severe than the defects detected for the E009 aircraft.
Margin of safety and reliabilities of the upper wing skins with the
baseline damage were obtained analytically. In addition, allowable
damage sizes were defined for various damage scenarios.

**Significance**

The certification methodology for composite aircraft structures
developed under a series of FAA/Navy sponsored programs was successfully
demonstrated on military aircraft components. However, the effects of
this methodology on structural design and certification requirements of
general aviation and commercial aircraft had not been examined. This
research initiates the examination for non-military components.

**Expected Results**

Upper wing skins are capable of tolerating damage induced under properly
controlled assembly procedures. However, poor assembly processes can
induce more severe damage in the structure, which may impose a threat to
the structural integrity. Therefore, assembly standards must be
established to minimize damage. Nondestructive inspection (NDI) after
final structural assembly should be performed if such standards do not
exist.

**Reference:** Kan, H.; Dyer, T.: Structural Integrity Evaluation of the
Lear Fan 2100 Aircraft, DOT/FAA/AR-95/13, May1996.

**Point of Contact:** P. Shyprykevich, AAR-431, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-4967, FAX: (609) 485-4004, e-mail:
pshyprykevich@admin.tc.faa.gov

## 

## The Effect of Preloading on Fatigue Damage in Composite Structures: Part 1

H. Hahn, J. Timmer, J. Barley-Cho, S. Lee, and S. Lim[^4]

**Research Objective**

Systematically separate the loading environment into component parts to
understand the interactions between each of the elements of the spectrum
on damage initiation, growth, and residual strength.

**Approach**

Two types of tension-tension fatigue tests were conducted on
quasi-isotropic laminates and damage was measured in the form of ply
cracks. Baseline fatigue tests were run at constant amplitudes ranging
from 20% to 60% of the ultimate tensile strength (UTS) while specimens
with preloads of 50% to 80% UTS were tested in fatigue at the same
amplitudes. The ply crack densities determined from edge replications in
preloaded specimens were compared to those without preload.

**Accomplishment Description**

The effect of a preload on subsequent fatigue damage growth depended on
the combination of both the preload level and the subsequent fatigue
stress level. Since the tension preload was always higher than the
tension-tension fatigue stress level, damage in preloaded specimens was
more severe in the low-cycle region than nonpreloaded specimens.
However, this preload induced damage did not grow any more if the
fatigue stress level was kept low. Preloads just above levels that cause
significant damage (70% UTS and above here) appeared to retard fatigue
damage development. At the highest tension-tension fatigue stress level
of 40% UTS, the high cycle damage in the form of matrix cracking
decreased with increasing preload level.

**Significance**

This report is the first of a series of reports that will provide
comprehensive documentation of damage induced by spectrum fatigue
loading in composite laminates. The effect of preloads on fatigue damage
in composite laminates has been examined in order to better understand
the modes of damage development and to assess the effect of proof
testing for airframe composite structures.

**Expected Results**

The results for this simplest form of variable loading will be used as a
building block with other data to eventually predict damage due to
fatigue spectrum loading for airframe composite structures.

**Reference:** Hahn, H.; Timmer, J., Bartley-Cho, J.; Lee, S.; Lim, S.:
The Effect of Preloading on Fatigue Damage in Composite Structures: Part
1, DOT/FAA/AR-95/79, April, 1996.

**Point of Contact:** P. Shyprykevich, AAR-431, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-4967, FAX: (609) 485-4004, e-mail:
pshyprykevich@admin.tc.faa.gov

## 

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## Comparative Evaluation of Failure Analysis Methods for Composite Laminates

C. T. Sun, B. J. Quinn, and J. Tao [^5], and D. W. Oplinger[^6]

**Research Objective**

To objectively study lamina and laminate failure criteria.

**Approach**

Comparisons among the commonly used failure criteria were made for
failure in unidirectional composites under various loading cases. From
these comparisons, the characteristics of these criteria were identified
and discussed.

**Accomplishment Description**

Over the last three decades, there have been continuous efforts in
developing failure criteria for unidirectional fiber composites and
their laminates. Currently, there exist a large number of lamina failure
criteria and laminate failure analysis methods. This effort furthered
the research on these subjects.

**Significance**

It was found that those lamina failure criteria which separate fiber and
matrix failure modes most accurately predict lamina and laminate
strength. Analytical methods that will accurately predict composite
laminate strength are needed for the design and certification of
composite aircraft structures.

**Expected Results**

With the aid of some limited experimental lamina and laminate strength
data available in the literature and new data generated by the authors,
failure criteria and laminate analysis methods that are mechanistically
sound and are capable of accurately predicting lamina and laminate
strengths for states of combined stresses will be developed.

**Reference:** Sun, C. T.; Quinn, B. J.; Tao, J; Oplinger, D. W.:
Comparative Evaluation of Failure Analysis Methods for Composite
Laminates, DOT/FAA/AR-95/109, May 1996.

**Point of Contact:** Donald Oplinger, AAR-431, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-4914, FAX: (609) 485-4569, e-mail:
Donald_Oplinger_at_ct27@admin.tc.faa.gov

# Crashworthiness 

## Survey and Analysis of Rotorcraft Flotation Systems

Mark Muller and Richard Greenwood[^7], Marvin Richards and Lindley
Bark[^8]

**Research Objective**

Evaluate rotorcraft flotation system performance in water-related
incidents and accidents and identify areas of potential improvement with
regard to rotorcraft flotation system operation and occupant fatality
reduction.

**Approach**

Four tasks were conducted: a survey of civilian and U.S. Navy rotorcraft
flotation technology, a review of civilian flotation accident and
incident data, definition of generic flotation system improvements, and
evaluation of available analytical methods for assessing flotation
system performance.

**Accomplishment Description**

After review of the performance data and state-of-the-art system
technology, four complexities associated with analytical modeling of the
ditching sequence were identified. These were buoyancy and stability,
wave action, water entry, and structural damage. Various existing
analytical methods were investigated for their treatment.

**Significance**

The design improvements are intended to address areas identified in the
performance data review and include use of auxiliary floats as well as a
refined arming and activation system for rotorcraft.

**Expected Results**

Generic design improvements to current flotation systems were suggested.
Updates to current FAA regulations were suggested.

**Reference:** Muller, M.; Greenwood, R.; Richards, M.; Bark, L.: Survey
and Analysis of Rotorcraft Flotation Systems, DOT/FAA/AR-95/53, May
1996.

**Point of Contact:** Gary Frings, AAR-431, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-4494, FAX: (609) 485-5785, e-mail:
Gary_Frings_at_ct27@admin.tc.faa.gov

## Transport Water Impact and Ditching Performance

Amit A. Patel and Richard Greenwood[^9]

**Research Objective**

Review and analyze worldwide transport accident data relative to water
impacts and ditching performance, compare the results of this study with
current FAA requirements to determine their adequacy and relevancy, and
conduct a survey of major worldwide airports to determine their
proximity to water.

**Approach**

The data were analyzed with respect to the airplanes' structural
integrity, breakup patterns, subsystem performance, cabin integrity, and
airline procedures that were or could be contributors to injuries and
fatalities. The current methods by which airframe manufacturers certify
their aircraft\'s ditching behavior were presented.

Because of the infrequency of unplanned water contacts and ditching
(planned) occurrences, a case study approach was taken in analyzing the
accident data. Eleven worldwide water impact accidents were identified
between the years 1959 and 1979. Of these, only one was classified as a
ditching occurrence. For the years 1980 to the present, three U.S.
water-related occurrences were identified. There were no ditching
occurrences identified during these years. All three occurrences
involved runway overruns.

**Accomplishment Description**

In deep water accidents, it was found that when the flight crew had at
least some degree of preparedness, trauma-caused injuries were minimized
while the majority of fatalities resulted from drowning. When the impact
was unexpected, however, the forces on the airplane were generally much
higher, resulting in a higher proportion of injuries and fatalities
caused by trauma. In shallow water incidents, usually occurring as a
result of runway overruns, drowning was not as common.

**Significance**

Injuries and fatalities in runway overruns are more likely to result
from excessive localized forces caused by the airplane's impact with
obstructions located in the area immediately beyond the end of the
runway. These localized forces were concentrated at the nose section of
the airplane and often led to fuselage breaks and separations.
Investigation of the incidents and understanding the force patterns may
help save passenger lives.

**Expected Results**

A survey of worldwide transport category airports was performed to
identify those airports located near significant bodies of water and to
analyze the operations at these airports. The airport database consisted
of 156 U.S. airports and 100 foreign airports. This information did not
exist prior to this research and will be incorporated into a Federal
Aviation Administration safety database.

**Reference:** Patel, A; Greenwood, R: Transport Water Impact and
Ditching Performance, DOT/FAA/AR-95/54, March 1996.

**Point of Contact:** Gary Frings, AAR-431, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-5781, FAX: (609) 485-5785, e-mail:
Gary_Frings_at_ct27@admin.tc.faa.gov

## 

## 

## Transport Water Impact, Part II

J. Tahliani and M. Muller[^10]

**Research Objective**

Document the over-water operating environment of jet transport aircraft
and is comprised of three sections: aircraft accident analysis, airport
water rescue, and emergency flotation devices.

**Approach**

Selected land accidents were placed in a hypothetical analysis to
predict their outcome had they impacted a body of water. A mathematical
model was used to capture the relative importance of various survival
factors in water impact accidents. Two impact scenarios were identified,
surface to ground (ditching) and runway overrun.

A survey of water rescue facilities, equipment, personnel, and
operations at twenty-three domestic airports was conducted.

The regulations pertaining to the design, certification, and use of
onboard emergency flotation devices including seat cushions, life
preservers, life rafts, and evacuation slides were studied. A technology
survey of commercially available devices was also conducted.

**Accomplishment Description**

The fatality rate for ditching accidents was 44 percent and for runway
overruns was 10.1 percent. Drowning was the major cause of death in
water impacts in both ditching and overrun scenarios. Postcrash fires,
deaths due to asphyxia, smoke inhalation, and thermal injuries were the
major postimpact hazards in land accidents in contrast to drowning from
entrapment, flotation device performance, and post evacuation exposure
in water impacts. Postimpact fatalities were almost two and a half time
higher in water accidents than in land accidents. However, there were a
greater number of impact fatalities in land accidents.

Larger airports were more likely to have on-site water rescue
capabilities. Among airports surveyed, 75 percent of large, 43 percent
of medium, and 25 percent of small airports had on-site water rescue
capability. Airports immediately adjacent to bodies of water were more
likely to have water rescue capabilities than those located within 5
miles of water. Airports that had provisions to keep rescue vessels
docked in the water had a significantly lower response time than
airports that do not have such provisions. The number, type, and
capabilities of water rescue vessels as well as the water rescue
equipment, personnel, and training varies greatly from airport to
airport. There are no regulations that require airports to operate and
maintain facilities, equipment, and personnel dedicated to water rescue
situations. No standardized requirements are in place for the number,
quantity, or type of water rescue vessels and equipment.

The hazard of hypothermia from water immersion is a major concern in
aircraft water impact. Emergency flotation devices are essential to
reducing this hazard. The use of technical standard order TSO C-13 life
preservers in place of the comparatively inferior TSO C-72 is desirable
in all aircraft. Improved designs should incorporate increased
protection from hypothermia. Although the performance of flotation seat
cushions has been debated, they are still highly essential on all
aircraft. In unplanned water accidents, they are likely to be the only
available means of flotation. The stowage location, retrievability, and
ease of unpacking and donning still remain the main causes of concern
for inflatable personal flotation devices (PFDs) and life preservers.
Basic regulatory amendments to improve field testing and demonstration
of PFDs were identified. Review of flotation equipment indicates that
raft stability, canopy design, packaging valise, and stowage location
may be improved.

**Significance**

Current federal regulations and National Transportation Safety Board
(NTSB) recommendations regarding airport water rescue were reviewed and
may be updated due to this research.

**Expected Results**

Recommendations on improving the existing regulations and incorporating
new features into the devices to improve performance and reliability of
the devices.

**Reference:** Tahliani, J.; Muller, M: Transport Water Impact, Part II,
DOT/FAA/AR-95/112, May 1996.

**Point of Contact:** Gary Frings, AAR-431, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-5781, FAX: (609) 485-5785, e-mail:
Gary_Frings_at_ct27@admin.tc.faa.gov

# Propulsion Systems and Fuels

## Spin Synchronous X-Ray Sinography for Nondestructive Imaging of Turbine Engines under Load

T. Kirchner, P. Burstein, and J. Youngberg[^11]

**Research Objective**

To develop and demonstrate a system for the early detection of
subcritical flaws in turbine engine components.

**Approach**

The Synchronous X-Ray Sinography (SXS) allows high resolution imaging
inspection of the interior of the rotating engine, particularly the
turbine disks and associated components, without engine disassembly. A
feasibility demonstration was conducted using a second stage turbine
section of a Lycoming T53 turboshaft engine.

**Accomplishment Description**

The second stage turbine engine was spun at 1800 rpm and a freeze-frame,
cross-sectional image was produced from the data. The SXS approach,
which is based on variations of computed tomography, demonstrated a
spatial resolution consistent with the detection of a 0.008-inch crack.
The first stage fan disk of a GE F101 engine was computer modeled to
show the effect of engine speed on crack opening. It was predicted that
a 0.3-inch long by 0.15-inch deep crack would open by 0.0015 inches
under load, an opening that would be detected by the SXS.

**Significance**

The SXS inspection can be conducted without engine disassembly,
therefore, allowing inspections to be done at much more frequent
intervals than tests that require dismantling the engine. The SXS could
also study regions of the engine that are far from the surface and that
are otherwise unobservable under load.

**Expected Results**

The demonstration showed the basic concepts of the SXS were feasible.
Such a system could detect cracks as small as 0.001 inches in turbine
engine components without disassembly of the engine.

**Reference:** Kirchner, T.; Burstein, P.; and Youngberg, J.: Spin
Synchronous X-Ray Sinography (SXS) for Nondestructive Imaging of Turbine
Engines Under Load, DT/FAA/AR-95/90, March 1996.

**Point of Contact:** Bruce Fenton, AAR-432, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-5158, FAX: (609) 485-4569, e-mail:
Bruce_Fenton_at_ct27@admin.tc.faa.gov

# Aging Aircraft

## Characterization of Early Stages of Corrosion Fatigue in Aircraft Skin

C. G. Schmidt, J. E. Crocker, J. H. Giovanola, C. H. Kanazawa, and D. A.
Shockey[^12]

**Research Objective**

To characterize the early stages of corrosion fatigue of the skin of
commercial aircraft fuselages.

**Approach**

An experimental technique was developed to produce corrosion fatigue
cracks in the laboratory in a manner that allowed examination of crack
nucleation and growth kinetics in both a statistical and deterministic
manner. A range of environments, material conditions, and loading
conditions were investigated experimentally to assess their effect on
the crack nucleation and growth rate.

**Accomplishment Description**

It was found that crack nucleation in bare material was slower than
crack nucleation in clad material. Crack nucleation from
crystallographic pits in the clad did not always occur at the largest
pit. For the factors considered within this study, it was found that in
aircraft skins, the crack nucleation mechanism that dominates depended
upon environment (e.g., pH) and the presence or absence of a clad layer.
The results suggest that corrosion fatigue in Aluminum clad 2024-T3
involves two competing mechanisms: hydrogen effects in the cladding
layer and pitting at constituent particles in the core alloy.

**Significance**

Corrosion fatigue in difficult-to-access areas of an aircraft fuselage
is an occasional problem in commercial aircraft and has the potential to
affect the structural integrity and the useful life of aircraft
structures. The research will address the synergistic effects of fatigue
and corrosion and quantitatively investigate the interaction of
corrosion and fatigue.

**Expected Results**

An improved deterministic understanding of the transition from corrosion
pit to short crack to long crack will result.

**Reference:** Schmidt, C. G.; Crocker, J. E.; Giovanola, J. H.;
Kanazawa, C. H.; and Shockey, D. A.: Characterization of Early Stages of
Corrosion Fatigue in Aircraft Skin, DOT/FAA/AR-95/108, February 1996.

**Point of Contact:** Dr. Thomas Flournoy, AAR-433, FAA William J.
Hughes Technical Center, Atlantic City International Airport, NJ 08405,
Phone: (609) 485-5327, FAX: (609) 485-4569, e-mail:
Thomas_Flournoy_at_ct27@admin.tc.faa.gov

## Antisymmetric Buffet Loads on Horizontal Stabilizers

Saeed Farokhi, Clay S. Mauk, and James E. Locke[^13]

**Research Objective**

To develop a method to model antisymmetric buffet design loads on
horizontal stabilizers with a known probability.

**Approach**

An experimental-computational method was developed using a rigid wind
tunnel model. The surface pressure was measured for a large number of
test conditions, including the most severe buffeting environment for
this type of aircraft: the tail immersed in the massively separated wake
of the wind. Modifications to a Beech King Air 200 one-sixth scale wind
tunnel model included the construction of a new horizontal stabilizer
instrumented with 12 miniature pressure transducers.

**Accomplishment Description**

Structural characteristics of the full-scale aircraft were estimated
using the Automated Structural Optimization System ASTROS program.
ASTROS, while not directly supporting buffet calculations, is written in
a flexible, high-level language and thus is easily adaptable.
Motion-dependent aerodynamics (stiffness and damping) were computed
using the proven doublet lattice method, which was incorporated into
ASTROS. The current method was validated using buffet pressure power
spectral densities from an existing reference. Based on the results, the
method was shown to be sound.

**Significance**

Under Federal Aviation Regulation 25-305(e), aircraft manufacturers are
required to demonstrate that the cumulative probability of an aircraft
encountering dangerous levels of buffet-induced rolling moment is below
the prescribed level. The current accepted method of meeting this
requirement involves a great deal of full-scale flight testing, which
can be very expensive. The new method would allow the design rolling
moment load to be estimated before the full-scale aircraft is
constructed. A standardized method would expedite the certification
process and produce consistent and repeatable results.

**Expected Results**

A method was developed to predict, with a known probability, the
antisymmetric buffeting of horizontal stabilizers in massively separated
flows for use in the design and certification of aircraft.

**Reference:** Farokhi, S.; Mauk, C. S.; and Locke, J. E.: Stochastic
Modeling of Antisymmetric Buffet Loads on Horizontal Stabilizers in
Massively Separated Flows, DOT/FAA/AR-95/7, March 1996.

**Point of Contact:** Thomas DeFiore, AAR-431, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-5009, FAX: (609) 485-4569, e-mail:
Thomas_DeFiore_at_ct27@admin.tc.faa.gov

## An Analysis of Ground-Flight Loads Measured on a B-727

William M. Cavage[^14], Thomas DeFiore[^15], and Terrence Barnes[^16]

**Research Objective**

To analyze some recent, high quality, independent data with high sample
rates to measure and correlate landing gear internal and external loads
experienced during takeoff, landing, and some selected abrupt ground
maneuvers.

**Approach**

The study used existing high speed B-727 data to compile the flight
acceleration, aircraft system, and force data for each maneuver into
data files. The flight information was analyzed to provide a typical
picture of each aircraft maneuver. Relationships describing takeoff,
landing, exits, braking, S-turn, and minimum radius turn aircraft
maneuvers were examined.

**Accomplishment Description**

Multiple time trace data plots of all maneuvers illustrated variable
relationships as expected. Pitch rate was calculated to be 3.43 degrees
per second, while the maximum was 4.63 degrees per second. Histograms
were plotted and mean, median, and standard deviation were calculated
for all takeoff and landing critical parameters although the small
number of samples caused a large amount of uncertainty. Measured values
were compared with calculated ones; there was good agreement between the
two although a great deal of scatter was found for the measured and
calculated braking drag shear. Measured accelerations agreed well wit
measured forces. Relationships between braking acceleration and drag
shear, lateral acceleration and side shear, and normal acceleration and
vertical shear were determined for the braking, exiting, and landing
tests, respectively. A linear relationship was found between side shear
and axle differential load.

**Significance**

These relationships provide some directly measured internal loads
information not available in previous research efforts.

**Expected Results**

The data will provide an understanding of the relationship between the
external conditions (e.g., runway exit speed) and the internal loads
(e.g., individual main gear loads) which may be of value in
understanding service problems on older airframes and will assist in the
development of the design loads and operating loads on new designs.

**Reference:** Cavage, W. M.; DeFiore, T.; Barnes, T.: An Analysis of
Ground-Flight Loads Measured on the Instrumented B-727 N40,
DOT/FAA/AR-95/82, October 1995.

**Point of Contact:** Thomas Defiore, AAR-431, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-5009, FAX: (609) 485-4569, e-mail:
Thomas_DeFiore_at_ct27@admin.tc.faa.gov

## Incorporating the Dripless Bubbler with a Scanner for Ultrasonic C-Scan Imaging of Defects in Bonded Aircraft Structures

D. K. Hsu and D. J. Barnard[^17]

**Research Objective**

To develop a focused beam, water coupled scanning method to detect
corrosion and disbonds in adhesively bonded structures and riveted lap
splices in an airline hangar environment.

**Approach**

A second generation Dripless Bubbler has been designed for use with a
commercial scanner, with improvements made to ensure a more compact and
portable system. Laboratory and field trials have been used to evaluate
design changes aimed at enhancing durability, operator ease of use, and
data interpretation.

**Accomplishment Description**

The new Dripless Bubbler design has been adapted to a Tektrend Panda
II^®^ scanner that is both light (8 lb.) and capable of conforming to
fuselage curvatures. A new couplant (water) handling system has been
developed that uses the compressed air readily available in a hangar
environment to power a diaphragm pump and venturi vacuum. This unit is
compact, light weight (7 lb.) , and operates on only 1-1.5 gallons of
water. Significant improvements to the durability and scanning action of
the Dripless Bubbler are due to the brush cup seals incorporated into
this second generation design. A recent change from nylon to
polypropylene brushes has made for increased wet bristle stiffness which
aides in maintaining a constant stand-off distance for the focused
immersion probes used. Tests on sample corrosion panels have shown that
c-scan images produced by this system offer the resolution and
sensitivity of immersion scans. The system has to date been on the
following aircraft: Beechcraft King Air at Iowa State University Flight
Service, Boeing 737 at Federal Aviation Administration Aging Aircraft
Nondestructive Inspection Validation Center (in Albuquerque, NM),
DC-9-30 at Midwest Express Airlines (Milwaukee, WI) and Boeing 747-200
at Northwest Airlines (Minneapolis, MN).

**Significance**

The Dripless Bubbler is a patented technology, licensed to a commercial
nondestructive testing equipment manufacturer. To date, more than 60
inquiries from industry, government, and academia have been received.
This device provides a durable, easy to use method to detect corrosion
and disbonds in aircraft structures.

**Expected Results**

A functioning prototype for conducting beta-site testing at the airlines
and for evaluation by airline inspection staff.

**Reference**: Barnard, D.; Hsu, D.: "Development and Testing of the
Dripless Bubbler Ultrasonic Scanner." Review of Progress in Quantitative
Nondestructive Evaluation, Vol. 16, presented in Brunswick, ME, July
28-August 2, 1996.

**Point of Contact:** David Galella, AAR-433, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, (609)
485-5784, FAX (609) 485-4005, e-mail: galellad@admin.tc.faa.gov

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## An Overview of the Engine Titanium Consortium

Bruce Thompson and Lisa Brasche[^18], Kevin Smith[^19], and Jon
Bartos[^20]

**Research Objective**

The Engine Titanium Consortium (ETC) was established to provide
reliable, cost-effective methods for the inspection of engine materials,
components, and hardware. A university and industry consortium, the ETC
brings together the fundamental support of university research with the
engineering talents of industry. Iowa State University, General
Electric, Pratt & Whitney, and Allied Signal have established a strong
team to develop and transfer inspection technology throughout the engine
life cycle.

**Approach**

The ETC accomplishes its objective in performance of four primary tasks:
ultrasonic inspection in production; fundamental studies in titanium;
probability of detection; and eddy current inspection inservice.

**Significance**

The Engine Titanium Consortium has brought together fundamental
understanding of university research with the engineering talents of
industry to address generic, technology base issues related to
inspection of engines. The ETC works closely with the engine lifing
community, the airlines, and the FAA New England Engine and Propeller
Directorate to ensure that the highest priority issues are addressed.
Because the end users have been involved in the planning,
prioritization, and development, implementation is facilitated.

**Expected Results**

Comprehensive, cost effective tools for inspection throughout the life
cycle of titanium.

**References:** Brasche, L: "Improving Safety of Aircraft Engines: A
Consortium Approach" Proceedings of the 1996 SPIE Conference,
Nondestructive Evaluation of Aging Infrastructure.

**Point of Contact:** Dr. Christopher Smith, AAR-433, FAA William J.
Hughes Technical Center, Atlantic City International Airport, NJ 08405,
609-485-5221, FAX, 609-485-4569, e-mail: smithc@admin.tc.faa.gov

## An Overview of the Contaminated Billet Study

Lisa Brasche[^21], Kevin Smith[^22], and Jon Bartos[^23]

**Research Objective**

To use a contaminated heat of Ti-6Al-4V billet to better understand the
effect of defects, namely hard alpha, and inspectability on engine
materials and components.

**Approach**

A 6300 pound heat of Ti-6Al-4V billet is being used by the Engine
Titanium Consortium (ETC) and the Turbine Rotor Materials Design
programs to assess the inspectability of hard alpha and the
morphological changes that occur with processing from billet to final
forging shape. A comprehensive program that includes destructive
characterization of 10 indications removed from the billet and 10
indications removed at various points in further processing to final
forged component is planned. As part of the development of the
probability of detection (POD) for billet inspection, additional studies
will be completed to assess if misses occurred.

**Accomplishment Description**

The heat has been inspected using multizone and conventional inspection
(5 MHz longitudinal focus and 5 MHz refracted longitudinal) transducers
at the RMI titanium facility in Niles, Ohio. Six of the 12 billets were
also inspected using a phased array inspection system at RMI. Analysis
of the results led to 64 indications detected with the multizone
inspection technique compared to 31 indications with the conventional
technique. Comparison of the two zoned inspection approaches, phased
array and multizone, indicated that the same indications were detected.
Destructive analysis of select indications is under way to determine if
indications were hard alpha or whether there were false calls.

**Significance**

The contaminated billet study provides a unique opportunity to
understand the effect of morphology and location of hard alpha defects
on inspectability. Portions of the heat are also being used to ascertain
the movement of defects during processing from billet to pancake forging
to final forging shape. Verification of deformation models is under way
as part of the FAA-funded Turbine Rotor Materials Design Program.

**Expected Results**

Comprehensive understanding of naturally occurring hard alpha defects
available in the public domain.

**References:** Brasche, L.; Smith, K.; Bartos, J.: "Overview of the
Contaminated Billet Study," Review of Progress in Quantitative
Nondestructive Evaluation, 1996.

**Point of Contact:** Dr. Christopher Smith, AAR-433, FAA William J.
Hughes Technical Center, Atlantic City International Airport, NJ 08405,
609-485-5221, FAX, 609-485-4569, e-mail: smithc@admin.tc.faa.gov

## An Element-Free Galerkin Method for Three-Dimensional Crack Growth Modeling

B. Moran, T. Belytschko, N. Sukumar, and T. Black[^24]

**Research Objective**

To explore the application of the Element-Free Galerkin (EFG) method to
three-dimensional linear elastic fracture mechanics, with particular
emphasis on stress-intensity factor (SIF) determination for stationary
as well as growing cracks.

**Approach**

The Element-Free Galerkin method requires only nodal data - no element
connectivity is required to describe the domain. The shape functions are
constructed using moving least square interpolation (MLS). Essential
boundary conditions are satisfied by choosing finite element nodes on
the boundaries where displacement is prescribed. In order to adapt the
method for crack problems, enriched basis functions, which include the
terms in the asymptotic expansion of the near-tip displacement field,
are used. A Galerkin procedure on the weak form results in the discrete
system of equations: Kd = f. A sparse iterative solver is used to solve
the linear system of equations. Domain integral methods are used to
evaluate the stress intensity factor along the three-dimensional (3D)
crack front. The EFG method allows for higher order continuity in the
displacements which facilitates the use of the surface as well as volume
forms of the domain integral.

**Accomplishment Description**

Two benchmark three-dimensional crack problems were considered in order
to assess the accuracy of the method and to establish settings for EFG
parameters. Firstly, a 3D edge-crack specimen under tension was
considered. The stress intensity factors obtained by, both, the volume
and surface forms of the domain integral were found to be in good
agreement with reported values in the literature. Secondly, the problem
of a penny-shaped crack in an infinite domain was considered (Figure
1a). The nodal discretization on the crack plane is shown in Figure 1b.
The SIF values obtained by the volume form of the domain integral were
compared to the exact solution (Figure 1c). Excellent agreement between
the numerical and exact solution was obtained.

**Significance**

Provides guidance for service life prediction of aging aircraft.

**Expected Results**

EFG method for crack growth modeling of planar as well as non-planar
crack fronts.

**Reference:** Sukumar, N.; Moran, B.; Black, T.; Belytschko, T.: An
Element-Free Galerkin Method for Three-Dimensional Fracture Mechanics.
Computational Mechanics, accepted for publication.

**Point of Contact:** Dr. Christopher Smith, AAR-433, FAA William J.
Hughes Technical Center, Atlantic City International Airport, NJ 08405,
609-485-5221, FAX, 609-485-4569, e-mail: smithc@admin.tc.faa.gov

![](media/image3.emf){width="8.164583333333333in"
height="4.892361111111111in"}

## The Effects of Beam Distortion On The Strength Of Ultrasonic Signals

P. D. Panetta, F. J. Margetan, I. Yalda, and R. B. Thompson[^25]

**Research Objective**

To gain an understanding of a variety of unusual, microstructurally
controlled, wave propagation phenomena in titanium alloys which cause
flaw signals to fluctuate, decreasing the reliability of flaw detection,
and back surface signals to fluctuate, adding errors to attenuation
corrections.

**Approach**

As a first step, the simple problem of fluctuations in back surface
signals, as occurs in measurements of attenuation, has been considered.
This problem is of interest because (a) the back surface signal can be
considered as an approximation to the signal from a very large flaw and
(b) these signals are used in titanium inspection as part of a
correction for sample specific attenuation. Fluctuations in these
signals render both the attenuation correction and flaw detection less
reliable than would be desired. To understand the origins of these
fluctuations, the results of classical attenuation measurements, based
on the amplitude of back surface echoes, have been compared to
point-probe measurements of attenuation, in which the energy transmitted
through a sample is directly measured, independent of phase, with a very
small probe.

**Accomplishment Description**

The results show that the attenuation measured in the classical
technique can be much higher than the attenuation inferred from the
point-probe measurements of transmitted energy. These are explained in
terms of significant distortions of the phase and amplitude of the beam
when it propagates through titanium, particularly in certain directions,
as compared to the behavior in homogeneous materials such as fused
quartz. The existence of these distortions will obviously influence the
response of small flaws in ways, which will be more fully determined in
future work.

**Significance**

The understanding of the wave propagation phenomena will lead to less
distortions and more accureate measurement of small flaws using
ultrasonic methods.

**Expected Results**

New more accurate techniques for inspection of aircraft engine
components.

**References:** Panetta, P. D.; Margetan, F. J.; Yalda, I.; Thompson, R.
B.: "Observation and Interpretation of Microstructurally Induced
Fluctuations of Back-Surface Signals and Ultrasonic Attenuation in
Titanium Alloys," Presented at the Twenty First Review of Progress in
Quantitative Nondestructive Evaluation, Brunswick, Maine, July 28-August
2, 1996.

Panetta, P. D.; Thompson, R. B.; Margetan, F. J.; Yalda, I.: "Ultrasonic
Microstructural Characterization and Modeling of Attenuation in Ti-17
and Ti-6Al-4V Alloys," Presented at American Society of Materials -
Titanium Materials Society Materials Week '96, Cincinnati, Ohio, October
7-10, 1996.

**Point of Contact:** Dr. Christopher Smith, AAR-433, FAA William J.
Hughes Technical Center, Atlantic City International Airport, NJ 08405,
609-485-5221, FAX, 609-485-4569, e-mail: smithc@admin.tc.faa.gov

## Transducer Design for High Sensitivity Ultrasonic Inspection of Titanium

P. J. Howard, D. C. Copley, and R. S. Gilmore[^26]

**Research Objective**

Investigate transducer design parameters such as frequency and beam
diameter to determine their affect on inspection sensitivity in titanium
(Ti) alloys. Develop a methodology for designing transducers to produce
a high sensitivity C-scan image based inspection in Ti-6Al-4V (Ti-6-4)
and Ti-Al5-Sn2-Zr2-Mo4-Cr4 (Ti-17) parts with flat sound entry surfaces.

**Approach**

A set of transducers spanning a wide range of transducer design
parameters was acquired from a single commercial manufacturer. These
transducers were used to collect C-scan images of small (1/64")
flat-bottom holes, tungsten-carbide inclusions, uncracked, unvoided
synthetic Ti-N (hard-alpha) inclusions, and grain noise data from eight
separate Ti alloy sample blocks. The images were analyzed statistically
and the results compared to measured transducer parameters.

**Accomplishment Description**

Each of the transducers was characterized experimentally by measuring
the frequency, half-amplitude beam diameter and pulse length. These
parameters were used to estimate the volume of the ultrasonic pulse. The
amplitude and signal-to-noise ratio (SNR) of each target and the maximum
amplitude of Ti grain noise images were extracted and plotted versus the
measured parameters. Plots of the target SNR vs. measured design
parameters are shown in the accompanying figure. These plots demonstrate
that the SNR from flat bottom holes and inclusions increased as the
volume of the ultrasonic pulse decreased.

**Significance**

Results demonstrate a method for increasing the sensitivity of
ultrasonic C-scan image based inspections of Ti alloys based on using
transducers with the smallest pulse volume practicable.

**Expected Results**

A methodology, validated by experiments, for selecting a transducer to
produce a given sensitivity for an ultrasonic nondestructive test.

**Reference:** Howard, P. J. and Gilmore R. S.: "Transducer Design for
High Sensitivity Ultrasonic Inspection of Titanium," Review of Progress
in Quantitative Nondestructive Evaluation*,* Vol. 16, (Plenum
Publishing, New York), June 1997.

**Point of Contact:** Dr. Christopher Smith, AAR-433, FAA William J.
Hughes Technical Center, Atlantic City International Airport, NJ 08405,
609-485-5221, FAX, 609-485-4569, e-mail: smithc@admin.tc.faa.gov

![](media/image4.emf){width="9.139583333333333in"
height="5.152083333333334in"}

## Eddy Current Probe Design for Edge Effect Reduction

S. Sharma, I. Elshafie, L. Udpa[^27]

**Research Objective**

To investigate alternate eddy current probe designs for reducing edge
effects in multilayer geometries.

**Approach**

A finite element (FE) model was generated for simulating multilayer
aircraft geometries and predicting the flux distribution and impedance
plane trajectories of various probe configurations. The probe
configurations were selected to reduce edge signal by shaping the flux
of the probe coil using auxiliary coils with appropriate currents.

**Accomplishment Description**

The two dimensional finite element model was used to visualize the
magnetic flux distribution of various probe geometries. Model results
were used to design a probe with a concentric flux shaping auxiliary
coil. The magnitude and phase of the current in the auxiliary coil was
optimized to reduce the spreading of flux inside the conductor and
thereby reduce the edge contribution to the signal. The probe design was
also tested using three dimensional FE model to predict the impedance
plane trajectory due to a crack in the vicinity of an edge. The model
results validate the feasibility of the flux shaping approach for edge
effect reduction. A prototype of the probe is currently being built for
experimental validation.

**Significance**

By reducing edge effects in multilayer geometries, eddy current methods
can more accurately characterize flaws in aircraft structures.

**Expected Results**

A candidate eddy current probe design for detecting defects close to
edges in multilayer geometries.

**Reference:** Sharma, S.; Elshafie, I.; Udpa, L.: "Finite Element
Modeling of Eddy Current Probes for Edge Effect Reduction," Presented at
the 21st Review of Progress in Quantitative Nondestructive Evaluation,
Brunswick, Maine, July 1996.

Sharma, S.; Elshafie, I.; Udpa, L: "Probe Design for Edge Effect
Reduction in Eddy Current Inspection," Presented at the SPIE Conference
on Nondestructive Evaluation of Aging Infrastructure, Scottsdale,
Arizona, December 1996.

**Point of Contact:** Dr. Christopher Smith, AAR-433, FAA William J.
Hughes Technical Center, Atlantic City International Airport, NJ 08405,
609-485-5221, FAX, 609-485-4569, e-mail: smithc@admin.tc.faa.gov

![](media/image5.emf){width="7.729166666666667in"
height="5.881944444444445in"}

## Air-Coupled C-Scan of Boeing Repair Coupon

D. E. Chimenti and O. I. Lobkis[^28]

**Research Objective**

To adapt, develop, and improve air-coupled ultrasonic inspection methods
for in-service defect detection and aircraft manufacturing inspections
of composite components.

**Approach**

Assess resolution and sensitivity limits of air-coupled ultrasonics and
defect detection versus several measurement parameters. Emphasize
composites and subsurface impact damage, internal delaminations, lack of
substrate adhesion, lay-up ply orientation defects, and material
property variations will be studied. Explore feasibility of development
of field-implementable inspection devices using air-coupled technology.

**Accomplishment Description**

Transmission air-coupled C-scans of a Boeing composite sample have been
performed and compared to conventional scans. The B-777 specimen was cut
from the tail section and repaired, in a dished, scarf-joint circular
area using Boeing methods and included several simulated defects; the
repair was complicated by a tapered bond zone. This sample was scanned
at 0.5 MHz with a higher-order plate mode in an optimized transmission
C-scan with the receiver placed to observe the directly transmitted
signal, monitored as a function of transducer position. Despite the
structural complexity, the air-coupled inspection yields an excellent
map of the defective area, showing all significant internal features.
The repaired area is a large circular feature transected by diagonal
lines of the prepreg tape borders. The inner circle shows the bond zone
of the second of two tapered repair inserts. Within this inner circle
are two clear areas (light gray scale) showing damage in the repaired
zone. These correlate almost perfectly with the Boeing sample
fabrication drawings and with water-coupled scans.

**Significance**

The feasibility of the air-coupled ultrasonic method for disbonds in
realistic aircraft structures.

**Expected Results**

Development of methods and procedures for air-coupled inspection and
materials characterization; system specifications for practical
air-coupled ultrasonic inspection system.

**References:** Lobkis, O; Chimenti, D: "Anisotropic Materials
Characterization Using Air-coupled Ultrasound," Review of Progress in
Quantitative Nondestructive Evaluation, Vol. 16, Eds. D. O. Thompson and
D. E. Chimenti (Plenum Press, New York), 1997.

**Point of Contact**: David Galella, AAR-433, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405 (609)
485-5784, FAX (609) 485-4005, e-mail: galellad@admin.tc.faa.gov

## Modeling the Relationship between Microstructure and Backscattered Noise

Y. K. Han and R. B. Thompson[^29]

**Research Objective**

To gain a fundamental understanding of the relationship between
microstructure and backscattered ultrasonic noise, in particular the
dependence of noise on frequency, propagation direction, and
microstructure, and to use this to provide feedback to the billet
producers regarding microstructures which would have low noise
characteristics and thus improved inspectability.

**Approach**

The approach builds on previous work of in which backscattered noise was
predicted based on the Born approximation. For single-phase, equiaxed
polycrystals, his predictions of noise as a function of grain size and
frequency were found to be good agreement with experiment. In the new
work, this technique has been applied to a more complex microstructure,
typical of titanium alloys, consisting of macrograins containing
colonies with crystallographically related orientations.

**Accomplishment Description**

Results have been obtained showing how the backscattering is controlled
by the relative values of frequency, macrograin size and shape, and
colony size and shape. At sufficiently long wavelengths, the
backscattering is found to be isotropic. At wavelengths approaching the
colony size, the anisotropy of the backscattering is controlled by the
colony elongation. At intermediate wavelengths, the anisotropy of the
backscattering is a complex function of the elongation of the
macrograins and colonies. These results represent a significant step in
understanding the relative role of texture and microstructure size and
shape in determining the noise. Although numerical results have only
been developed for certain idealized microstructures and
macrostructures, the tools used in the computations have sufficient
generality to be applied to the microstructures developed in real engine
alloys.

**Significance**

Provides technical information to modify processing for reducing
ultrasonic noise in the inspection of aircraft engine components.

**Expected Results**

New nondestructive evaluation techniques for aircraft.

**Reference:** Han, Y. K. and Thompson, R. B.: "Ultrasonic
backscattering in duplex microstructures: theory and application to
titanium alloys," Metallurgical and Materials Transactions. To be
published in January, 1997.

**Point of Contact:** Dr**.** Christopher Smith, AAR-433, FAA William J.
Hughes Technical Center, Atlantic City International Airport, NJ 08405,
609-485-5221, FAX, 609-485-4569, e-mail:
smithc@admin.tc.faa.gov![](media/image6.emf){width="7.71875in"
height="5.479166666666667in"}

## Ultrasonic Microscopy of Synthetic Spherical TiN Inclusions in Forged Ti-6Al-4V

J. B. Deaton, Jr., M. F. X. Gigliotti, L. C. Perocchi, R. S. Gilmore,
and E. J. Nieters[^30]

**Research Objective**

Fabricate Ti-6-4 test standards containing embedded 4/64" (#4) spherical
Ti-2.9 wt. % N inclusions, and evaluate the ultrasonic detectability of
these #4 spherical Ti-2.9 wt. % N inclusions at a 1" depth in the Ti-6-4
standards.

**Approach**

Synthetic Ti-2.9 without N inclusions were prepared by arc-melting and
bonding Ti sponge and TiN powder. Spherical seeds were machined to shape
from the TiN ingot and bonded between two Ti-6-4 forged blocks to create
the final test standard. Ultrasonic C-scan images were collected of both
spherical and flat-faced inclusions at a depth of 1" beneath the ground
surfaces of the respective Ti-6-4 test standards. Images were acquired
using 5 and 10 MHz f/4 and 5, 10, and 20 MHz f/7 focused beam immersion
transducers.

**Accomplishment Description**

Ti-6-4 test standards were produced with embedded #4 Ti-2.9 wt. % N
inclusions that were uncracked and unvoided and which remained nominally
spherical after the HIP bonding. The signal-to-noise ratio (SNR) was
calculated for the various inclusion geometries and transducer
parameters to estimate detectability of the spherical inclusions as
compared with the detectability of flat-faced inclusions of similar
nitrogen content, as shown in the attached figure.

**Significance**

The highest SNR for all inclusion geometries was achieved with a 20 MHz
f/7 transducer. The higher material noise levels near the diffusion bond
line must be corrected for when predicting the detectability of these
synthetic TiN inclusions.

**Expected Results**

Estimates of detectability of naturally occurring TiN inclusions of
varying geometries in Ti-6-4.

**References:** Deaton, J; Gigliotti, M; Perocchi, L; Gilmore, R;
Nieters, E: "Ultrasonic Microscopy of Synthetic Spherical TiN Inclusions
in Forged Ti-6Al-4V," presented at the 23^rd^ Annual Review of Progress
in Quantitative Nondestructive Evaluation, Brunswick, ME, August 1996.

**Point of Contact:** Dr. Christopher Smith, AAR-433, FAA William J.
Hughes Technical Center, Atlantic City International Airport, NJ 08405,
609-485-5221, FAX, 609-485-4569, e-mail: smithc@admin.tc.faa.gov

![](media/image7.emf){width="7.877083333333333in"
height="5.577083333333333in"}

## Development of Ultrasonic Models to Guide the Design and Evaluation of Inspection Systems for Titanium Rotating Components

R. B. Thompson, F. J. Margetan, I. Yalda, and C. P. Chiou[^31]

**Research Objective**

To develop an integrated set of models to guide the design and
evaluation of inspection systems for titanium rotating components. These
should include the effects of microstructure on both flaw response and
competing noise.

**Approach**

The complex microstructure of titanium alloys influences the ultrasonic
detection of flaws in a variety of ways. Backscattered noise competes
with flaw signals and material inhomogeneties distort the beam, leading
to fluctuations in flaw signals and potential errors in standard
corrections for attenuation. These effects must be taken into account in
designing inspection systems and evaluating their performance. Models,
based on the scattering of ultrasound from hard-alpha inclusions and
microstructure are being developed to simulate inspections and thereby
guide these tasks. These models are all based on the same formal
framework, Auld's electromechanical reciprocity relation, and use
various approximations to predict signal and noise under different
conditions.

**Accomplishment Description**

Three classes of models are in varying states of development. Most
complete are models for backscattered noise, which can predict root mean
square (RMS) noise and gated peak-to-peak noise distributions from
knowledge of the details of the measurement system and the noise
generating capacity of the materials (described by factors that can be
determined experimentally or predicted from knowledge of the
microstructure). Additional models are under development to predict the
signals reflected from naturally occurring hard-alpha inclusions,
including the independent contributions of pores, cracks and the
diffusion zone, and how these flaw signals are modified by
microstructural inhomogeneties. These models are finding application in
assessing probabilities of detection and selecting probes for optimum
flaw detectability.

**Significance**

Models which simulate inspections will guide the design and evaluation
of inspection systems without expensive fabrication costs.

**Expected Results**

Inspection systems for titanium rotating components.

**Reference:** Thompson, R. B.: "Use of UT Models in Titanium Inspection
Development," Presented at the Twenty First Review of Progress in
Quantitative Nondestructive Evaluation, Brunswick, Maine, July 28-August
2, 1996.

**Point of Contact:** Dr. Christopher Smith, AAR-433, FAA William J.
Hughes Technical Center, Atlantic City International Airport, NJ 08405,
609-485-5221, FAX: 609-485-4569, e-mail: smithc@admin.tc.faa.gov

## Modeling Ultrasonic Inspectability in Forgings

T. A. Gray[^32]

**Research Objective**

To use ultrasonic measurement modeling capabilities to define
inspectability of small defects in near net-shaped forgings, such as are
used in the manufacture of jet engine turbine disks.

**Approach**

Ultrasonic inspectability, defined as signal amplitude or as percent
screen height relative to a distance-amplitude-curve (DAC), is simulated
using the Gaussian-Hermite ultrasonic beam model and the Kirchoff
approximation for elastic wave scattering from planar or spheroidal
scatterers within the volume of a forging. A geometrical description of
the forging, including surface curvatures, normal vectors, etc., is
obtained from computer-aided design (CAD) representations of the
component. Signal amplitude maps for assumed ultrasonic scans of the
forging are simulated and the results are converted either to raw binary
image files or to Microsoft Windows bitmap files.

**Accomplishment Description**

Pratt & Whitney Aircraft supplied a CAD file containing the geometry of
a titanium forging. The inspectability simulation software was used to
compute #1 (1/64 inch diameter) flat-bottomed hole (FBH) signal
amplitudes as a function of position within the forging cross-section
for an assumed set of scanned parameters. Results were computed both
with and without an assumed DAC compensation.

**Significance**

Visualization technique will enhance usability of these techniques.

**Expected Results**

A method to predict and visualize internal defects in forgings using
ultrasonic techniques.

**Reference:** Gray, T. A.: "Application of Measurement Models to
Specification of Ultrasonic Inspections," Review of Progress in
Quantitative Nondestructive Evaluation, vol. 16*,* D. O. Thompson and D.
E. Chimenti, eds., Plenum Press, New York.

**Point of Contact:** Dr. Christopher Smith, AAR-433, FAA William J.
Hughes Technical Center, Atlantic City International Airport, NJ 08405,
609-485-5221, FAX, 609-485-4569, e-mail: smithc@admin.tc.faa.gov

## Improved Methodology for Predicting Probability of Detection (POD) of Synthetic Hard Alpha Inclusions in Titanium

William Q. Meeker, Shuen-Lin Jeng, Chien-Ping Chiou, R. Bruce
Thompson[^33]

**Research Objective**

Develop methodology to estimate probability of detection (POD) for
detecting synthetic hard alpha inclusions in titanium. The methodology
will be used to predict POD, probability of false alarm (PFA) and
receiver operating characteristic (ROC) function curves.

**Approach**

The methodology is based on estimating distributions of signals from
flaws in noise and of the distribution noise alone. The methodology uses
combinations of physical modeling of an inspection process (based on the
theory of ultrasonic wave scattering), along with laboratory and
production data, to estimate nondestructive evaluation (NDE) capability.
A physical prediction model is used to predict the effect of making
specific changes to the inspection setup (e.g., probe characteristics).
Using the model-specified signal distribution (e.g., signal strength as
a function of distance of the flaw from the beam center), probability
theory is used to describe variabilities that can be modeled
mathematically (e.g., the random position of a flaw relative to scan
lines and, correspondingly, the effect of changing scan increment and
gate width). A statistical prediction model is used to quantify that
part of the scatter observed in actual inspection data that is not
accounted for by the physical model.

**Accomplishment Description**

We have used available experimental data to assess the adequacy of the
model predictions and to characterize the unexplained variability. We
used a statistical model for the generalized deviations between the
actual data and the physical model. These deviations are for a
parametric distribution in the Box-Cox family of distributions.
Interestingly, the Box-Cox transformation parameter seems to be
remarkably constant (at about .3) over different experimental
situations. This distribution, along with model signal predictions, can
be used to generate POD curves. For a given noise distribution, ROC
curves can also be generated.

**Significance**

This research provides an important step in the development of the
Engine Titanium Consortium POD methodology for titanium billet
inspection. The next step is to combine the results of this study with
current research on signal response from real flaws obtained in the
Contaminated Billet Study.

**Expected Results**

A methodology to predict POD, PFA, and ROC of titanium inspection
techniques.

**Reference:** Meeker, W. Q.; Jeng, S. L.; Chiou, C. P.; Thompson, R.
B.: "Improved Methodology for Predicting POD of Detecting Synthetic Hard
Alpha Inclusions in Titanium," Review of Progress in Quantitative
Nondestructive Evaluation 16, Plenum Press.

**Point of Contact:** Dr. Christopher Smith, AAR-433, FAA William J.
Hughes Technical Center, Atlantic City International Airport, NJ 08405,
609-485-5221, FAX, 609-485-4569, e-mail: smithc@admin.tc.faa.gov

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height="3.9090277777777778in"}
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height="3.892361111111111in"}

## Hand-Held, Flexible Eddy Current Probe for Inspection of Curving Surfaces

Thadd C. Patton, Robert Filkins, James Fulton, John Young, Kristina
Hedengren, Carl Granger, and Tom Hewton[^34]

**Research Objectives**

Develop single coil derivatives of the Eddy Current Array Probe (ECAP)
for manual and automatic scanning applications that can be used with
conventional, single-channel, eddy scope instruments, such as the Nortec
19. The new single element eddy current sensors, designated (SECAP) were
designed to cover an area previously covered by an array of sensors.

**Approach**

The single element eddy current sensors, are similar to the original
ECAP sensors and have the following properties: (1) multiple eddy
current coil elements are replaced by a single larger eddy current coil
element; (2) the layout and design of the drive and sense coils are
optimized for single coil detection sensitivity; (3) preamplification
and signal conditioning are provided to the differential sense output to
improve signal-to-noise ratio; and (4) the shape of the flexible circuit
is modified to promote bending flexibility over curving surfaces. The
new sensor elements were incorporated into a hand-held probe with
internal preamplification; and used to demonstrate wide area inspection
coverage by hand scanning.

**Accomplishment Description**

The differential, reflection style, SECAP sensor elements were
incorporated into a hand-held probe with internal preamplification to
demonstrate wide area hand scanning. The probe housing is built around
the preamplification circuitry; and serves to shield and protect the
internal components of the probe while providing an ergodynamic hand
grip for manual positioning of the probe. The sensor elements of the
SECAP probe are interchangeable within the same probe holder, thus
providing adaptation to a wide range of inspection conditions. The probe
as shown was demonstrated at the 1996 Airline Transportation Association
Nondestructive Testing Forum, held in Seattle, Washington.

**Significance**

Results demonstrate the feasibility of detecting 0.030" fatigue cracks
in titanium using a single large detector, whereby reducing part
inspection times by greater than four times. Upon completion, a family
of wide-area scanning probes for the inspection of smoothly curving
engine components will be available for commercialization.

**Expected Results**

Presently, the SECAP probe is a developmental prototype designed for the
inspection of generally smooth, curving engine surfaces. Future versions
of the prototype will be designed for tight convex and concave surfaces
as well as for radius corner geometries. Also of consideration are a
reduction in the size of the probe body and preamplification board, the
design optimization of the SECAP coil elements, as well as an
improvement in its ergodynamic design and ruggedness. The general
surface SECAP probe is currently planned for beta-site testing in 1997.

**References:** Hurley, D; Hedengren, K; Howard, P; Kornrumpf, W;
Sutton, G;Young, J: "An Eddy Current System for Aircraft Engine
Inspection," Review of Progress in Quantitative Nondestructive
Evaluation, Vol. 13A, eds. D. O. Thompson and D. E. Chimenti (Plenum
Press, New York, 1994), p. 1111.

Fulton, J; Hedengren, K; Young, J; Filkins, R; Patton, T: "Optimizing
the Design of Multilayer Eddy Current Probes - A Theoretical and
Experimental Study," Review of Progress in Quantitative Nondestructive
Evaluation, Vol. 16, eds. D. O. Thompson and D. E. Chimenti (Plenum
Press, New York, 1997, - to be published).

Patton, T.; Filkins, R.; Fulton, J.: Hedengren, K.; Young, J.; Granger,
C.; Hewton, T.: "Development of a Hand-Held, Flexible Eddy Current Probe
for Inspection of Curving Surfaces," Review of Progress in Quantitative
Nondestructive Evaluatio*n,* Vol. 16, eds. D. O. Thompson and D. E.
Chimenti (Plenum Press, New York, 1997, - to be published).

**Point of Contact:** Dr. Christopher Smith, AAR-433, FAA William J.
Hughes Technical Center, Atlantic City International Airport, NJ 08405,
609-485-5221, FAX, 609-485-4569, e-mail: smithc@admin.tc.faa.gov

## Eddy Current Probe Design

Norio Nakagawa[^35], David Raulerson and Joe Chao[^36],

**Research Objective**

To improve the detection capability of nondestructive evaluation
inspections used on field hardware.

**Approach**

An electromagnetic field modeling capability which accepts input from a
standard computer aided design (CAD) package will be developed based on
the boundary element method (BEM). Modeling techniques such as BEM and
finite element method (FEM) will be utilized to determine
electromagnetic field characteristics for conventional element designs.
This will include applications of the modeling data to design elements
with controlled field shapes in an effort to increase sensitivity to key
flaw features while reducing effects from geometry. To test the accuracy
of the model, validation will be performed which includes correlation of
predicted and measured data for artificial as well as real flaws.

**Accomplishment Description**

The key feature added this year is the crack modeling capability.
Specifically, progress has been made in two areas: First, the core BEM
code has been enhanced so that it now includes crack modeling
capability. Second, additional geometry pre-processing software that
assists model users to create cracks and their mesh interactively has
been developed. The overall procedure is now established, which guides a
user to follow a stream-lined processes, starting with part and probe
designs on CAD, generating a crack and scan plans, and then running the
BEM code. In addition, the code validation effort has been continued.
For solenoid coils, the code has been validated against edge signals and
some crack signals. Similar validation efforts with ferrite-core probes
are in progress.

**Significance**

The resulting software simulation capability to examine field eddy
current inspections has a significant potential to improve element
designs, inspection strategies, and overall performance of
safety-critical inspections. In particular, the computer model provides
this capability without incurring excessive costs, thus reducing
expensive empirical prototyping requirements.

**Expected Results**

The entire package, namely the general 3-D BEM code with crack modeling
capability plus the geometry pre- and post-processing procedure, is
expected to become available in 1998.

**References:** Chao, J; Raulerson, D; Nakagawa, N: "Boundary Element
Method Based Probe Design Model Validation," Review of Progress in
Quantitative Nondestructive Evaluation, 1996.

Chao, J; Nakagawa, N; Raulerson, D.; Moulder, J: "A General Boundary
Integral Equation Approach to Eddy Current Crack Modeling," Review of
Progress in Quantitative Nondestructive Evaluation, 1996.

**Point of Contact**: Dr. Christopher Smith, AAR-433, FAA William J.
Hughes Technical Center, Atlantic City International Airport, NJ 08405,
609-485-5221, FAX, 609-485-4569, e-mail: smithc@admin.tc.faa.gov

## Fatigue Crack Initiation and Growth in Riveted Specimens

Z. M. Connor, W. Li., M. E. Fine, and J. D. Achenbach[^37]

**Research Objective**

To compare fatigue crack development in riveted Aluminum clad aluminum
alloy specimens using an optical microscope and a scanning acoustic
microscope.

**Approach**

Riveted specimens were cycled in fatigue until cracks of various lengths
were observed using a microscope which was mounted to the testing
machine. A scanning acoustic microscope was used to obtain a C-scan of
each specimen in the area containing the fatigue crack. These specimens
were then carefully examined at various magnifications using high
resolution metallographic optical microscopes. Micrographs were taken of
areas containing fatigue cracks.

**Accomplishment Description**

With optical microscopy, surface rumpling consisting of microcracks and
roughness was an early indication of fatigue cracking on the flush
surface of the riveted lap joint specimens. Propagating cracks grew out
of the rumpled region. While the optical microscope reveals only surface
cracks, the scanning acoustic microscope in the C-scan mode images
subsurface cracks projected onto a plane. Such two-dimensional images
reveal subsurface cracks prior to their observance on the specimen
surface

**Significance**

A much better understanding of how fatigue cracks develop at rivets was
obtained. This will be an aid to inspection with current hardware as
well as an aid to improving nondestructive evaluation hardware.

**Expected Results**

Data on fatigue crack initiation and growth at rivets to aid in
development of inspection techniques and life prediction methodologies.

**Reference:** Connor, Z. M; Li, W.; Fine, M. E.; Achenbach, J. D.:
"Fatigue Crack Initiation and Growth in Riveted Specimens," Presented at
First International Conference on Fatigue Damage in Structural
Materials, Hyannis, Massachusetts, September 22-27, 1996.

**Point of Contact:** Dr. Christopher Smith, AAR-433, FAA William J.
Hughes Technical Center, Atlantic City International Airport, NJ 08405,
609-485-5221, FAX, 609-485-4569, e-mail: smithc@admin.tc.faa.gov

## A Study of Fatigue Crack Generation and Growth in Riveted Aluminum Clad 2024-T3 Specimens

Z. M. Connor, M. E. Fine, and B. Moran[^38]

**Research Objective**

To investigate microscopically the initiation and propagation of fatigue
cracks in riveted specimens as a function of fatigue cycles. To prepare
specimens with characteristic cracks for nondestructive evaluation
studies and calibration.

**Approach**

Specimens were fabricated under carefully controlled conditions
endeavoring to keep the specimens as nearly identical as possible. Each
lap joint specimen was assembled from two Aluminum clad 2024-T3 aluminum
alloy panels using three 2017-T4 aluminum alloy rivets. Specimens were
loaded in uniaxial tension with an R ratio of 0.1 and a maximum load of
4.0 kN. Cracks were viewed at a magnification of 40x with an optical
microscope mounted on an X-Y micrometer stage. Crack length measurements
were taken from the edge of the rivet head using the X-Y micrometer
stage.

**Accomplishment Description**

Crack growth measurements were obtained on five specimens at intervals
of every 10,000 cycles. The results of these measurements were plotted
as crack length vs. number of cycles as shown in the attached figure.
The number of cycles until a crack approximately 2 mm in length was seen
in the plate near the rivet head was defined as N~i~. The average N~i~
was 310,000 cycles with a standard deviation of 80,000 cycles. With in
situ optical microscopy, microcracks were found to be the earliest
indication of fatigue cracking on the surface of lap joint panels
riveted with flathead chamfered rivets. The propagating crack emerged
from the microcracks.

**Significance**

A much better understanding of how fatigue cracks develop at rivets was
obtained. This will be an aid to inspection with current hardware as
well as an aid to improving nondestructive evaluation hardware.

**Expected Results**

A method for producing in situ controlled damage, cracked riveted
specimens for nondestructive evaluation research and development.

**Reference:** Connor, Z. M.; Fine, M. E. ; Moran, B.: "A Study of
Fatigue Crack Generation and Growth in Riveted Alclad 2024-T3
Specimens," Presented at FAA-NASA Symposium on Continued Airworthiness
of Aircraft Structures, Atlanta, Georgia, August 28-30, 1996.

**Point of Contact:** Dr. Christopher Smith, AAR-433, FAA William J.
Hughes Technical Center, Atlantic City International Airport, NJ 08405,
609-485-5221, FAX, 609-485-4569, e-mail:
smithc@admin.tc.faa.gov![](media/image10.emf){width="7.281944444444444in"
height="5.3277777777777775in"}

## Economical Eddy Current

Kevin Smith, Rob Stephan, and Dave Raulerson[^39]

**Research Objective**

Provide generic tools applicable across engine make and model for engine
shop use. Design goals include low cost to capability ratio;
portability; use in on-wing, module, and detail part levels.

**Approach**

Inspection technology in the commercial aviation industry has
traditionally been developed to address known problems (reactive
approach). The Engine Titanium Consortium (ETC) has worked with the
airlines to define common needs that occur across manufacturers and
across engine models. With the problem definition in hand, generic
tools, which allow controlled scanning and digital data acquisition,
have been developed. The ETC has developed a portable scanner and data
acquisition system for use in airline overhaul and maintenance shops.
The portable scanner consists of generic motor control and motion
hardware with application specific tooling for probe positioning and
manipulation. Adapter plates are used to mate the scanner to a variety
of engine disks by matching the bolt hole patterns for a given disk. A
lunch box computer has been used for data acquisition. The system
communicates with a variety of commercially available instruments.

**Accomplishment Description**

To date, bore scans (axial motion), web scans (radial motion), and bolt
hole scans have been demonstrated on engine hardware from all three
Original Equipment Manufacturers (OEMs) involved in ETC. Betasite
testing has been completed at United Airlines were efforts focused on
inspection of the 14^th^ and 15^th^ stages of the JT9D compressor disks.
Betasite testing is under way with Northwest Airlines and Allied Signal,
and is planned in 1997 with American Airlines. In addition to the
intended jet engine applications, minimal adaptation has enabled the
technology to be applied to inspection of propeller blades, reportedly
saving the industry over \$10M.

**Significance**

Cost effective tooling that can be used on multiple applications
enhances the potential for controlled scanning in an airline maintenance
environment. This has the potential to increase the sensitivity,
reliability, and productivity of engine inspection.

**Expected Results**

Generic, cost effective tools for inservice inspection of aircraft
engines.

**Reference:** Raulerson, D. A.; Smith, K. D.; Stephan, R. R.:
"Economical Approach to Automated Eddy Current Inspection," Presented at
the SPIE conference, Scottsdale, Arizona, December 3-5, 1996.

**Point of Contact:** Dr. Christopher Smith, AAR-433, FAA William J.
Hughes Technical Center, Atlantic City International Airport, NJ
08405,609-485-5221, FAX, 609-485-4569, e-mail: smithc@admin.tc.faa.gov

## Eddy Current Methods for Crack Detection 

J. C. Moulder[^40]

**Research Objective**

Develop new methods for detection and characterization of small fatigue
cracks in layered aircraft structures using pulsed eddy current
techniques.

**Approach**

A novel scanned, pulsed eddy-current instrument developed for use in
detection of hidden corrosion in lap splices is being adapted for
detecting small fatigue cracks in layered aircraft structures. The
advantages of pulsed eddy current techniques include the wide bandwidth
attainable, which permits a single probe and a single measurement to
provide information over a broad frequency range. Pulsed eddy current
signals can be used to characterize cracks and locate them in the depth
of the material. Time gating of the pulsed eddy-current signals provides
a means to discriminate against interfering signals from lift off, air
gaps, and fasteners.

**Accomplishment Description**

The pulsed eddy current instrument has been significantly improved by
the addition of scanning capability and the ability to select pulse
repetition rate and amplitude under computer control. Now,
two-dimensional scans of surface areas can be performed and data
presented in a pseudo-color, C-scan format. Commercial probes of many
different designs can be used easily, since the amplitude and timing of
the excitation pulse have been brought under software control. The
ability to discriminate crack signals from interfering signals caused by
lift off, air gaps, and fasteners has been demonstrated on simulated lap
splice specimens with electro-discharge machine (EDM) slots in the
second layer. Slots as small as 60 mils can be detected in the second
layer of 65-mil lap splices. A reference standard for the DC-9 rudder
rib structure provided by Northwest Airlines has also been examined with
the instrument to evaluate potential applications to field inspection of
this complicated, 4-layer structure. Potential eddy current (PEC)
measurements were able to detect all the simulated defects in the
standard and to determine the location of each of the flaws by using the
time-gating feature to isolate individual flaws. The instrument has been
licensed to Sierra-Matrix, a California-based company, for
commercialization.

**Significance**

The principal advantage of the instrument developed in this program, as
compared to the commercially available pulsed eddy current instrument,
is that our instrument can be adapted for use with any commercial probe
and used in any type of geometry normally inspected with eddy currents.
The commercial instrument employs a rotating Hall-effect sensor and is
designed specifically for fastener inspection. The ability to scan and
image areas up to 8 x 20 inches in size is a considerable advantage. The
time-gating feature included in this instrument is a unique capability
that has no parallel in commercial instrumentation.

**Expected Results**

Pulsed eddy current methods to detect small fatigue cracks in layered
structures.

**References:** Bieber, J. A.; Shaligram, S. K.; Rose, J. H.; Moulder,
J. C.: "Time-Gating of Pulsed Eddy Current Signals for Defect
Characterization and Discrimination in Aircraft Lap-joints," Review of
Progress in Quantitative Nondestructive Evaluation Vol. 16, edited by D.
O. Thompson and D. E. Chimenti, Plenum, New York, in press.

Moulder, J. C.; Bieber, J. A.; Ward, W. W.; Rose, J. H.: "Scanned
Pulsed-Eddy-Current Instrument for Non-Destructive Inspection of Aging
Aircraft," Proceedings of SPIE Conference on Nondestructive Evaluation
of Aging Aircraft, Airports, and Aerospace Hardware, Vol. 2945, SPIE,
Bellingham, WA, 1996, p. 2.

**Point of Contact:** David Galella, AAR-433, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ (609)
485-5784, FAX (609) 485-4569, e-mail: galellad@admin.tc.faa.gov

## Eddy Current Methods for Corrosion Detection

J. C. Moulder[^41]

**Research Objective**

Develop and evaluate eddy current inspection methods for the
quantitative determination of metal skin thickness in lap joints.
Techniques will be sought that provide quantitative information on the
amount of metal thinning induced by corrosion in multi-layer lap joints.

**Approach**

Eddy current techniques are being developed to characterize
corrosion-induced thinning of aircraft skin in lap joints. Two methods
have been developed: a swept-frequency method using conventional
instruments and a new technique using pulsed eddy currents. The first
method is based on multi-frequency measurements of probe impedance. A
pulsed eddy current method based on a newly developed prototype
instrument provides similar results in a faster, more economical field
instrument. Development of a theoretical model for the pulsed eddy
current technique was carried out in a related program funded by Air
Force Office of Scientific Research (AFOSR).

**Accomplishment Description**

Development and testing of the scanned pulsed eddy current instrument
has been completed this year. A new feature, time-gating, has been added
to the software that permits discrimination against interfering signals
from lift off, air gaps, or fasteners. Time-gating also provides a means
to determine the location of detected corrosion in multi-layered
structures. A number of field demonstrations have been carried out,
including demonstrations at the Aging Aircraft Nondestructive Inspection
Validation Center (AANC), at the Air Transport Association (ATA)
Nondestructive Testing Forum in Seattle, WA, at United Airlines
Maintenance Facility in Oakland, CA, and at Kelly Air Force Base,
McClellan Air Force Base, and Robins Air Force Base. A round-robin test
on a Boeing-supplied corrosion specimen was carried out to compare
quantitative results from ultrasound, x-ray, thermal wave, and eddy
current methods. The instrument has been licensed to Sierra-Matrix, a
California-based company, for commercialization

**Significance**

Both of these new techniques are capable of quantitatively determining
the thickness of both layers in a lap splice and discriminating between
loss of metal and larger than normal gaps in the lap joint. This is in
contrast to the techniques presently in use, which are qualitative
rather than quantitative and do not provide information about corrosion
in the second layer. The swept frequency technique, when implemented
with available commercial equipment, is a difficult and time-consuming
measurement. The new pulsed eddy current instrument is fast and
portable, and provides essentially the same amount of information.

**Expected Results**

Improvements to both hardware and software are in progress. A new
electronics card is being designed that will incorporate a DSP board and
constant current excitation. A major upgrade in the software is also
underway, with the major goal being added flexibility in use of
different scanners and data acquisition boards, as well as improved
imaging. A technology transfer program is planned in collaboration with
AANC, the Center for Aviation Systems Reliability, a commercial vendor,
original equipment manufacturers and commercial airlines to develop
practical field applications for the new scanned pulsed eddy current
instrument.

**References:**. Bieber, J. A.; Shaligram, S. K.; Rose, J. H.; Moulder,
J. C.: "Time-Gating of Pulsed Eddy Current Signals for Defect
Characterization and Discrimination in Aircraft Lap-Joints," Review of
Progress in Quantitative Nondestructive Evaluation Vol. 16, edited by D.
O. Thompson and D. E. Chimenti, Plenum, New York, in press.

Moulder, J. C.; Bieber, J. A.; Ward, W. W.; Rose, J. H.; "Scanned
Pulsed-Eddy-Current Instrument for Non-Destructive Inspection of Aging
Aircraft," Proceedings of SPIE Conference on Nondestructive Evaluation
of Aging Aircraft, Airports, and Aerospace Hardware, Vol. 2945, SPIE,
Bellingham, WA, 1996, p. 2.

**Point of Contact:** David Galella, AAR-433, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ (609)
485-5784, FAX (609) 485-4569, e-mail: galellad@admin.tc.faa.gov

## 

## Frequency-Shifted Sagnac Interferometer for Ultrasound Measurement 

Pavel Fomitchov, Sridhar Krishnaswamy, and J. D. Achenbach[^42]

**Research Objective**

Develop equipment and a technique for laser-based ultrasonics to detect
small cracks and to characterize other defects in aircraft structures.
In particular, develop a robust frequency-shifted Sagnac interferometer
to detect and measure pzt- and laser-generated ultrasonic signals.

**Approach**

The Sagnac interferometer is a truly path-matched device and it is a
more robust alternative to the heterodyne or path-stabilized Michelson
interferometer. The interferometer is suitable for the measurement of
ultrasonic surface and plate waves arising from laser- or pzt-generated
sources or from acoustic emissions. The device will form part of a
complete laser-based ultrasonic nondestructive evaluation system. This
system can be used to detect and characterize discrete defects such as
fatigue cracks as well as distributed regions of reduced material
properties.

**Accomplishment Description**

We have developed a frequency-shifted Sagnac interferometer to measure
ultrasonic signals. The main advantage of the proposed interferometer is
improved signal-to-noise ratio, which has been achieved using an optical
frequency shifting technique for biasing to quadrature and for
elimination of parasitic interference between sampling beams and other
beams in the interferometer. The results of application of the Sagnac
sensor for artificial crack detection and detection of laser generated
ultrasonic waves are shown in Fig. 1, 2. A first prototype of the
interferometer was demonstrated at the Air Transport Association
meeting, Sept. 30-Oct. 1, 1996, Seattle, Washington.

**Significance**

The measurement of ultrasonic signals by laser interferometric
techniques has many advantages for applications to nondestructive
evaluation. These include noncontact point detection, remote placement
of equipment using fiber-optics, easy scanning, absolute displacement
calibration, both broad band and narrow band signal detection, wide
frequency band measurements, and applicability to curved surfaces.

**Expected Results**

A robust low-cost frequency-shifted Sagnac interferometer suitable for
field applications to detect and measure pzt- and laser-generated
ultrasonic signals in aircraft structures.

**References:** Fomitchov, P.; Krishnaswamy, S.; Achenbach, J.:
"Fiberized Sagnac Interferometer for Ultrasound Measurement," Program
Quantitative NDE, vol. 15, edited by D.O. Thompson and D.E. Chimenti, p.
645-650 (Plenum Press, New York, 1996).

Fomitchov, P.; Steckenrider, J.; Krishnaswamy, S.; Achenbach, J.:
"Frequency-shifted low-noise Sagnac sensor for ultrasonic measurements,"
Review of Progress in Quantitative NDE, vol. 16.

Fomitchov, P.; Krishnaswamy, S.; Achenbach, J.: "Compact Phase-Modulated
Sagnac Interferometer for Ultrasound Detection," submitted to Optics and
Laser Technology (Nov. 1996).

**Point of Contact**: David Galella, AAR-433, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ (609)
485-5784, FAX (609) 485-4569, e-mail: galellad@admin.tc.faa.gov

![](media/image11.emf){width="8.544444444444444in"
height="3.8208333333333333in"}

## 

## Self-Compensating Ultrasonic Technique

I. N. Komsky and J. D. Achenbach[^43]

**Research Objective**

To adapt a self-compensating ultrasonic technique that has been
developed at Northwestern University to detect and characterize defects
in aircraft structures.

**Approach**

An ultrasonic self-compensating technique has been further developed.
The approach does not require a separate calibration for each test. The
measuring device as a unit can be placed on the structure. The technique
requires the transducers to be fired in a specific sequence. An
appropriate combination of the measurements then yields data that is
uncoupled from the measurement device. Work is in progress on the
detection and characterization of radial cracks emanating from fastener
holes. The careful selection of frequencies, insonification angles, and
transducer spacing optimizes the penetration of ultrasound into second
layers. Applications to crack detection in a second layer are
specifically directed to the detection and characterization of fatigue
cracks in the second layer of a bolted spar-cap/strap connection of the
DC-10, the rear spar of a DC-9/MD-80 wing, and the second layer of a
Boeing 767 multilayered structure. A related technique developed on this
task has been used for detection of material loss by corrosion and
detection of stress corrosion cracks in a tee cap of the DC-9 wing box.

**Accomplishment Description**

The self-compensating ultrasonic technique has been utilized to detect
and image fatigue cracks, beneath the repair angle, in the rear spar aft
flange in the wings of DC-9 and MD-80 aircraft. Work was carried out to
image flaws in the second layer of a multilayered configuration with
PR1422 sealant in between the doubler and the rear spar. Two sets of
specimens with notches (0.062" - 0.125" in lengths, 0.032" - 0.064" in
depths) have been machined to model fatigue cracks around fasteners. The
transducer configuration has been developed to detect and characterize
cracks underneath the tapered doubler in the area covered by the
fastener nut. A prototype search unit was developed and fabricated.
Images of the artificial flaws were successfully generated in laboratory
tests and the results have been presented to Nondestructive evaluation
(NDE) engineers at DAC, Northwest Airlines, and Midwest Express. Initial
modifications of the first design have been made to fit the actual
spacing in the DC-9/MD-80 wing (upon request of DAC and Northwest
Airlines). The unit has been demonstrated.

A set of calibration specimens to model a DC-10 spar-cap/strap
connection was fabricated. This calibration set included specimens with
artificial cracks of different sizes and orientations as well as
specimens with implanted fatigue cracks. The first prototype of a
rotational scanning system was manufactured and integrated with the data
acquisition system and the ultrasonic probes. The inspection system has
been tested on the calibration specimens. Images of rectangular and
triangular artificial cracks (0.050" - 0.250" in lengths, 0.050" -
0.0250" in depths) as well as implanted fatigue cracks on the top and at
the bottom of the second layer have been successfully generated. The
ultrasonic technique developed for the DC-9 wing box inspection has also
had successful spin-offs to other applications.

**Significance**

The self-compensating ultrasonic technique makes it possible to detect
and size fatigue cracks, stress corrosion cracks, and material corrosion
loss in internal layers of airplane structures. The inspection can be
accomplished from the outside of the structure. Ultrasonic waves
penetrate from the external layers of a multilayered structure through
the sealant to the internal layer.

**Expected Results**

An ultrasonic inspection technique to replace the current procedure for
ultrasonic testing of a spar-cap/strap connection of the DC-10 with a
procedure for crack sizing. After review by the staff and approval by
the Federal Aviation Administration (FAA) the new procedure for an
ultrasonic inspection of the DC-10 spar-cap/strap connection will be
available as a Supplemental Inspection Document for nondestructive
inspection of the DC-10. Work will be continued on the ultrasonic
procedure for application of the self-compensating technique to
ultrasonic inspection of the DC-9/MD-80 rear spar, and Boeing 767
structures.

**References:** Komsky, I. N.: "Ultrasonic Scanning of Airplane and
Bridge Structures," Paper summaries of the first U.S.-Japan Symposium on
Advances in NDT, June, 1996.

Komsky, I. N.; Achenbach, J. D.: "Ultrasonic Imaging of Corrosion and
Fatigue Cracks in Multilayered Airplane Structures," Proceedings of the
SPIE Conference on Nondestructive Evaluation Techniques for Aging
Infrastructure and Manufacturing, December, 1996.

Komsky, I. N.; Achenbach, J. D.: "Optimization of an Imaging Procedure
for Ultrasonic Inspection of Airplane and Bridge Structures,"
Proceedings of the SPIE Conference on Nondestructive Evaluation
Techniques for Aging Infrastructure and Manufacturing, December, 1996.

**Point of Contact:** David Galella, AAR-433, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ (609)
485-5784, FAX (609) 485-4569, e-mail: galellad@admin.tc.faa.gov

## Self-Focusing Ultrasonic System

J. D. Achenbach and W. A. K. Deutsch[^44]

**Research Objective**

Develop a technique for self-focusing of a linear array on internal
defects, surface defects and small cracks in thin sheets.

**Approach**

Focused transducers are frequently used to increase the amplitude of
ultrasonic signals. Since a focused transducer has a fixed focal point,
mechanical movement of the transducer in both the normal and lateral
directions is generally required to focus on a defect, resulting in long
inspection times and extra costs for a scanning system. Alternatively a
transducer array together with a phased array technique can be used to
focus ultrasound with the advantage that the focal point can be moved
electronically without mechanical motion of the transducer array. In
this work a simple idea for electronic self-focusing of a linear array,
previously developed at Northwestern University, has been extended to
Rayleigh and Lamb Waves. The self-focusing procedure first measures the
backscattered signals for a first transmission by a single element of
the array. A cross correlation technique is used to determine the
time-of-flight differences of the back scattered signals to the elements
of the array. These time delays are subsequently used to adjust the
times of excitation of the elements for transmission focusing on the
defect. Using the differences in arrival times once more the
backscattered signals after transmission focusing are aligned for
reception focusing.

**Accomplishment Description**

A prototype system consisting of a linear array, together with the
required electronics has been built, and the necessary software has been
developed. Experimental results have demonstrated the ability of the
system to self-focus surface waves on a defect. For the case that there
is more than one defect, the technique has been extended to focus on the
defect that produces the largest backscattered signal. The technique has
also been used for Lamb wave focusing. Experiments have been
successfully performed to focus the first antisymmetrical Lamb mode on a
small crack-like (100 µm) defect in an aluminum sheet.

**Significance**

Development of self-focusing ultrasonic systems will offer the aviation
industry improved capability to detect small cracks, high density
inclusions, hard interstitial defects, and stress induced porosity in
aircraft structures and engine components.

**Expected Results**

A prototype and a methodology, validated by experiments, for
self-focusing of body waves, surface waves and Lamb waves on small
defects whose location is not known a-priori.

**References:** Deutsch, W.; Cheng, A.; Achenbach, J.: "Self-Focusing
Surface Wave Array," presented at Review of Progress in Quantitative
Nondestrucive Evaluation, MA, July 28-August 2, 1996.

**Point of Contact:** Dr. Christopher Smith, AAR-433, FAA William. J.
Hughes Technical Center, Atlantic City International Airport, NJ 08405,
609-485-5221, FAX = 609-485-4569, e-mail: smithc@admin.tc.faa.gov

![](media/image12.emf){width="6.706944444444445in"
height="4.248611111111111in"}

## Quantitative Thermal Wave Imaging of Corrosion Thinning

R.L. Thomas, L.D. Favro, and P.K. Kuo[^45]

**Research Objective**

To establish the usefulness of thermal wave imaging (TWI) as a
noncontacting, large area nondestuctive inspection technique for the
evaluation of material defects, corrosion and damage in adhesively
bonded and composite structures, and to determine the applicability of
TWI to the inspection of adhesive bonds in tear strap geometry's, rear
surface skin corrosion, bonded composite joints, and bonded boron-epoxy
composite doublers on aircraft.

**Approach**

To investigate TWI for inspection of adhesive bonds in tear strap
geometries, rear surface skin corrosion, as well as several common
composite defects, in close collaboration with manufacturers and
maintenance departments of air carriers.

**Accomplishment Description**

We have developed a real-time algorithm for quantitative measurement of
the percentage material loss from the rear surface of an aircraft skin.
The algorithm is capable of measuring corrosion loss of less than a
percent. This sensitivity and quantitatively checked by comparison with
ultrasonic time-of-flight and profilometer/micrometer measurements on a
series of corrosion test panels.

**Significance**

The thermal wave technique has excellent potential for rapid, wide-area
detection, imaging, and quantitative measurement of rear surface skin
corrosion on aircraft, either through the use of the uncorroded skin
itself as a calibration, or an external calibration thickness standard.
If the latter is used, the method provides absolute thickness
measurements.

**Expected Results**

Thermal wave imaging and measurement methodology and instrumentation.

**References:** Han, X.; Favro, L.; Ahmed, T.; Ouyang, Z.; Wang, L.;
Wang, X.; Zhang, F.; Kuo, P.; Thomas, R.: "Quantitative Thermal Imaging
of Corrosion on Aircraft," Review of Progress in Quantitative
Nondestructive Evaluation, Vol. 16, Ed. D.O. Thompson and D.E. Chimenti,
Plenum Press, New York, to be published (1997).

**Point of Contact**: David Galella, AAR-433, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ (609)
485-5784, FAX (609) 485-4569, e-mail: galellad@admin.tc.faa.gov

![](media/image13.emf){width="4.317361111111111in"
height="4.018055555555556in"}

Comparison of thermal and ultrasonic measurements of

a corrosion test panel with direct, profilometer measurement

of rear surface material loss.

## 

## 

## Nondestructive Evaluation Parameters - Failure Property Relationships of Adhesive Bonds

H. A. Aglan[^46]

**Research Objective**

To correlate parameters detectable by nondestuctive evaluation (NDE)
techniques (thermal and ultrasonic) with critical failure properties,
mainly the residual strength and fatigue lifetime of industry standard
bonded boron/epoxy repair patches. The effect of the number of plies on
the critical failure properties is also to be investigated.

**Approach**

Mechanical fatigue aging of adhesive joints is to be performed.
Boron/epoxy doublers with various number of plies will be fabricated and
used to repair pre-cracked aluminum substrates. The extent of adhesive
and doubler damage induced by the mechanical cyclic aging is to be
assessed.

**Accomplishment Description**

Monotonic and fatigue tension tests were conducted on multiple cracked
aluminum substrates bonded with boron/epoxy patches of different number
of plies (2, 4, and 6) to evaluate their residual strength and fatigue
behavior. Analysis of the results revealed that the 4 and 6 ply
boron/epoxy doublers can bring up the residual strength of the
pre-cracked aluminum substrate to the ultimate strength of the aluminum.
Post failure examination of specimens indicated that the failure is
always in the aluminum away from the boron/epoxy patch. A considerable
retardation in the crack growth was observed in the repaired pre-cracked
substrates with 4 and 6 plies. The repaired substrates with 4 and 6
plies exhibit considerable crack retardation without catastrophic
failure. Fatigue induced damage in the form of disbond between the
boron/epoxy doubler and the aluminum substrate has been captured using
thermal wave imaging. The size of the disbond tapered down in the
direction of the propagating crack in the aluminum substrate, forming a
triangular shaped disbond.

**Significance**

Boron/epoxy doublers with 4 plies or more increase the residual strength
of pre-cracked aluminum substrate to a value equal to the strength of
uncracked aluminum substrate. A considerable enhancement of the fatigue
lifetime was observed due to the crack arrest nature of the boron/epoxy
doublers with more than 4 plies.

**Expected Results**

Construct the relationship between the residual strength as well as
fatigue lifetime and the number of plies in the boron/epoxy doublers. In
addition, the mechanisms by which the boron/epoxy doublers arrest flaws
and defects in aluminum substrate will be established in view of the
fatigue crack propagation kinetics of the joints.

**References**: Aglan, H.; Abdo, Z.: "An Innovative Approach to Fatigue
Disbond Propagation in Adhesive Joints," Journal of Adhesion Science and
Technology, Vol. 10, No. 3 (1996), 183-198.

Aglan, H.; Abdo, Z.: "Effect of Surface Treatment on the Fatigue Failure
Behavior of Structural Adhesive Joints," Journal of Materials Science
Letters, Vol. 15 (1996), 469-472.

Aglan, H.; Allen, C.; Rowell, T.; Elleithy, R.; Ahmed, T.; Thomas, R.:
"Fatigue Failure Evaluation of Boron/Epoxy Patches Using Thermal Wave
Imaging" submitted for publication in Vol. 16, Review of Progress in
Quantitative Nondestructive Evaluation.

**Point of Contact:** David Galella, AAR-433, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ (609)
485-5784, FAX (609) 485-4569, e-mail: galellad@admin.tc.faa.gov

## Feasibility Study of a Rotorcraft Health and Usage Monitoring System (HUMS): Usage and Structural Life Monitoring Evaluation

B. Dickson, J. Cronkhite, S. Bielefeld, L. Killian, and R. Hayden[^47]

**Research Objective**

To evaluate two techniques, Flight Condition Recognition (FCR) and
Flight Load Synthesis (FLS), for usage monitoring and assess the
potential benefit of extending the retirement intervals of lifelimited
components, thus reducing the operator\'s maintenance and replacement
costs. Both techniques involve indirect determination of loads using
measured flight parameters and subsequent fatigue analysis to calculate
the life expended on the lifelimited components. To assess the potential
benefit of usage monitoring, the two usage techniques were compared to
current methods of component retirement. In addition, comparisons were
made with direct load measurements to assess the accuracy of the two
techniques.

**Approach**

The data that was used for the evaluation of the usage monitoring
techniques was collected under an independent HUMS flight trial program,
using a commercially available HUMS and data recording system. The usage
data collect from the HUMS trial aircraft was analyzed off-line using PC
based software that included the FCR and FLS techniques. In the future,
if the technique prove feasible, usage monitoring would be incorporated
into the onboard HUMS. The benefit of usage monitoring was identified
under work accomplished during the first phase of this activity.

**Accomplishment Description**

For the selected dynamic components analyzed, the results of the
evaluation of the FCR and FLS techniques indicate a potential for

extending retirement lives. This is due to the damage accumulation rate
for the FCR and FLS techniques being slower ("slow clock") than the
current method using actual flight hours as the basis for retirement
times. Of course, the benefits of usage monitoring are dependent on how
the rotorcraft is operated. Based on the mission flown for this
rotorcraft, which flew work crews to offshore oil platforms, the flight
hours charged against retirement times could be reduced by 50% or
greater.

**Significance**

The operator would gain a considerable payback in reduced maintenance
costs due to extension of retirement intervals.

**Expected Results**

The FCR technique, which only modifies the helicopter maneuver spectrum
relative to the manufacturer\'s baseline, was considered more practical
and lower risk to implement compared to the FLS technique. However, the
FLS technique could be refined to overcome shortcomings found and
benefit rotorcraft.

**Reference:** Dickson, B; Cronkhite, J.; Bielefeld, S.; Killian, L.;
Hayden, R.: Feasibility Study of a Rotorcraft Health and Usage
Monitoring System (HUMS): Usage and Structural Life Monitoring
Evaluation, DOT/FAA/AR-95/9, February 1996.

**Point of Contact:** William Emmerling, AAR-432, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-4009, FAX: (609) 485-8846, e-mail:
william_emmerling@admin.tc.faa.gov

## Feasibility Study of a Rotorcraft Health and Usage Monitoring System (HUMS): Results of Operator\'s Evaluation

Raylund Romero and Harold Summers[^48], James Cronkhite[^49]

**Research Objective**

To evaluate the feasibility of a state of the art health and usage
monitoring system (HUMS) to provide monitoring of critical mechanical
systems on the helicopter, including motors, drive train, engines and
lifelimited components.

**Approach**

The implementation of HUMS and cost integration with current maintenance
procedures was assessed from the operator's viewpoint in order to
achieve expected benefits from these systems, such as enhanced safety,
reduced maintenance cost and increased availability. An operational HUMS
was used as a basis for this study that was installed and operated under
an independent flight trial program. The HUMS equipment and software
were commercially available.

**Accomplishment Description**

The HUMS used in the flight trial program generally demonstrated a high
level of reliability in monitoring the rotor system, engines, drive
train and lifelimited components. The system acted as a sentinel to warn
of impending failures. A worn tail rotor pitch bearing was detected by
HUMS, which had the capability for self testing to diagnose system and
sensor faults.

**Significance**

Examples of potential payback to the operator with HUMS were identified,
including reduced insurance cost through enhanced safety, lower
operating costs derived from maintenance credits, increased aircraft
availability and improved operating efficiency.

**Expected Results**

The success in realizing the potential benefits from HUMS technology was
found to depend on the operator, helicopter manufacturer, regulator
(FAA), and HUMS supplier working together. The interfacing of HUMS with
current operational procedures was assessed to require only minimal
revisions to the operator's maintenance manuals.

**Reference:** Romero, R.; Summers H.; Cronkhite, J.: Feasibility Study
of a Rotorcraft Health and Usage Monitoring System (HUMS): Results of
Operator\'s Evaluation, DOT/FAA/AR-95/50, February 1996.

**Point of Contact:** William Emmerling, AAR-432, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-4009, FAX: (609) 485-8846, e-mail:
william_emmerling@admin.tc.faa.gov

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## Corrosion and Corrosion Fatigue of Airframe Materials

Robert P. Wei and D. Gary Harlow[^50]

**Research Objective**

The development of basic understanding of the processes of localized
corrosion and corrosion fatigue crack nucleation and growth in high
strength aluminum alloys used in airframe construction, the formulation
of kinetic models for these elemental processes, and the integration of
these models into probabilistic models that can provide guidance in
formulating methodologies for service life prediction.

**Approach**

In support of the National Aircraft Research Program of the Federal
Aviation Administration (FAA), Lehigh University undertook a
multidisciplinary program of research to study corrosion and corrosion
fatigue of airframe materials. The program is complemented by the Air
Force Office of Scientific Research (AFOSR).

**Accomplishment Description**

Summarizes research performed under the FAA sponsored (Phase 1) program
for the period from 15 June 1992 to 14 June 1995. Experimental efforts
during this period were focused upon (1) characterizations of the
chemical, microstructural and statistical aspects of pitting corrosion
and upon the kinetics of pitting of 2024-T3 aluminum alloy in aqueous
environments; (2) establishment of the criteria for the transition from
pitting to corrosion fatigue crack growth (i.e., crack nucleation); and
(3) studies of corrosion fatigue crack growth particularly in its early
stage.

**Significance**

The development and demonstration of a mechanistically based probability
approach for service life prediction and the formulation of a
probability model for particle induced corrosion pit growth that
pertains to multiple-site-damage (MSD) analysis in aircraft.

**Expected Results**

Methodologies to provide guidance for service life prediction of aging
aircraft.

**Reference:** Wei, R.; Harlow, G.: Corrosion and Corrosion Fatigue of
Airframe Materials, DOT/FAA/AR-95/76, 1996.

**Point of Contact:** Dr. Thomas Flournoy , AAR-433, FAA William J.
Hughes Technical Center, Atlantic City International Airport, NJ 08405,
Phone: (609) 485-5327, FAX: (609) 485-4569, e-mail:
Thomas_Flournoy_at_ct27@admin.tc.faa.gov

## Stochastic Modeling of Antisymmetric Buffet Loads on Horizontal Stabilizers in Massively Separated Flows

Saeed Farokhi, Clay S. Mauk, and James E. Locke[^51]

**Research Objective**

Identify buffet loading as a structural design load.

**Approach**

The surface pressure was measured for a large number of test conditions,
including the most severe buffeting environment for this type of
aircraft: the tail immersed in the massively separated wake of the wing.
Modifications to the Beech King Air 200 one-sixth scale wind tunnel
model included the construction of a new horizontal stabilizer
instrumented with 12 miniature pressure transducers. Structural
characteristics of the full-scale aircraft were estimated using the
ASTROS (Automated STRuctural Optimization System) program.
Motion-dependent aerodynamics (stiffness and damping) were computed
using the proven doublet lattice method, which is incorporated into
ASTROS. Due to difficulties with the data acquisition system, the
current approach was validated with buffet pressure power spectral
densities from an existing reference.

**Accomplishment Description**

A modern method was developed to model antisymmetric buffet design loads
on horizontal stabilizers with a known probability, utilizing a rigid
wind tunnel model.

**Significance**

Provides technical information concerning stabilizer antisymmetric
buffet loads.

**Expected Results**

Aircraft manufacturers will be able to predict stabilizer antisymmetric
buffet loads early for use in the certification program.

**Reference:** Farokhi, S.; Mauk, C.; Locke, J.: Stochastic Modeling of
Antisymmetric Buffet Loads on Horizontal Stabilizers in Massively
Separated Flows, DOT/FAA/AR-95/7, March 1996.

**Point of Contact:** Thomas DeFiore, AAR-432, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-5009, FAX: (609) 485-4569, e-mail:
Thomas_DeFiore_at_ct27@admin.tc.faa.gov

## Fracture Testing Of Large-Scale Thin Sheet Aluminum Alloy

R. DeWit, R. J. Fields, S. R. Low III, D. E. Harne, T. Foeke[^52]

**Research Objective**

Characterize the process by which cracks propagate and link up on
large-scale, precracked aluminum alloy panels.

**Approach**

Extended grips and test fixtures were specially designed to tension load
the panel specimens in a 1780-kN capacity universal testing machine. Ten
single sheets of bare 2024-T3 aluminum alloy, approximately 4 m high,
2.3 m wide, and 1 mm thick were fabricated with simulated through cracks
oriented horizontally at midheight. Using existing information, a test
matrix was set up to explore regions of failure controlled by fracture
mechanics, with additional tests near the boundary between plastic
collapse and fracture. In addition, a variety of multiple site damage
(MSD) configurations were included to distinguish between various
proposed linkage mechanisms. All tests but one used antibuckling guides.
Three specimens were fabricated with a single central crack, six others
had multiple cracks on each side of the central crack, and one had a
single crack but no antibuckling guides.

**Accomplishment Description**

The data were analyzed by two different procedures: (1) the plastic zone
model based on the residual strength diagram and (2) the R-curve. The
first three tests determined the basic material properties; these
results were used in the analysis of the subsequent tests with MSD
cracks. The two analyses procedures provided good predictions of the
residual strength , linkups stress, and fracture stress of the panels
with MSD. The critical fracture stress predicted by both procedures are
virtually the same.

**Significance**

The large panels were used in this test to avoid net section control
failure. The two analyses procedures: (1) which is analytically based
and (2) experimentally based, produce the same results.

**Expected Results**

Experimental data for use in the verification of models for service life
prediction of aluminum alloy on aging aircraft.

**Reference:** DeWit, R.; Fields, R.; Low III, S.; Harne, D.; Foeke, T.:
Fracture Testing Of Large-Scale Thin Sheet Aluminum Alloy,
DOT/FAA/AR-95/11, February 1996.

**Point of Contact:** Dr. Paul Tan, AAR-433, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-6665, FAX: (609) 485-4569, e-mail:
Paul_Tan_at_ct27@admin.tc.faa.gov

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## Development Of A D Sight Aircraft Inspection System: Phase II

F. Karpala and O. L. Hageniers[^53]

**Research Objective**

Summarize the results of the second phase of development for a
nondestructive inspection (NDI) system based on D Sight.

**Approach**

Field trips were made to Northwest Airlines, in Minneapolis-St. Paul,
Air Canada, in Montreal, and United Airlines, in San Francisco A field
trip to the Royal Air Force in England is also summarized to include the
experience of corrosion detection with the original prototype developed
in Phase 1. The new hardware performed well in the field trials with the
exception of connectors. The new software philosophy using a turtle
diagram (planar surface drawing) graphical user interface was
particularly well received by NDI personnel.

**Accomplishment Description**

The accelerated corrosion testing continued from the first phase and
provided additional support for the high sensitivity of D Sight to
pillowing in corroded lap splices down to the 2 to 3 percent level.

**Significance**

The computer modeling and D Sight simulation provides a useful way to
study the potential of D Sight on new joint configurations as well as
draw attention to configurations that may be unacceptable for visual
inspection due to the low levels of pillowing that develop. The fact
that D Sight is measuring the extent of pillow deflections rather than
material loss also draws attention to techniques that may be reporting
false material loss simply because of aircraft manufacturing tolerances
especially at low corrosion levels.

**Expected Results**

A nondestructive inspection (NDI) system for aircraft based on D Sight.

**Reference:** Karpala, F.; Hageniers, O. L.: Development Of A D Sight
Aircraft Inspection System: Phase II, DOT/FAA/AR-95/15, March 1996.

**Point of Contact:** David Galella, AAR-433, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-5784, FAX: (609) 485-4569, e-mail: galellad@admin.tc.faa.gov

## Light Shaping Diffusers For Improved Visual Inspection Of Aircraft

Richard N. Shagam[^54]

**Research Objective**

Study representative flashlights available to inspectors of commercial
aircraft and description of the Light Shaping Diffuser (LSD), which
significantly improves the quality of the illuminating beam.

**Approach**

A review of the many different styles and varieties of flashlights on
the market today. Metal or plastic cases and reflectors, environmentally
sealed, adjustable or fixed beam spread, switch activated or
twist-on/twist-off, the battery size and number, and the type of bulb
are just some of the various differences to be found in flashlights.
Many share a common problem however, that of uneven illumination. A
flashlight beam having a high peak illumination value may also contain
areas of relatively brighter and dimmer intensities spread throughout
the beam.

**Accomplishment Description**

Tests of the LSD at aircraft inspection facilities demonstrated that the
LSD is a positive improvement to inspection flashlights. Ninety percent
of inspectors that responded to a survey distributed during the 1994 Air
Transport Association Nondestructive Testing Forum said the LSD-equipped
flashlight was an improvement over their standard flashlight.

**Significance**

The LSD is an inexpensive optical component, which is an improvement
over existing aircraft inspection flashlights.

**Expected Results**

LSD-equipped flashlights may replace present available commercial models
for aviation maintenance technicians.

**Reference:** Shagam, R.: Light Shaping Diffusers For Improved Visual
Inspection Of Aircraft, DOT/FAA/AR-95/32, April 1996.

**Point of Contact:** David Galella, AAR-433, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-5784, FAX: (609) 485-4569, e-mail: galellad@admin.tc.faa.gov

## Investigation of Fuselage Structure Subject to Widespread Fatigue Damage

M. L. Gruber, C. J. Mazur, K. E. Wilkins, and R. E. Worden[^55]

**Research Objective**

Focus on developing representative data on the effects of widespread
fatigue damage (WFD) on fuselage lap joints using unique Boeing testing
capabilities and expertise in airplane analysis and design.

**Approach**

The program was divided into three technical tasks: Task I consisted of
six tests conducted to characterize the 2024-T3 sheet material that was
used for the skins of the two curved fuselage pressure test panels
fabricated for Task II. Static, crack growth rate, and fracture
toughness data were generated. Task II involved curved pressure panel
tests conducted in the Boeing 127-inch-radius pressure test fixture. Two
panels representative of typical wide-body crown fuselage construction
were fabricated and tested. The testing focused on the outer skin lap
joint upper fastener row with and without simulated multiple site damage
(MSD). Lead cracks were grown in the lap joints by pressure cycling from
an initial 5-inch sawcut to two frame bays (40 inches). Task III
consisted of performing analyses to predict the crack growth and
residual strength of the pressure test panels. Stress intensity factors
were obtained from a material and geometric nonlinear finite element
model developed using standard Boeing practices. Crack growth rate and
residual strength predictions were made and correlated with test results
from the pressure tests.

**Accomplishment Description**

Test results indicated that the crack growth rate was faster in the
presence of MSD. The frame spanning the 40-inch crack was then cut and
residual strength tests conducted by gradually increasing the pressure
until dynamic panel failure occurred. Test results indicated that panel
residual strength was reduced 20 percent in the presence of MSD for the
panel geometry tested. The crack growth predictions correlated very well
with test results for small and no MSD cases.

**Significance**

The development of analytical prediction capability would be a major
step forward in developing an understanding of structures susceptible to
WFD.

**Expected Results**

Data for aging aircraft that can realistically simulate airplane
pressure loading.

**Reference:** Gruber, M.; Mazur, C.; Wilkins, K.; Worden, R.:
Investigation of Fuselage Structure Subject to Widespread Fatigue
Damage, DOT/FAA/AR-95/47, February 1996.

**Point of Contact:** Dr. Paul Tan, AAR-433, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-6665, FAX: (609) 485-4569, e-mail:
Paul_Tan_at_ct27@admin.tc.faa.gov

## 

## Automated Inspection of Aircraft

C. J. Alberts, C. W. Carroll, W. M. Kaufman, and M. W. Siegel[^56]

**Research Objective**

Develop robotic tools to assist aircraft inspectors by automating the
collection, archiving, and postprocessing of inspection data.

**Approach**

Studied the task of aircraft inspection, compiled the functional
requirements for an automated system to inspect skin fastener rows, and
developed a conceptual design of an inspection robot. The system was
designed to be sufficiently flexible to allow for the incor-poration of
new sensor technologies, including those being developed by other
organizations participating in the FAA's National Aging Aircraft
Research Program. A prototype of the robotic inspection system, (the
Automated Nondestructive Inspector or ANDI), was developed. The first
phase of system development resulted in a laboratory system that
demonstrated the ability to adhere to the surface of an aircraft panel
and deploy a standard eddycurrent sensor.

**Accomplishment Description**

Several operational limitations of the robotic inspection system were
exposed during the laboratory testing and field demonstration. Specific
problems include this prototype robot's speed, absence of a high-level
of operator control, and mechanical reliability.

**Significance**

As a result of these experiments, insight has been gained into means for
improving speed, ease of operation, and mechanical performance.

**Expected Results**

Establish the technical feasibility of using robotic nondestructive
inspection (NDI) systems in major aircraft maintenance facilities.

**Reference:** Alberts, C.; Carroll, C.; Kaufman, W.; Siegel, M.:
Automated Inspection of Aircraft, DOT/FAA/AR-95/48, June 1996.

**Point of Contact:** David Galella, AAR-433, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-5784, FAX: (609) 485-4569, e-mail: galellad@admin.tc.faa.gov

## 

## 

## 

## Validation of the Magneto-Optic/Eddy-Current Imager

Vanessa J. Brechling[^57], Floyd Spencer[^58]

**Research Objective**

Validate the MagnetoOptic/EddyCurrent Imager (MOI).

**Approach**

The MOI was used for inspections as part of an experiment designed to
assess the reliability of detecting a crack originating within fastener
holes in thin aluminum structures. Nondestructive test (NDT) personnel
performed the inspections in aircraft maintenance facilities. The
resultant probabilities of detection as a function of crack length for
the MOI were comparable to the values obtained for sliding probe eddy
current inspections.

The angle that a crack emanates from the rivet hole influences detection
rates. A 90 percent detection rate was achieved for 0.079 inch cracks
that were horizontal (in the direction of scan). The same detection rate
was achieved for 0.095 inch cracks, when the cracks were as much as 22°
off-horizontal. The stated crack lengths are as measured from the rivet
shank. There is no indication that the noted effect of the angle
direction on detection rates is a limitation of the MOI technology as
opposed to an inspector "bias" that could be addressed in procedures and
training.

The economic analysis considers the effects of individual factors that
contribute to the cost-effectiveness of the MOI. The possible returns to
the investment for a representative maintenance facility are calculated
using the net present value methodology. Specific characteristics are
defined for the representative facility, and then they are varied to
account for the differences in the maintenance community.

**Accomplishment Description**

In the sensitivity analysis, it is discovered that the most influential
factor on the economics of the MOI is the proportion of the inspection
time savings that can be transferred to decreasing aircraft downtime. If
the time savings achieved with the MOI can be implemented in the
inspection schedule to achieve at least a 10 percent decrease in
aircraft downtime, then the investment in the MOI is cost-effective for
the representative facility. This assumes that there is no need to strip
paint on the aircraft.

**Significance**

Productivity improvements and decreased aircraft downtime will result
from faster inspection techniques.

**Expected Results**

Even if inspection time savings are not transformed to decreased
aircraft downtime, if the MOI allows for the elimination of the
requirement to strip paint on one or more aircraft then it would be a
cost-effective investment for aviation maintenance.

**Reference:** Brechling, V.; Spencer, F: Validation of the
Magneto-Optic/Eddy-Current Imager, DOT/FAA/AR-95/100, November 1995.

**Point of Contact:** Dr. Christopher Smith, AAR-433, FAA William J.
Hughes Technical Center, Atlantic City International Airport, NJ 08405,
Phone: (609) 485-5221, FAX: (609) 485-4569, e-mail:
smithc@admin.tc.faa.gov

## 

## 

## A Methodology For The Economic Assessment Of Nondestructive Evaluation Techniques Used In Aircraft Inspection

Vanessa J. Brechling[^59]

**Research Objective**

Addresses the cost implications of: Nondestructive Evaluation (NDE)
techniques used for inspecting aircraft and a methodology is provided
for the economic evaluation of emerging NDE methods.

**Approach**

A model is described for the evaluation of the financial benefits net of
costs that an individual aircraft operator may expect to receive if it
employs a specific NDE technique. The calculation of the net present
value (NPV) of an investment in a new NDE technique to an aircraft
carrier or maintenance facility is described.

The net benefits to society were assessed upon the adoption of the new
technique. This analysis is particularly appropriate to investment
decisions that may be mandated by Federal Aviation Administration rule.
A social assessment of the costs and benefits of a new NDE technique
requires the assessment of all costs and benefits to society. The net
benefits to all aircraft operators in the industry are included in the
social calculation. In addition, costs to the government or any other
public institution, as well as any benefits that may accrue to members
of society from the airline industry's use of the new technique, are
included.

**Accomplishment Description**

Issues such as the measurement of benefits, the economic impact of an
improvement of the probability to detect flaws, heterogeneity in
aircraft maintenance facilities, estimation in the presence of
uncertainty, and the practical problem of obtaining the data necessary
to apply the methodology are addressed. It acknowledges that some of the
relevant factors are not measurable and that some subjective data may
have to be examined regarding the impact of these factors.

**Significance**

The availability of a formal cost benefit analysis for capital
investment decisions in inspection technology will facilitate the
introduction cost-effective inspection technology, and prevent the
expenditure of funds on the development and application inappropriate or
ineffective technologies.

**Expected Results**

A methodology to assess the economic benefit of nondestructive
inspection methods.

**Reference:** Brechling, V.: A Methodology For The Economic Assessment
Of Nondestructive Evaluation Techniques Used In Aircraft Inspection,
DOT/FAA/AR-95/101, November 1995.

**Point of Contact:** Dr. Christopher Smith, AAR-433, FAA William J.
Hughes Technical Center, Atlantic City International Airport, NJ 08405,
Phone: (609) 485-5221, FAX: (609) 485-4569, e-mail:
smithc@admin.tc.faa.gov

## 

## 

## 

## 

## 

## 

## 

## 

## 

## Tire Test Correlation: Radial Versus Bias-Ply Tires

Rich Anderson[^60]

**Research Objective**

Correlate the temperature performance of a radial tire with a bias-ply
tire of identical size under controlled laboratory dynamometer
conditions.

**Approach**

The general effects of increases in load and ground speed on the
temperature profiles of each tire were compared.

**Accomplishment Description**

The results indicated that the bias-ply tire used during the tests was
more adversely affected by increases in load and speed than the radial
tire.

**Significance**

Inspection and retreading intervals for radial tires have been
identified.

**Expected Results**

The data contained in this report could be used to review radial tire
inspection and retreading criterion.

**Reference:** Anderson, R.: Tire Test Correlation: Radial Versus
Bias-Ply Tires, DOT/FAA/AR-TN95/97, March 1996.

**Point of Contact:** William Nissley, AAR-432, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-4147, FAX: (609) 485-8846, e-mail:
William_Nissley@admin.tc.faa.gov

## 

## Flight Loads Data For A Boeing 737-400 In Commercial Operation

D. Skinn, P. Miedlar, and L. Kelly[^61]

**Research Objective**

Provide information that will enable the Federal Aviation Administration
(FAA) to better understand and control those factors that influence the
structural integrity of commercial transport aircraft.

**Approach**

During this prototype data collection program, 593 flights of
operational flight loads were collected. Of these, 535 flights
representing 817.7 hours, provided usable data. The National Aeronautics
and Space Administration (NASA) developed the specifications for the
recording system, defined the recording format, reduced the data to time
histories of engineering units, and tested and evaluated the algorithms
for data reduction and statistical reporting. The University of Dayton
Research Institute (UDRI) received the flight loads data and data review
software from NASA. UDRI developed software to reduce the flight loads
data and obtain additional parameters such as derived gust velocity and
continuous turbulence gust intensity.

**Accomplishment Description**

The data reduction includes, but is not limited to, analysis of c.g.,
accelerations, airspeeds, altitudes, flaps usage, and takeoffs and
landings. Data are typically presented in cumulative distribution
function or cumulative counts normalized to nautical mile or 1000 hours.

**Significance**

The data collection program is part of a joint FAA/NASA effort to
develop a flight recorder to obtain statistical loads data on commercial
transport (FAR Part 25) aircraft during routine operations.

**Expected Results**

Comparisons of typical usage with published Federal Aviation
Regulations.

**Reference:** Skinn, D.; Miedlar, P.; Kelly, L.: Flight Loads Data For
A Boeing 737-400 In Commercial Operation, DOT/FAA/AR-95/21, April 1996.

**Point of Contact:** Thomas Defiore, AAR-431, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-5009, FAX: (609) 485-4569, e-mail:
Thomas_DeFiore_at_ct27@admin.tc.faa.gov

## Axial Crack Propagation and Arrest in a Pressurized Fuselage

M. Kosai, A Shimamoto, and A. S. Kobayashi[^62]

**Research Objective**

Establish a practical crack kinking criterion that will account for the
presence of multiple site damage (MSD) in a pressurized airplane
fuselage.

**Approach**

The Ramulu‑Kobayashi crack kinking criterion was modified to predict
self‑similar crack propagation along a line of MSD and subsequent
kinking near a tear strap. Instrumented biaxial test specimens and
small‑scale fuselage rupture experiments were conducted to generate the
crack kinking and crack velocity data which were then input to a large
deformation, elastodynamic, finite element model of the fracture
specimen. The computed mixed mode I and II stress-intensity factors and
large axial stress preceding the propagating crack were used to evaluate
the self‑similar crack extension and the crack kinking criterion along
the crack trajectory.

**Accomplishment Description**

The biaxial specimen tests showed that in spite of the presence of plane
strain mode II stress-intensity factor and the large axial stress ahead
of the propagating crack, axial crack extension continued in the axial
direction if the skin is weakened by MSD. The crack propagated through
the tear strap in the presence of a continuous MSD and kinked when the
MSD terminated at the tear strap. However, if there is a short secondary
crack under the tear strap, the lead crack kinked at the long groove end
but reconnected to the secondary crack. The small‑scale fuselage rupture
tests showed that the pre‑existing axial through crack along the
stringer immediately kinked upon propagation due to the mixed mode I and
II state caused by the one‑sided opening of the crack flap. The
diagonally running crack subsequently turned circumferentially along the
tear straps or the two cylinder ends.

**Significance**

A practical crack kinking criterion was developed and verified for use
in predicting crack behavior in pressurized aircraft fuselages. Such a
criterion can be used in predicting residual strength in aircraft
structures.

**Expected Results**

A validated criteria for predicting axial crack propagation and arrest
in an aging aircraft.

**Reference:** Kosai, M.; Shimamoto, A.; Kobayashi, A. S.: Axial Crack
Propagation and Arrest in a Pressurized Fuselage, DOT/FAA/AR-95/43,
September 1996.

**Point of Contact:** Dr. Paul Tan, AAR-431, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-6665, FAX: (609) 485-4569, e-mail:
Paul_Tan_at_ct27@admin.tc.faa.gov

## 

## 

## 

## 

## Implications of Corrosion Pillowing on the Structural Integrity of Fuselage Lap Joints

Nicholas C. Bellinger and Jerzy P. Komorowski[^63]

**Research Objective**

Determine the effect that pillowing, caused by the presence of corrosion
products, has on fuselage lap joints.

**Approach**

Finite element techniques were utilized for analyses.

**Accomplishment Description**

The stress in a fuselage lap joint increases as the pillowing caused by
the corrosion by‑products increases. Depending on the location of the
corrosion and severity of the pillowing, the maximum stress in a joint
could shift to other skins (i.e. second or third layers) causing
premature crack initiation. The high aspect ratio cracks would make
detection difficult thus decreasing the probability of detection and
increasing the risk of premature failure.

**Significance**

Even a relatively small amount of pillowing can lead to yielding around
the rivet which must be accounted for in any analyses.

**Expected Results**

Verification of the model for service life prediction of corrosion
pillowing in an aging aircraft.

**Reference:** Bellinger, N. C.; Komorowski: "Implications of Corrosion
Pillowing on the Structural Integrity of Fuselage Lap Joints," FAA-NASA
Symposium on Continued Airworthiness of Aircraft Structures, August
28-30, 1996.

**Point of Contact:** Dr. Paul Tan, AAR-431, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-6665, FAX: (609) 485-4569, e-mail:
Paul_Tan_at_ct27@admin.tc.faa.gov

## 

## Introduction of Bonded Composite Doublers to Commercial Aircraft

Dennis Roach[^64], Aubrey Carter and John Marshall[^65]

**Research Objective**

Research and validate doubler applications on the commercial aircraft
fleet.

**Approach**

By focusing on a specific commercial aircraft application‑reinforcement
of the L‑1011 door frame and encompassing all application issues such as
design, analysis, installation, and nondestructive inspection, this
program is designed to prove the capabilities of composite doublers. In
addition to the L‑1011 project, a series of fatigue and strength tests
have been conducted on coupons and subsize test articles.
Tension‑tension fatigue and residual strength tests attempted to grow
engineered flaws in coupons with composite doublers bonded to aluminum
skin. Also, structures that modeled key aspects of the door corner
installation were subjected to extreme tension, shear, and bending
loads. In this manner it was possible to study the performance of the
Lockheed‑designed composite doubler using realistic aircraft load
scenarios.

**Accomplishment Description**

The data acquired was used to validate finite element models and
associated damage tolerance analyses.

**Significance**

The use of bonded composites may offer the airframe manufacturers and
airline maintenance facilities a cost effective technique to extend the
lives of their aircraft. Repairs and reinforcing doublers using bonded
composites have numerous advantages over mechanically fastened repairs
including corrosion resistance, light weight, high strength, elimination
of rivets, and installation time savings.

**Expected Results**

This research will provide technical data and service experience to
promulgate the use of composite doublers by U. S. operators.

**Reference:** Roach, D.; Carter, A.; Marshall, J: "Introduction of
Bonded Composite Doublers to Commercial Aircraft," Submitted for the
11th Technical Conference on composite Materials, American Society for
Composites, October 7-9, 1996.

**Point of Contact:** Peter Versage, AAR-433, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-5816, FAX: (609) 485-4569, e-mail:
Peter_Versage_at_ct27@admin.tc.faa.gov

## Corrosion of Aluminum Alloys in the Presence of Fire-Retardant Aircraft Interior Materials

J. E. Talia and J. Chaudhuri [^66]

**Research Objective**

Evaluate the potential for fire‑retardant materials used in aircraft
interiors to cause corrosion of aluminum structural alloys.

**Approach**

Service Difficulty Reports (SDR) data were studied to measure the
magnitude of the problem. Corrosion incidences were reviewed for several
aircraft types, and the most frequent locations for corrosion were
identified as fuselage, windows, and frames.

Laboratory experiments were designed and conducted for corrosion testing
of common aircraft structural alloys (Al 2024‑T3 and Al 7075‑T6) in the
presence of aircraft interior materials. Tests were conducted for
Ultrasuede, Glenlivit, and Highland Wool, common interior materials.
Accelerated corrosion test conditions included ambient and high
temperatures and high‑humidity conditions. Control specimens of the
aluminum alloys were tested as a baseline without exposure to the fire
retardants.

**Accomplishment Description**

For most test conditions, corrosion was increased for all three
fire‑retardant materials compared to the baseline tests without fire
retardants. Chemical composition from studies of the fire‑retardants
revealed substantial halogens in Ultrasuede and Glenlivit, but
negligible halogens in Highland Wool. Thus the corrosion results cannot
be attributed solely to halogens.

**Significance**

This work was initiated based on evidence of some corrosion problems in
general aviation, small business jets, and commuter aircraft in areas
near fire‑retardant interior fabrics and will provide data in order to
alleviate those problems.

**Expected Results**

Data will be utilized to alleviate corrosion problems in general
aviation, small business jets, and commuter aircraft in areas near
fire‑retardant interior fabrics.

**Reference:** Talia, J. E.; Chaudhuri, J.: Corrosion of Aluminum Alloys
in the Presence of Fire-Retardant Aircraft Interior Materials,
DOT/FAA/AR-95/81, October 1995.

**Point of Contact:** Dr. Thomas Flournoy, AAR-433, FAA William J.
Hughes Technical Center, Atlantic City International Airport, NJ 08405,
Phone: (609) 485-5327, FAX: (609) 485-4569, e-mail:
Thomas_Flournoy_at_ct27@admin.tc.faa.gov

## Maintenance Resource Management (MRM) Training and Crew Resource Management (CRM) Training: Analysis, Prescriptions, and Proscriptions

Ronald J. Lofaro[^67]

**Research Objective**

Examine the history and current status of CRM and the current efforts
and status of MRM. Provide an analysis of the similarities and
differences between MRM and CRM and will proceed to extract those
principles and practices of CRM that are relevant and appropriate for
developing and implementing MRM training. Present a prescription for MRM
training development, with proscriptions.

**Approach**

It has become an accepted fact that MRM training can-and,-should-be
developed using CRM training as its paradigm. On a higher level, it has
also been accepted that MRM, as a guiding philosophy and working set of
principles, should be based on CRM. While this last statement may be
accurate, the use of CRM training as a model for MRM training is replete
with serious problems. These problems are addressed.

**Accomplishment Description**

CRM and MRM are much less similar than has been realized and the
unquestioned use of CRM training as a paradigm for MRM training can,
has, and will lead to inappropriate and ineffective MRM training.

**Significance**

Recent developments and changes to CRM at both American and Delta
airlines are troublesome because they are using CRM to develop MRM.

**Expected Results**

This human factors research provides the groundwork for alerting the
airlines to the dangers of using CRM to develop MRM.

**Reference:** Lofaro, R.: "Maintenance Resource Management (MRM)
Training and Crew Resource Management (CRM) Training: Analysis,
Prescriptions, and Proscriptions," Aircraft Maintenance Technology, Vol.
7, No. 6, Ft. Atkinson, WI, September 1996.

**Point of Contact:** Dr. R. Lofaro, AAR-433, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-4501, FAX: (609) 485-4569, e-mail:
Ronald_Lofaro_at_ct27@admin.tc.faa.gov

## Visual Inspection Research Project Report on Benchmark Inspections

Floyd W. Spencer[^68]

**Research Objective**

Provide a benchmark measure of capability for visual inspections
performed under conditions that are realistically similar to those
usually found in major airline maintenance facilities.

**Approach**

Neither completely a laboratory study nor completely field observations.
A link between field and laboratory is provided by using visual
inspection tasks on a real airplane combined with other more controlled
tasks.

Twelve inspectors from four different airlines served as the subjects of
the experiment. Each subject spent two days at the Aging Aircraft
Nondestructive Inspection Validation Center (AANC) performing 10
different inspection tasks. Data collection consisted both of notes
taken by monitors and videotaping of the inspection tasks.

**Accomplishment Description**

Substantial inspector-to-inspector variation in performance was
observed. This observation has direct bearing on determining sample
sizes necessary to study the impact of visual inspection factors or the
effectiveness of specific interventions. On a specific task of looking
for cracks from beneath rivet heads the 90 percent detection rate crack
length for 11 inspectors ranged from 0.16 to 0.36 inch, with the 90
percent detection rate for the twelfth inspector being 0.91 inch.
Similar variations were observed in other inspection tasks.

Performance levels were task specific. Thus, an inspector's good
performance (relative to other inspectors) on one task does not
necessarily indicate a relatively good performance on other tasks. The
search component, as opposed to the decision component, of visual
inspection was the larger factor in determining performance levels for
most of the inspectors.

Major factors associated with performance in this research were the use
of job cards, thoroughness as reflected by total time, peripheral visual
acuity, and general aviation experience. Specifically, some of the
inspectors used job cards only to define the inspection task area.
Others used the information contained within the job card to direct
their attention to likely flaw locations The inspectors with lower
peripheral visual acuity scores showed a decline in performance on
certain tasks. Better performances were observed among inspectors with
more aviation experience as well as those that were more deliberate in
their inspections as reflected by the time taken on all tasks.

**Significance**

Visual inspection is the first line of defense for safety-related
failures on aircraft and provides the least expensive and quickest
method of assessing the condition of an aircraft and its parts. As such,
its reliability should be high and well-characterized.

**Expected Results**

Most of the research related to visual inspection has been in the area
of visual search conducted for industrial products and medical images.
This research targets aircraft specific visual inspections over a
variety of tasks. The tasks chosen represent different accessibility
levels, as well as different visual complexity levels. The benchmark
study is the first step in determining and characterizing the
reliability of visual inspection.

**Reference:** Spencer, F.: Visual Inspection Research Project Report on
Benchmark Inspections, DOT/FAA/AR-96/65, October 1996.

**Point of Contact:** Dr. Christopher Smith, AAR-433, FAA William J.
Hughes Technical Center, Atlantic City International Airport, NJ 08405,
Phone: (609) 485-5221, FAX: (609) 485-4569, e-mail:
smithc@admin.tc.faa.gov

## 

## 

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## 

## 

## 

## Variation in Load Factor experience of Fokker F27 and F28 Operational Acceleration Exceedance Data

J. B. de Jonge and P. A. Hol[^69]

**Research Objective**

Provide information that will enable the Federal Aviation Administration
(FAA) to better understand and control those factors that influence the
structural integrity of commercial transport aircraft.

**Approach**

Analyses and statistical summaries of normal acceleration data collected
from Fokker F27 and F28 aircraft representing 470,000 flights, which
were made by 101 aircraft belonging to 51 different operators.

**Accomplishment Description**

This activity supports the overall objectives of the FAA transport
flight loads data collection program which are (a) to determine whether
the loading spectra being used or developed for the design and test of
both small and large aircraft are representative of operational usage
and (b) to develop structural design criteria for future generations of
small and large aircraft.

**Significance**

This report demonstrated that even with limited flight parameters, in
this case only vertical acceleration, it is possible to identify
significant differences in aircraft operational usage. If aircraft
maintenance and inspection actions could be scheduled on actual usage,
considerable time and operating expense could be saved with no
degradation in safety.

**Expected Results**

Structural design criteria for future generations of small and large
aircraft.

**Reference:** de Jonge, J. B.; Hol, P. A.: Variation in Load Factor
experience of Fokker F27 and F28 Operational Acceleration Exceedance
Data, DOT/FAA/AR-96/114, December 1996.

**Point of Contact:** Thomas DeFiore, AAR-431, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-5009, FAX: (609) 485-4569, e-mail:
Thomas_DeFiore_at_ct27@admin.tc.faa.gov

## Engineering Approach to Damage Tolerance Analysis of Fuselage Skin Repairs

J. Bakuckas, P. Tan, and C. Bigelow[^70] C. Chen and J. Yu[^71]

**Research Objective**

Develop a new user‑friendly software tool, Repair Assessment Procedure
and Integrated Design (RAPID), capable of static strength and damage
tolerance analyses of simple fuselage skin repairs.

**Approach**

A simplified engineering approach to static and damage tolerance
analyses of riveted fuselage skin repairs has been incorporated in
RAPID. In this study, the damage tolerance analysis methodology in RAPID
was evaluated in terms of the fastener loads, stress‑intensity factor
solutions, crack growth, residual strength, and inspection schedule
calculations. Three example problems, each representing a typical
fuselage skin repair configuration, were analyzed. The analysis results
obtained from RAPID were compared with results generated using a
Representative Original Equipment Manufacturer (ROEM) method and a
special purpose finite element program for fracture mechanics analysis
and crack growth simulation in layered two­ dimensional structures.

**Accomplishment Description**

In general, results generated using RAPID were in good agreement with
results generated using the ROEM method and the finite element code.

**Significance**

A critical issue identified by the aviation industry is the need to
examine the effects of repairs on the structural integrity of aircraft.
The incorporation of damage tolerance methodologies in the maintenance
and repair practices of aging aircraft is required in order to insure
their continued airworthiness and operational safety. In December 1978,
the Federal Aviation Administration (FAA) amended their Fatigue
Evaluation Requirements to include a damage tolerance philosophy.
However, the majority of the current aircraft repairs are still designed
using static strength approaches. The resources needed for damage
tolerance designs of repairs are lacking, particularly for small
operators and independent repair facilities.

**Expected Results**

New user‑friendly software tool, Repair Assessment Procedure and
Integrated Design (RAPID), capable of static strength and damage
tolerance analyses of simple aircraft fuselage skin repairs.

**Reference:** Bakuckas, J.; Chen, C.; Yu, J.; Tan, P.; Bigelow, C.:
Engineering Approach to Damage Tolerance Analysis of Fuselage Skin
Repairs, DOT/FAA/AR-95/75, November 1996.

**Point of Contact:** John Bakuckas, AAR-431, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-4784, FAX: (609) 485-4569, e-mail:
John_Bakuckas@admin.tc.faa.gov

## 

## The Role of Fretting Fatigue on Aircraft Rivet Hole Cracking

D. Heoppner, C. Elliot III, and M. Moesser[^72]

**Research Objective**

Investigate the role of fretting corrosion and fretting fatigue on
aircraft rivet hole cracking.

**Approach**

A sensitivity study was conducted to determine the effects of fretting
on the fatigue lives of 2024-T3 clad sheet aluminum alloy riveted joint
specimens prepared with either FV or CE rivets using either C‑squeeze
riveting or a rivet gun with bucking bar riveting procedures. A method
was developed to predict coefficient of friction characteristics within
a fretted contact during the nucleation of a crack. This required
development of a system capable of determining the coefficient of
friction at controllable slip amplitudes as small as 80 microinches and
a verification test system. A finite element method was used to
calculate the state of stress at CE rivet locations where
fretting‑nucleated cracks were observed during the sensitivity study
portion of this grant program.

**Accomplishment Description**

It was found that fretting damage led to crack nucleation in all failed
specimens. It also was concluded that with respect to fretting fatigue
lives, based only on the results of this research program, the better
overall rivet is the 7050 FV rivet. In spite of the previous conclusion,
seven of eight specimens riveted by FV/C‑squeeze procedures had rivet
heads crack. This was concluded to have caused a reduction in specimen
lives.

**Significance**

A lap joint panel removed from an aircraft used in service was
investigated for evidence of fretting induced cracking in and adjacent
to the rivet holes. Cracks were found in all of the rivet holes that
were inspected. Generally, they had nucleated in regions where there was
evidence of fretting. This is significant because it indicates a
potential for multiple‑site damage occurring more rapidly than might be
anticipated from a fatigue analysis or testing that did not consider
fretting.

**Expected Results**

Fretting in riveted joints is a potentially major cause of crack
nucleation in aircraft skin structure. Its role in crack nucleation was
characterized in this work.

**Reference:** Heoppner, D.; Elliot, C.; Moesser, M.: The Role of
Fretting Fatigue on Aircraft Rivet Hole Cracking, DOT/FAA/AR-96/10,
October 1996.

**Point of Contact:** Dr. Thomas Flournoy, AAR-433, FAA William J.
Hughes Technical Center, Atlantic City International Airport, NJ 08405,
Phone: (609) 485-5327, FAX: (609) 485-4569, e-mail:
Thomas_Flournoy_at_ct27@admin.tc.faa.gov

# Catastrophic Failure Prevention

## Fiber-Reinforced Structures for Small Turbine Engine Containment

Pepin Associates[^73]

**Research Objective**

To design two fiber-reinforced structures for lightweight containment of
turbine rotor failures.

**Approach**

The first design was a hybrid core sandwich panel that can be used a
part of the airframe or nacelle structure and as a containment panel, if
required. The second was a collar or ring place close to the turbine
case wall of a turbofan, turboprop, or turboshaft engine.

**Accomplishment Description**

The program was focused on the design of structures to contain a
1-million in-lb. T053 tri-hub rotor burst using the lowest weight
containment structure possible. A combination of design, test article
fabrication, and spin pit testing were used. Design modification and
subsequent testing developed an understanding of the relationship
between reinforcement architecture, static elastic behavior, and dynamic
impact behavior of the ring and the panel. Additional spin pit testing
was done to evaluate changes in geometry and panel-to-panel joint
designs. Program results showed the potential of the hybrid sandwich
panel for application to containment barrier structures on aircraft and
helicopters.

**Significance**

Design characteristics were developed that can guide engine and airframe
structure designs to minimize weight and space requirements for
containment of internal turbine engine failures.

**Expected Results**

Design characteristics for structures to contain small rotor disk
bursts.

**Reference:** Fiber Reinforced Structures for Small Turbine engine
Fragment Containment (Phase II), DOT/FAA/AR-95/110, July 1996.

**Point of Contact:** Robert Pursel, AAR-432, FAA William J. Hughes
Technical Center, Atlantic City International Airport, NJ 08405, Phone:
(609) 485-6343, FAX: (609) 485-8846, e-mail:
Robert_Pursel@admin.tc.faa.gov

[^1]: Military Aircraft Division, Northrop Grumman Corporation, One
    Northrop Avenue, Hawthorne, CA 90250

[^2]: Military Aircraft Division, Northrop Grumman Corporation, One
    Northrop Avenue, Hawthorne, CA 90250

[^3]: Military Aircraft Division, Northrop Grumman Corporation, One
    Northrop Avenue, Hawthorne, CA 90250

[^4]: Mechanical, Aerospace and Nuclear Engineering Dept., UCLA, Los
    Angeles, CA 90024

[^5]: School of Aeronautics and Astronautics, Purdue University, West
    Lafayette, IN 47907

[^6]: AAR-431, FAA William J. Hughes Technical Center, Atlantic City
    International Airport, NJ 08405

[^7]: Galaxy Scientific Corporation, 2500 English Creek Ave., Bldg. 11
    Egg Harbor Township, NJ 08234

[^8]: Simula, Inc., 2500 English Creek Ave., Bldg. 11 Egg Harbor
    Township, NJ 08234

[^9]: Galaxy Scientific Corporation, 2500 English Creek Ave., Bldg. 11
    Egg Harbor Township, NJ 08234

[^10]: Galaxy Scientific Corporation, 2500 English Creek Ave., Bldg. 11,
    Egg Harbor Township, NJ 08234

[^11]: Foster-Miller, Inc., 350 Second Avenue, Waltham, MA 02154-1196

[^12]: SRI International, 333 Ravenswood Avenue, Menlo Park, CA
    94025-3493

[^13]: Aerotech Engineering and Research Corporation, 3125 W. 6th
    Street, Suite C, Lawrence, KS 66049-3101

[^14]: Galaxy Scientific Corporation, 2500 English Creek Avenue,
    Building 11, Pleasantville, NJ 08232

[^15]: FAA William J. Hughes Technical Center, AAR-431, Atlantic City
    International Airport, NJ 08405

[^16]: FAA Northwest Mountain Directorate, AMN-105N, 1601 Lind Ave.,
    S.W., Renton, WA 98055

[^17]: Center for Nondestructive Evaluation, Iowa State University,
    Ames, IA 50011

[^18]: Center for Nondestructive Evaluation, Iowa State University,
    Ames, IA 50011

[^19]: Pratt & Whitney Engines, West Palm Beach, FL 33410

[^20]: General Electric Aircraft Engines, Cincinnati, OH 45215

[^21]: Center for Nondestructive Evaluation, Iowa State University,
    Ames, IA 50011

[^22]: Pratt & Whitney Engines, West Palm Beach, FL 33410

[^23]: General Electric Aircraft Engines, Cincinnati, OH 45215

[^24]: Department of Civil Engineering, Northwestern University,
    Evanston, IL 60208

[^25]: Center for Nondestructive Evaluation, Iowa State University,
    Ames, Iowa 50011

[^26]: General Electric Aircraft Engines, Cincinnati, OH 45215-6301

[^27]: Center for Nondestructive Evaluation, Iowa State University,
    Ames, IA 50011

[^28]: Center for Nondestructive Evaluation, Iowa State University,
    Ames, IA 50011

[^29]: Center for Nondestructive Evaluation, Iowa State University,
    Ames, Iowa 50011

[^30]: GE Corporate Research & Development, P. O. Box 8, Schenectady, NY
    12301

[^31]: Center for Nondestructive Evaluation, Iowa State University,
    Ames, Iowa 50011

[^32]: Center for Nondestructive Evaluation, Iowa State University, 1915
    Scholl Road, Ames, IA 50011

[^33]: Center for Nondestructive Evaluation, Iowa State University,
    Ames, Iowa 50011

[^34]: General Electric Aircraft Engines, Cincinnati, OH 45215-6301

[^35]: Center for Nondestructive Evaluation, Iowa State University,
    Ames, IA 50011

[^36]: Pratt & Whitney Engines, West Palm Beach, FL 33410

[^37]: Northwestern University, Evanston, IL 60208-3020

[^38]: Northwestern University, Evanston, IL 60208-3020

[^39]: Pratt & Whitney Engines, West Palm Beach, FL 33410

[^40]: Center for Nondestructive Evaluation, Iowa State University,
    Ames, IA 50011

[^41]: Center for Nondestructive Evaluation, Iowa State University,
    Ames, IA 50011

[^42]: Center for Quality Engineering and Failure Prevention,
    Northwestern University, Evanston, IL 60208-3020

[^43]: Center for Quality Engineering and Failure Prevention,
    Northwestern University, Evanston, IL 60208-3020

[^44]: Center for Quality Engineering and Failure Prevention,
    Northwestern University, Evanston, IL 60208-3020

[^45]: Institute for Manufacturing Research and Department of Physics,
    Wayne State University, Detroit, MI 48202

[^46]: Department of Mechanical Engineering , Tuskegee University,
    Tuskegee, AL 36088

[^47]: Bell Helicopter Textron Inc., 600 E. Hurst Blvd., Fort Worth,
    Texas, 76101

[^48]: Petroleum Helicopters Inc., P.O. Box 90808, Lafayette, Louisiana,
    70509

[^49]: Bell Helicopter Textron Inc., 600 E. Hurst Blvd., Fort Worth,
    Texas, 76101

[^50]: Lehigh University Department of Mechanical Engineering and
    Mechanics, 19 Memorial Drive West, Bethlehem, PA 18015

[^51]: Aerotech Engineering & Research Corporation, 3125 W. 6th St.
    Suite C, Lawrence, KS 66049-3101

[^52]: National Institute of Standards and Technology Metallurgy
    Division, Gaithersburg, MD 20899

[^53]: Diffracto Ltd. Windsor, Ontario N8T 3B7

[^54]: Sandia National Laboratories, Albuquerque, New Mexico 87185

[^55]: Boeing Commercial Airplane Group, Seattle Wash. 98124

[^56]: Carnegie Mellon Research Institute, Pittsburgh, PA 15230-2950

[^57]: Northwestern University, 1936 Sheridan Road, Evanston, Illinois
    60208-4040

[^58]: Sandia National Labs, Albuquerque, NM 87185

[^59]: Northwestern University, Evanston, Illinois 60208-4040

[^60]: Galaxy Scientific Corporation, Egg Harbor Township, NJ 08234-5562

[^61]: University of Dayton Research Institute, Dayton Ohio 45469-0120

[^62]: University of Washington, Seattle WA. 98195

[^63]: Structures, Materials, and Propulsion Laboratory, Institute for
    Aerospace Research, Ottawa, Ontario, Canada

[^64]: Sandia National Laboratories, Albuquerque, NM 87185

[^65]: Delta Air Lines, Hartsfield Atlanta Airport, Atlanta, Ga. 30320

[^66]: National Institute for Aviation Research, Mechanical Engineering
    Department, Wichita State University, Wichita, KS 67260-0035

[^67]: William J. Hughes Technical Center, AAR-433, Atlantic City
    Airport, NJ 08405

[^68]: Sandia National Laboratories, P. O. Box 5800 Albuquerque, NM
    87185

[^69]: National Aerospace Laboratory, P. O. Box 90502 1006 BM Amsterdam
    The Netherlands

[^70]: Federal Aviation Administration, William J. Hughes Technical
    Center, Atlantic City International Airport, NJ 08405

[^71]: McDonnell Douglas Aerospace, 1510 Hughes Way, Long Beach, CA
    90810

[^72]: Department of Mechanical Engineering University of Utah, 3209
    Merrill Engineering Center Salt Lake City, UT 84112

[^73]: Pepin Associates, 15 Holly Street, Scarborough, ME 04074