$$TR Avt 275 All
$$TR Avt 275 All
$$TR Avt 275 All
ORGANIZATION ORGANIZATION
AC/323(AVT-275)TP/930 www.sto.nato.int
Final report on best practices and lessons learned for aircraft structural,
propulsion, and mechanical systems safety and reliability. The report
Annex, STO-TR-AVT-275-A, contains Chapters 10, 23, and 29.
AC/323(AVT-275)TP/930 www.sto.nato.int
Final report on best practices and lessons learned for aircraft structural,
propulsion, and mechanical systems safety and reliability. The report
Annex, STO-TR-AVT-275-A, contains Chapters 10, 23, and 29.
The NATO Science and Technology Organization
Science & Technology (S&T) in the NATO context is defined as the selective and rigorous generation and application of
state-of-the-art, validated knowledge for defence and security purposes. S&T activities embrace scientific research,
technology development, transition, application and field-testing, experimentation and a range of related scientific
activities that include systems engineering, operational research and analysis, synthesis, integration and validation of
knowledge derived through the scientific method.
In NATO, S&T is addressed using different business models, namely a collaborative business model where NATO
provides a forum where NATO Nations and partner Nations elect to use their national resources to define, conduct and
promote cooperative research and information exchange, and secondly an in-house delivery business model where S&T
activities are conducted in a NATO dedicated executive body, having its own personnel, capabilities and infrastructure.
The mission of the NATO Science & Technology Organization (STO) is to help position the Nations’ and NATO’s S&T
investments as a strategic enabler of the knowledge and technology advantage for the defence and security posture of
NATO Nations and partner Nations, by conducting and promoting S&T activities that augment and leverage the
capabilities and programmes of the Alliance, of the NATO Nations and the partner Nations, in support of NATO’s
objectives, and contributing to NATO’s ability to enable and influence security and defence related capability
development and threat mitigation in NATO Nations and partner Nations, in accordance with NATO policies.
The total spectrum of this collaborative effort is addressed by six Technical Panels who manage a wide range of
scientific research activities, a Group specialising in modelling and simulation, plus a Committee dedicated to
supporting the information management needs of the organization.
• AVT Applied Vehicle Technology Panel
• HFM Human Factors and Medicine Panel
• IST Information Systems Technology Panel
• NMSG NATO Modelling and Simulation Group
• SAS System Analysis and Studies Panel
• SCI Systems Concepts and Integration Panel
• SET Sensors and Electronics Technology Panel
These Panels and Group are the power-house of the collaborative model and are made up of national representatives as
well as recognised world-class scientists, engineers and information specialists. In addition to providing critical
technical oversight, they also provide a communication link to military users and other NATO bodies.
The scientific and technological work is carried out by Technical Teams, created under one or more of these eight
bodies, for specific research activities which have a defined duration. These research activities can take a variety of
forms, including Task Groups, Workshops, Symposia, Specialists’ Meetings, Lecture Series and Technical Courses.
The content of this publication has been reproduced directly from material supplied by STO or the authors.
ISBN 978-92-837-2255-7
Single copies of this publication or of a part of it may be made for individual use only by those organisations or
individuals in NATO Nations defined by the limitation notice printed on the front cover. The approval of the STO
Information Management Systems Branch is required for more than one copy to be made or an extract included in
another publication. Requests to do so should be sent to the address on the back cover.
ii STO-TR-AVT-275
Table of Contents
Page
STO-TR-AVT-275 iii
3.3 Aging Aircraft Programme Conduct Issues 3-4
3.3.1 Pan-Organisation Support 3-4
3.3.2 Systems Sampling 3-4
3.3.3 Identification of Target Systems and Zones with Limited A Priori 3-5
Information
3.3.4 System Performance Assessment 3-5
3.3.5 System Interconnection Issues 3-6
3.3.6 Risk-Based Reporting 3-6
3.4 Wider Airworthiness Issues 3-6
3.4.1 Electrical Wiring and Interconnection Systems (EWIS) Integrity 3-6
3.4.2 Fault Reporting and Corrective Action Systems 3-7
3.4.3 Configuration Control Systems 3-7
3.4.4 Corrosion of Systems Components 3-7
3.4.5 Polymer Degradation 3-8
3.5 Conclusions 3-8
3.6 Acknowledgements 3-8
3.7 References 3-9
iv STO-TR-AVT-275
Part 2: Aircraft Structural Systems 2-i
Chapter 6 – Airworthiness of Structures (United States) 6-1
6.1 Introduction 6-1
6.1.1 Initial Airworthiness 6-1
6.1.2 Continued Airworthiness for Operations Up To ~50% SLL 6-2
6.1.3 Continued Airworthiness for Operations Approaching SLL 6-2
6.1.4 Continued Airworthiness for Operations Beyond SLL 6-2
6.1.5 USAF Aircraft Structural Safety Experience 6-3
6.2 References 6-3
STO-TR-AVT-275 v
9.3.1 Continuing Airworthiness for Initial Operations 9-4
9.3.2 Continuing Airworthiness for Operations Approaching the SLL 9-4
9.3.3 Continuing Airworthiness for Operations Beyond Original SLL 9-4
9.4 Enhanced Teardowns 9-5
9.5 References 9-5
1 Chapter 10, “Enhanced Wing Teardown in Support of F-16 Structural Life Management” is located in “Continuing
Airworthiness of Aging Systems: Annex” (STO-TR-AVT-275-A).
vi STO-TR-AVT-275
Chapter 12 – SU-22 Fighter-Bomber Aircraft Service Life 12-1
Extension Programme Supported by Operational Load
Monitoring Implementation
12.1 Introduction 12-1
12.2 Service Life Extension Program for SU-22 12-1
12.3 OLM System Design and Implementation On SU-22UM3K 12-3
12.4 In-Flight Load Determination 12-4
12.5 Summary 12-6
12.6 References 12-6
STO-TR-AVT-275 vii
Chapter 16 – Engine Structural Integrity Program (ESIP) 16-1
16.1 Introduction 16-1
16.2 Program Requirements 16-1
16.2.1 A Defined Program 16-1
16.2.2 At Regular Intervals 16-1
16.2.3 Reassess and Document Actual Usage Conditions 16-2
16.2.4 Compare Actual Usage Against Design Usage Conditions 16-2
16.2.5 Determine Life Consumption Based on Actual Usage (of 16-2
Critical Components, as a Minimum) Accordingly
16.3 Other Requirements 16-3
16.3.1 Reliability Tracking 16-3
16.4 Outcomes 16-3
16.5 References 16-4
viii STO-TR-AVT-275
19.6 Continuing Airworthiness for USAF Mechanical Equipment 19-7
and Subsystems
19.7 Summary 19-7
19.8 References 19-7
2 Chapter 23, “Life Extension of Tornado Aircraft Subsystems (Germany)” is located in “Continuing Airworthiness of Aging
Systems: Annex” (STO-TR-AVT-275-A).
STO-TR-AVT-275 ix
23.4 Tornado Life Extension 23-6
23.4.1 Introduction 23-6
23.4.2 The Life Extension Certification Task 23-6
23.5 Applying the Four Phase Process for Certification 23-7
23.5.1 Phase A 23-7
23.5.1.1 Equipment 23-7
23.5.1.2 Supporting Structure 23-9
23.5.1.3 Panavia Standard Parts 23-9
23.5.2 Phase B 23-10
23.5.3 Phase C 23-11
23.5.4 Phase D 23-12
23.6 Examples of Work Performed During Phase C 23-12
23.6.1 Fuel System 23-12
23.6.1.1 System Schematic/Description and Work Breakdown 23-12
23.6.1.2 Interfaces 23-15
23.6.1.3 Failure Modes 23-15
23.6.1.4 Life Extension Measures 23-16
23.6.2 Secondary Power System (SPS) 23-25
23.6.2.1 System Schematic/Description and Work 23-25
Breakdown
23.6.2.2 Interfaces 23-26
23.6.2.3 Failure Modes 23-26
23.6.2.4 Life Extension Measures 23-27
23.7 Summary/Outlook 23-32
23.8 References 23-33
x STO-TR-AVT-275
Chapter 26 – Su-22 FSFT Study Case and Mi-8 Hard Landing 26-1
Test Study Case
26.1 Swept-Wing Jet Aircraft FSFT 26-1
26.1.1 Introduction 26-1
26.2 Service Life Extension Program 26-1
26.2.1 Summary of the Su-22 FSFT 26-3
26.3 Military Transport Helicopter Harsh Landing Test 26-3
26.3.1 Introduction to Transport the Helicopter Harsh Landing Case 26-3
26.3.2 Problem of Harsh Landings 26-4
26.3.3 Test Setup 26-4
26.3.4 Instrumentation 26-5
26.3.5 Release Mechanism 26-6
26.3.6 Drop Test Assumptions 26-6
26.3.7 Drop Test Execution 26-6
26.3.8 Test Summary 26-8
26.4 References 26-9
3 Chapter 29, “Corrosion Management Using Environmental Sensor Data” is located in “Continuing Airworthiness of Aging
Systems: Annex” (STO-TR-AVT-275-A).
STO-TR-AVT-275 xi
29.3 Measurement of Corrosivity of the Environment 29-2
29.3.1 Measurement Approach 29-2
29.3.2 Sensor Suite 29-2
29.3.3 Measurement Setup 29-3
29.4 Measurement Locations 29-3
29.4.1 Airbase 29-3
29.4.2 Ship 29-3
29.4.3 Helicopter 29-4
29.5 Example of Re-Scheduling of Recommended Corrosion Inspection 29-4
29.6 Corrosivity Measurement Results 29-5
29.6.1 In-Service Measurement 29-5
29.6.2 Sensor Data and Corrosion Observations 29-7
29.6.3 Surface Contamination Effect 29-7
29.6.4 Conclusions from In-Service Measurements 29-8
29.7 Modelling of Corrosivity 29-8
29.8 Corrosion Modelling Results 29-9
29.8.1 Fitting of the Corrosion Model 29-9
29.8.2 Modelling Insights 29-10
29.9 Implications for Maintenance 29-11
29.9.1 Relative Humidity During Storage 29-11
29.9.2 Use of Dry Air Units 29-11
29.9.3 Sensor Data for Corrosion Maintenance Scheduling 29-11
29.10 Conclusions for Continued Airworthiness 29-11
29.11 References 29-12
xii STO-TR-AVT-275
List of Figures
Figure Page
Figure 7-1 Schematic for Projecting Crack Size Distribution as a Function 7-1
of Flight Hours
Figure 7-2 POD Sensitivity Analysis for a Fatigue Critical Location 7-8
Figure 10-1 Location of the Five IAT Strain Sensors in the F-16 of 10-2
the RNLAF
Figure 10-2 Experimental Validation of the CSI Concept for the Lower 10-3
Wing Skin at BL 120
Figure 10-3 One of the Test Specimens Used in the Assessment of Digital 10-5
X-ray Application for Detecting Cracked F-16 Pre-Block
40 Wing Spars
Figure 10-4 Assessment of the Capability to Detect Cracked F-16 10-5
Pre-Block 40 Wing Spars Using One-Step Process Digital
X-Ray Equipment at a Tilt Angle of 30°
Figure 10-5 Test Article and Test Setup Used in the Comparative 10-7
Test Program
Figure 10-6 Examples of the Varying Installation Quality of the Two 10-7
Fastener Systems
Figure 10-7 Workflow of the Enhanced F 16 Block 15 Wing Teardown 10-11
Program
Figure 10-8 Schematic of Crack Opening Procedure 10-11
Figure 10-9 Root Area of Lower Wing Skin with Several NDI Indications 10-12
Figure 10-10 Crack Findings in RHS Wing at 4,200 FH 10-12
STO-TR-AVT-275 xiii
Figure 10-11 Hoisting of the Test Article Prior to Installation in the Test Rig 10-15
Figure 10-12 Wing Suspension in the Test Rig 10-15
Figure 10-13 Impression of the WDET Setup 10-16
Figure 10-14 Outline of the Process to Compute the Hydraulic Actuator 10-17
Loads Spectrum
Figure 10-15 Validation of Flight Loads Computation Process 10-17
Figure 10-16 The Computed WDET Load Spectrum vs. the LM Generated 10-18
Spectra for Two Different RNLAF Usage Spectra
Figure 10-17 Observed Fracture Surface Pattern for Marker Load 10-19
Sequence #3
Figure 10-18 Composite Photograph that Shows the Wing Deflections 10-20
During the Applied Up and Down Bending Limit Load Cases
Figure 10-19 Multi-Site Damage in Lower Wing Skin at BL 71.00 10-21
Rectangular Cutout
Figure 10-20 Crack Findings in LHS Wing at 18,200 FH 10-22
Figure 10-21 Fatigue Cracks Selected for Quantitative Fractography 10-24
Figure 10-22 Fractographic Analysis of the Lead Crack at the BL 71 10-25
Rectangular Cutout in the Lower Skin
Figure 10-23 Crack Growth Curves Derived from the Markers on the 10-25
Fracture Surfaces
Figure 10-24 Use of the Lead Crack Fatigue Lifing Framework to 10-27
Establish the Lower Limit on the Economic Service Life of
the RNLAF F-16 Block 15 Wing Box
Figure 10-25 Required Stochastic and Deterministic Input Data for a 10-29
Structural Risk Analysis
xiv STO-TR-AVT-275
Figure 18-1 Probability of Failure in the Safe Life Approach 18-6
Figure 18-2 Schematic Representation of the Damage Tolerance Based Life 18-7
Cycle Management Approach
Figure 18-3 Flow Diagram of the Damage Tolerance Assessment Algorithm 18-8
Figure 19-1 Steps Required for Completing the Five MECSIP Tasks 19-4
STO-TR-AVT-275 xv
Figure 25-7 A Scheme of MiG-29 Aircraft 25-5
Figure 25-8 Location of a Selected Strain Gauge Used for Loads 25-6
Monitoring of MiG-29 Vertical Stabilizer
Figure 25-9 Example of Strain Gauge of OLM System 25-7
Figure 25-10 Examples of Normalized Strain Records from Strain 25-7
Gauge and Canonical Flight Parameter for Steady and
High Manoeuvre Flights
Figure 25-11 Example of CFD Computation for a Particular Flight 25-11
Configuration
Figure 25-12 Example of the Aerodynamic Force Distribution on the 25-12
Surface of the Stabilizer Based on CFD Computations
Figure 25-13 Vertical Stabilizer of the Aircraft with Indication of the 25-13
Axis with Respect to which Bending Momentum is Computed
Figure 25-14 Examples of Normalized Records of Strain Gauge, Canonical 25-14
Flight Parameter and Bending Momentum for Steady and
High Manoeuvre Flights
xvi STO-TR-AVT-275
Figure 28-6 Flaw Indication and Temperature Profile Plotted for Blue Line 28-4
Figure 28-7 Adhesive Layer After Patch Removing, with Evidence of 28-4
Voids and Areas Without Bonding
Figure 28-8 Inspection Using MOI 308 TDF 28-5
Figure 28-9 Thermography Result 28-5
STO-TR-AVT-275 xvii
Figure 31-7 Repair Technology of Cracks in a Skin’s Section and 31-6
Puncture of a Skin’s Sections
Figure 31-8 Repair Technology of Dented Trailing Section 31-6
Figure 31-9 The Section with Repairs 31-6
Figure 31-10 The Drying Equipment of the Structure 31-7
Figure 31-11 Water – Boiling Points at Vacuum Pressure 31-7
Figure 31-12 Surface Preparation for Bonding and Type of Failure 31-8
During Wedge Test (ASTM D3762) of the Joint Subjected
to 100% Humidity Exposure
Figure 31-13 Example of the Cure Cycle Curve 31-8
Figure 31-14 Schema of Device and Vacuum Bag During Cure Cycle of Patch 31-9
xviii STO-TR-AVT-275
List of Tables
Table Page
Table 18-1 Failure Modes and Life-Limiting Properties for Turbine 18-1
Engine Components
STO-TR-AVT-275 xix
Table 27-1 Information about NDT Techniques Used and Their 27-1
Capabilities
Table 29-1 Sensors and Measurement Techniques Used in the LS2A 29-3
Sensor Suite
Table 30-1 Information about MRB Used in Polish Air Force 30-1
xx STO-TR-AVT-275
List of Acronyms
AELEV-MIL Aero-elastic simulation tool for the lifing of military aircraft (Dutch acronym)
ASIP Aircraft Structural Integrity Program
BL Butt Line
EC Eddy Current
EIDS Equivalent Initial Damage Size
EPAF European Participating Air Forces
EPS Equivalent Pre-Crack Size
EOL End Of Life
g gravity
GE General Electric Measurement and Control
STO-TR-AVT-275 xxi
NASM National Aerospace Standard/Military
NDI Non-Destructive Inspection
NI Neutral Interface
NLR Royal Netherlands Aerospace Centre
Nz Normal acceleration (vertical)
QF Quantitative Fractography
xxii STO-TR-AVT-275
Acknowledgements
Thanks to Marko Yanishevsky for helping with technical editing on the chapters initially received from the
authors.
STO-TR-AVT-275 xxiii
AVT-275 Membership List
CO-CHAIRS
Dr. Min LIAO* Dr. Kimberli JONES*
National Research Council Canada United States Air Force
CANADA UNITED STATES
Email: Min.Liao@nrc-cnrc.gc.ca Email: kimberli.jones.1@us.af.mil
MEMBERS
Mr. Wolfgang BIENENDA* Mr. Marcin KURDELSKI*
Airbus Air Force Institute of Technology
GERMANY POLAND
Email: wolfgang.bienenda@airbus.com Email: marcin.kurdelski@itwl.pl
ADDITIONAL CONTRIBUTORS
Mr. Charles BABISH Mr. Wieslaw BERES
United States Air Force National Research Council Canada
UNITED STATES CANADA
Email: charles.babish@us.af.mil Email:Wieslaw.Beres@nrc-cnrc.gc.ca
4* Contributing Author
xxiv STO-TR-AVT-275
Mr. Yvan CARON Major Sean LEITHEAD
Department of National Defence (DND) Royal Canadian Air Force
CANADA CANADA
Email: YVAN.CARON@forces.gc.ca Email: Sean.Leithead@forces.gc.ca
STO-TR-AVT-275 xxv
Captain Sergiusz SZAWŁOWSKI (ret.) Col. Janusz ZAWISLAK
Air Force Institute of Technology Inspectorate for Armed Forces Support
POLAND POLAND
Email: sergiusz.szawlowski@itwl.pl Email: j.zawislak@ron.mil.pl
PANEL/GROUP MENTOR
Mr. Jerzy KOMOROWSKI
NRC Aerospace (retired)
CANADA
Email: jerzy.komorowski@jpwkaero.com
TECHNICAL EVALUATOR
Mr. Charles BABISH
United States Air Force
UNITED STATES
Email: charles.babish@us.af.mil
xxvi STO-TR-AVT-275
Continuing Airworthiness of Aging Systems
(STO-TR-AVT-275)
Executive Summary
Due to high procurement costs associated with replacing aging aircraft fleets, NATO Nations are frequently
required to operate their aircraft for longer than the original design life. Because of the extended life
requirements and the fact that aircraft have likely flown more severely than designed, aging aircraft issues
are a significant problem when it comes to maintaining the airworthiness of these aircraft. Given the unique
situation each NATO nation experiences because of the different operational environments, maintenance
procedures and flight envelopes, an opportunity exists for the various nations to share their best practices
relating to ensuring the airworthiness of aging aircraft systems within this Task Group.
A workshop, AVT-222, was held in October 2015 at which time various NATO nations made presentations
on how they were ensuring the safety of aging aircraft (airframe and systems). The results from the
workshop highlighted the aging aircraft issues, which the nations are experiencing, that are seriously
impacting aircraft airworthiness and availability. At the conclusion of the workshop, it was agreed that a
focus on common maintenance airworthiness issues would help in strengthening the Maintenance
Organizations’ ability to meet airworthiness requirements. AVT-275 was created to fulfil this role.
The objective of this Task Group was to develop a technical report containing the best practices and lessons
learned that exist in NATO nations for aircraft structural, propulsion, and mechanical systems.
The documented best practices on continuing airworthiness of aging aircraft systems aim at capturing the
unique aptitudes that have developed in each participating nation. With these documented practices, the
different NATO nations will be able to adopt them as they see appropriate. This has the potential to decrease
the cost of maintaining aging aircraft systems since the procedures/processes have already been developed
by other nations.
The primary motivation is that NATO Nations have similar methods (design criteria, analysis, testing, usage
monitoring, aircraft damage surveillance, etc.) and results (acceptable mishap rates and tolerable experience
associated with discovering unanticipated issues that must be corrected) for ensuring continuing
airworthiness for the aircraft structure; but that there is a diverse approach with mixed results for
safety-critical systems in various NATO military aircraft. Conversely, key differences in how some Nations
sustain aircraft may lead to modifications in what should be done for other safety-critical systems, for
example safe-life approach, and factors used for safety-by-inspection.
Aircraft structural systems have the most developed integrity program, with Propulsion systems well-defined
and becoming matured. Mechanical systems are progressing, but many lessons from structural and
propulsion integrity programs could be used to further mechanical systems integrity program efforts.
STO-TR-AVT-275 ES - 1
Employing aircraft structures-like methods may improve the overall aircraft continued airworthiness,
especially at or beyond the original service life limit.
ES - 2 STO-TR-AVT-275
Navigabilité continue des systèmes vieillissants
(STO-TR-AVT-275)
Synthèse
Le coût d’acquisition élevé associé au remplacement des parcs d’aéronefs vieillissants oblige fréquemment
les pays de l’OTAN à exploiter leurs aéronefs plus longtemps que la durée de vie théorique. À cause
du prolongement de la durée d’exploitation et du fait que les appareils ont probablement connu des
conditions de vol plus difficiles que prévu, les problèmes des aéronefs vieillissants sont un sujet important
dès lors qu’il s’agit de maintenir la navigabilité. Étant donné la situation unique de chaque pays de l’OTAN,
en raison des différences d’environnements opérationnels, de procédures de maintenance et de domaines de
vol, les divers pays ont l’opportunité de partager au sein de ce groupe de travail leurs meilleures pratiques
relatives au maintien de la navigabilité des systèmes d’aéronefs vieillissants.
Un séminaire, l’AVT-222, s’est tenu en octobre 2015, pendant lequel divers pays de l’OTAN ont fait
des présentations sur leur façon d’assurer la sécurité des appareils vieillissants (cellules et systèmes).
Les résultats du séminaire ont mis en lumière les problèmes que les pays rencontrent avec les aéronefs
vieillissants et qui entament gravement la navigabilité et la disponibilité des appareils. Lors de la conclusion
du séminaire, il a été convenu qu’il serait utile de faire le point sur les problèmes courants d’entretien
de la navigabilité, pour renforcer la capacité des organisations de maintenance à respecter les exigences
de navigabilité. L’AVT-275 a été créé dans ce but.
L’objectif de ce groupe de travail était de rédiger un rapport technique contenant les meilleures pratiques
et les enseignements des pays de l’OTAN en matière de systèmes de propulsion et de systèmes structurels
et mécaniques. Les meilleures pratiques documentées sur la navigabilité continue des systèmes d’aéronefs
vieillissants visent à rendre compte des aptitudes uniques développées par chaque pays participant.
Les différents pays de l’OTAN seront ainsi en mesure d’adopter ces pratiques s’ils les jugent adéquates. Cela
est susceptible de réduire le coût de maintenance des systèmes d’aéronefs vieillissants, puisque
les procédures/processus ont déjà été élaborés par d’autres pays.
La principale motivation est que les pays de l’OTAN ont des méthodes (critères de conception, analyse,
essais, suivi de l’usage, surveillance des avaries, etc.) et des résultats (taux d’accident admissibles
et expérience tolérable associée à la découverte de problèmes non anticipés qui doivent être corrigés)
similaires pour assurer la navigabilité continue de la structure de l’aéronef, mais que l’approche des systèmes
critiques sur le plan de la sécurité diffère, avec des résultats mitigés dans divers aéronefs militaires
de l’OTAN. À l’inverse, des différences essentielles d’entretien des aéronefs dans certains pays pourraient
modifier les interventions à mener sur d’autres systèmes critiques sur le plan de la sécurité, par exemple
la démarche de vie sûre et les facteurs utilisés pour la tolérance de détérioration.
STO-TR-AVT-275 ES - 3
Les systèmes structurels des aéronefs disposent du programme d’intégrité le plus développé, avec
des systèmes de propulsion bien définis et arrivant à maturité. Les systèmes mécaniques progressent, mais
de nombreux enseignements des programmes d’intégrité structurelle et de la propulsion pourraient faciliter
les travaux des programmes d’intégrité des systèmes mécaniques. L’utilisation de méthodes semblables
à celles appliquées aux structures d’aéronefs pourrait améliorer la navigabilité continue globale des aéronefs,
en particulier à la fin de la durée de vie théorique ou au-delà.
ES - 4 STO-TR-AVT-275
Part 1: CONTINUING AIRWORTHINESS
POLICIES AND APPROACHES
STO-TR-AVT-275 Part 1 - i
Part 1 - ii STO-TR-AVT-275
Chapter 1 – CONTINUING AIRWORTHINESS
POLICIES (UNITED STATES)
Charles A. Babish IV
Air Force Life Cycle Management Center
UNITED STATES
The USAF airworthiness process requires a continual evaluation of fielded aircraft throughout their life cycle
to ensure that all aircraft are in a condition for safe operation, i.e., that they have not exceeded the approved
service life limit, that environmental (e.g., corrosion) or other factors have not degraded airworthiness, that
aircraft are properly maintained in accordance with approved maintenance documentation, and the aircraft
are operated in accordance with the approved flight manual within the approved mission usage. The USAF
recognizes that the initial service life limit is not a fixed value and the airworthiness process includes
procedures for determining service life extension potential and for executing a service life extension program
when desired and practical [4].
STO-TR-AVT-275 1-1
CONTINUING AIRWORTHINESS POLICIES (UNITED STATES)
1.2 REFERENCES
[1] Air Force Instruction 63-101, “Acquisition and Sustainment Life Cycle Management”, May 2017.
[4] AWB-1009, USAF Airworthiness Bulletin 1009, “Airworthiness Flight Authorizations – Military Type
Certificate (MTC)/Military Flight Release (MFR)”, March 2016.
[5] AWB-150, USAF Airworthiness Bulletin 150, “Airworthiness (AW) Risk Assessment and
Acceptance”, September 2017.
1-2 STO-TR-AVT-275
Chapter 2 – CONTINUING AIRWORTHINESS PROGRAM
OF THE ROYAL CANADIAN AIR FORCE (RCAF)
2.1 INTRODUCTION
This chapter presents a brief summary of the Royal Canadian Air Force (RCAF) continuing airworthiness
(CAW) program including the major/unique features of the RCAF CAW policy/approach. It also provides
the latest RCAF references related to the RCAF CAW program, as well as the lessons learned and future
challenges/gaps.
Under the provisions of the Aeronautics Act, the Canadian Minister of National Defence (MND) is
responsible for military aviation, including foreign aircraft within Canada. Under the Aeronautics Act, the
MND delegates powers and responsibility to the Airworthiness Authority (AA), Technical Airworthiness
Authority (TAA), Operational Airworthiness Authority (OAA), and Airworthiness Investigative Authority
(AIA). Among them, the TAA is responsible for the regulation of the technical airworthiness aspects of
design, manufacture, maintenance and material support of aeronautical products and the determination of the
airworthiness acceptability of those products prior to operational service. The technical airworthiness
program, which is outlined in the Technical Airworthiness Manual (TAM) [1], consists of three primary
elements: Initial Airworthiness, Continuing Airworthiness, and Disposal. Initial Airworthiness includes
the processes and activities necessary to ensure that aeronautical products are airworthy (including
Type Certification and Equipment Certification), prior to their introduction into service.
STO-TR-AVT-275 2-1
CONTINUING AIRWORTHINESS PROGRAM
OF THE ROYAL CANADIAN AIR FORCE (RCAF)
Under the aircraft usage and condition monitoring, the TAM stipulates that there shall be processes for
monitoring the condition of the Aircraft Structure, Mechanical Systems, Propulsion Systems and Aircraft
Electrical Wiring Interconnection Systems. While the requirement to monitor the condition of the Aircraft
Structure has been included in the TAM from the beginning, it was only in November 2015, with the release
of Change 7 of the TAM, were Annexes for monitoring of Mechanical Systems, Propulsion Systems and
Aircraft Electrical Wiring Interconnection Systems populated.
The requirement for an Aging Aircraft Assessment (AAA) was initially introduced under the aircraft usage
and condition monitoring requirement in June 2012 with release Change 6 of the TAM. At that time it was to
apply to aircraft structures and systems. However, at Change 6 of the TAM, the Annexes describing the
monitoring requirement for the Mechanical Systems, Propulsion Systems, and Aircraft Electrical Wiring
Interconnection Systems had not been populated. As all of the AAAs activities at DND were related to the
aircraft structure, Change 7 of the TAM only included the detail requirements specified for an Aging Aircraft
Structural Assessment (AASA).
The AASA is to be carried out no later than 15 years after the production of the first aircraft delivered to the
RCAF. As of Dec 2017, all of the RCAF fleets that have passed the 15 year mark have undergone an AASA.
The approaches taken to perform the AASA has differed based on the fleet and the Aircraft Structural
Integrity Program (ASIP) for the fleet. For the fleets that pronominally follow a civilian role, such as the
CC144 (Bombardier Challenger CL601/604 variants) and CC150 (Airbus A310), AASA relied heavily on
the work done by the Original Equipment Manufacturer (OEM) under the Federal Aviation Administration
(FAA) mandated program for Aging Airplane Safety [2]. For the predominately military role fleets, a
combination on physical audit and review of the structural maintenance records were performed as part of
the AASA.
The level of effort to perform the AASAs was largely dependent on the strength of the ASIP for the fleet in
question. Fleets that had a strong ASIP required little effort to perform the AASA as there was already very
good documentation on the actual usage and physical condition of the airframe. Overall, the RCAF fleets are
well maintained and no significant issues were raised during the AASAs. This is largely due to having
healthy ASIPs for the RCAF fleets. This is not to say that some of the aircraft did not have corrosion issues,
but these were already being monitored and addressed as part of the ASIP.
2-2 STO-TR-AVT-275
CONTINUING AIRWORTHINESS PROGRAM
OF THE ROYAL CANADIAN AIR FORCE (RCAF)
The RCAF RARM process includes five steps for risk management, i.e., Hazard Identification, Risk
Assessment, Risk Control, RARM Approval, and Risk Tracking [1], which are briefly defined as:
1) Hazard Identification. Identifies the potential mishaps to an aeronautical product that may arise from
a particular root cause in conjunction with possible hazard conditions;
2) Risk Assessment. Identifies the level of risk by taking into consideration the severity and probability
of occurrence of the potential hazard;
3) Risk Control. Identifies the options that are available to eliminate, reduce and/or mitigate the
assessed risk and the rationale for the selection of the recommended options;
4) RARM Approval. Identifies the requirement for specific authorized individuals to approve a RARM
dependent on the level of risk identified. Describes the approval responsibilities, which include
verifying the technical airworthiness content and ensuring that risk management procedures were
followed; and
5) Risk Tracking. Identifies measures to ensure the RARM is appropriately documented and that risk
control measures are implemented and monitored for effectiveness.
In the RARM, a key airworthiness risk index, defined as in Figure 2-1, is measured by a product combining
both Hazard Severity and Hazard Probability. The risk index definitions are aligned with other military and
civil standards including FAA/JAR 25-1309, AC 25-1309, SAE ARP 4761, and MIL-STD-882. Detailed
definitions for the Hazard Severity, i.e., Catastrophic (A), Hazardous (Severe Major) (B), Major (C), Minor
(D), and Negligible (E), are documented in Ref. [1]. Detailed definitions for qualitative Hazard Probability
(Probability of Occurrence), i.e., Frequent (Level 1), Probable (Level 2), Remote (Level 3), Extremely Remote
(Level 4), and Extremely Improbable (Level 5), are also documented in Ref. [1]. The corresponding
quantitative Hazard Probability Thresholds (HPT) per flight hour, such as 10-3, 10-5, 10-7, or 10-9, are
presented in Figure 2-2. Note that the HPT defined in Figure 2-2 also depends on the type of RCAF aircraft
platforms, such that an aircraft carrying passengers will require lower HPT than an unmanned air vehicle.
The RARM process has affected all RCAF fleets (DND-AD-2007-01), covering all the airworthiness and
safety-related aircraft systems including hydraulics, structural, mechanical, avionics, etc. [3]. Today, the
RARM has become the single most critical decision-making tool for the RCAF air fleet management [4].
When there are sufficient data available, a quantitative risk assessment can be performed to substantiate the
assignment of a risk index number. When a qualitative risk assessment indicates a high or medium risk, a
detailed quantitative risk assessment is often requested to calculate the hazard probability, i.e., the single
flying hour probability of failure, to gain additional confidence in the decision making. In practice, the
RCAF’s first intention is always to conduct a quantitative risk assessment; when there are not enough data, a
qualitative risk assessment is performed supported by engineering judgement [4].
STO-TR-AVT-275 2-3
CONTINUING AIRWORTHINESS PROGRAM
OF THE ROYAL CANADIAN AIR FORCE (RCAF)
2-4 STO-TR-AVT-275
CONTINUING AIRWORTHINESS PROGRAM
OF THE ROYAL CANADIAN AIR FORCE (RCAF)
Since the introduction of RARM into TAA process in 2001, it is shown that RARM is an effective tool for
managing the RCAF fleets continuing airworthiness. A recent NRC work on RARM data mining has been
investigating some statistics of the RARM database, including number of RARMs for all fleets and specific
fleet; risk level/index (low, med, high) distributions; RARM cases distributions for different aircraft systems;
RARM hazard cause/effect distribution; risk assessment technology solutions; qualitative vs. quantitative
risk assessment. So far the study shows: there are about 1040 RARM cases for all fleets, including both
closed and still open. Some RARMs have multiple versions and the risk index was revised/lowered in the
final versions. As a snapshot in 2018, about half of them are ALOS (acceptable level of safety), 6.5% Low,
38.3% Medium, and about 5.9% High and Extremely High (Figure 2-3(a) and Figure 2-3(b)) shows the
RARM case distribution for different aircraft components/systems, it is noted that about 65% of RARMs are
for “other” systems which mainly include avionics and software, etc.
(a) RARM Risk Index Distribution. (b) RARM Cases Distribution for Different Systems.
2.8 REFERENCES
[1] Department of National Defence of Canada, Technical Airworthiness Manual (TAM), Document No. C-
05-005-001/AG-001, 2015, Change 7, 23 November 2015.
[2] US Department of Transportation, Federal Aviation Administration, 14 CFR Parts 119, 121, 129, 135,
and 183, Aging Airplane Safety; Final Rule, 2 February 2005.
[3] Department of National Defence of Canada (August 2007), “Review of Open Records of Airworthiness
Risk Management”, Airworthiness Directive DND-AD-2007-01.
STO-TR-AVT-275 2-5
CONTINUING AIRWORTHINESS PROGRAM
OF THE ROYAL CANADIAN AIR FORCE (RCAF)
[4] Liao, M., Kainth, T., Renaud, G., and Bombardier, Y., “Quantitative Risk Analysis to Support
Continued Airworthiness Management for Aircraft Structures”, NATO-RTO-AVT-222, Brussels,
Belgium, October 2014.
2-6 STO-TR-AVT-275
Chapter 3 – LESSONS FROM AGING
AIRCRAFT SYSTEMS AUDITS
Steve Reed
Technical Fellow, Dstl, Ministry of Defence
UNITED KINGDOM
3.1 INTRODUCTION
The United Kingdom Ministry of Defence (UK MOD) has been undertaking Aging Aircraft Structural
Audits (AASA) since the early 1990s. These audits were initiated following the well-known Aloha Flight 243,
Boeing 737 pressure cabin failure in 1988 [1]. However, several civil accidents in the mid-1990s, including
TWA Flight 800 [2] and Swiss Air Flight 111 [3], made the aviation community aware of the need to consider
aging effects of electrical and mechanical systems alongside structural implications. For the UK MOD the need
to understand these aging systems issues was reinforced starkly by the loss of Nimrod XV230 over Afghanistan
in 2006. Thereafter, the inclusion of aircraft systems aspects into the UK MOD’s Aging Aircraft Audit (AAA)
process was introduced.
The UK MOD / Industry Systems Airworthiness Advisory Group (SAAG) and Aging Aircraft Programme
Working Group (AAPWG) were established, chaired by the UK MOD Military Aviation Authority (MAA)
to sit alongside the existing structures and propulsion MOD / Industry groups. One of the early outputs from
the SAAG was SAAG Paper 001 [4], which focused upon initial lessons identified from these early AAA
including systems elements (AASysA), Condition Survey (CS) and other associated integrity programmes.
This chapter is primarily a precis of SAAG Paper 001 focusing on the critical issues, with updated references
identified where appropriate. Readers wishing to gain a fuller picture of these issues should read SAAG
Paper 001, which is freely available on the UK MOD website [4].
Lessons from these initial AASysA and CS, including programmes conducted on the Nimrod, VC10, C130,
Sentry, Tucano and Historic Aircraft were collated, with gratefully acknowledged inputs from across the
UK MOD and industrial support base. Common recommendations were identified under the categories of:
regulatory or policy issues, programme conduct issues and wider airworthiness issues, and the major issues
are discussed in the following sections of this chapter. The key recommendations from this study were:
• To include a requirement to validate the effectiveness of integrity assurance methods by undertaking
a detailed physical condition survey of a sample of aircraft from within each fleet;
• To reinforce the airworthiness requirements of the AAA, rather than the cost of ownership and
availability issues;
• To include a requirement to address the effectiveness of emergency systems within the AAA; and
• To streamline the three independent audits (structures, systems and propulsion) into a single AAA.
All four of these key recommendations have since been incorporated into the UK MOD Regulatory
Framework [5]. Moreover, the recommendations discussed below and in SAAG Paper 001 have been used
as the foundation for the UK MOD Aging Aircraft Research and Development Programme since the
publication of the paper.
STO-TR-AVT-275 3-1
LESSONS FROM AGING AIRCRAFT SYSTEMS AUDITS
3-2 STO-TR-AVT-275
LESSONS FROM AGING AIRCRAFT SYSTEMS AUDITS
maintained by the MAA. These papers when then referenced from the appropriate sections of the regulatory
framework. This ensured that papers had suitable credibility, were endorsed by the relevant authorities, were
easily accessible and had longevity. SAAG Paper 001 was the first non-structures paper developed using this
approach. Other GM followed soon thereafter, including Refs. [7], [8], [9], [10]. This has proven to be an
extremely successful method of knowledge capture, retention and continuous improvement.
Consequently, the regulatory framework was amended to specify that a CS of a sample of the fleet was
required to assess the performance of the measures used to assure an airworthy aircraft and to identify signs
of aging (i.e., degradation) [5]. In addition, the MOD formed a MOD / Industry Condition Survey Working
Group (CSWG) (a SAAG sub-committee) within which SMEs developed guidance material on the conduct
of CS [10], now referenced from the AAA regulations [5].
Ideally, at the initiation of an AAA existing information would be readily available to identify all the known
system hazards and inter-system hazards across the aircraft, from a detailed functional safety, operational
safety and physical safety perspective. This would then form the basis of the AAA approach. However, for
legacy aircraft fleets in particular, the available systems safety information and its source were found to vary
considerably. Enhancements to the regulations [5] were recommended to emphasise the importance of
clearly defining the scope of the AAA and the rationale for exclusions.
Therefore, it was considered that the continued performance and adequacy of each emergency system should
be ascertained against the system requirements. Overheat detection devices provide an excellent example to
STO-TR-AVT-275 3-3
LESSONS FROM AGING AIRCRAFT SYSTEMS AUDITS
illustrate this point. Evidence identified that although the overheat detectors on some platforms may have
been subject to electrical continuity checks, the actual performance of the detector had not been verified
during the life of the aircraft, particularly if these sensors had been identified as ‘on condition’ items with no
refurbishment or hard-life requirements.
Following these recommendations, the regulatory article for AAA [5] was amended to ensure that
emergency systems were subject to specific attention within an AAA. Moreover, the systems integrity
management regulations [6] were also amended to ensure that sensor-related emergency systems had an
appropriate calibration policy in place.
However, experience identified shortfalls in the information readily available to validate these systems
design and usage assumptions. It is important to understand the key sensitivities within a target system or
zone and to be able to measure appropriate parameters to validate assumptions. For example, when
considering the traditional fire triangle of oxygen, heat or ignition source, and fuel, the usual focus within a
zone for an ignition source might be a spark from an electrical system. However, auto-ignition of fuels can
occur without the presence of a flame or spark. The exact conditions for auto-ignition are influenced by
factors such as surface geometry, closed or open environment, local air velocities and fluid residence time
[16]. Nevertheless, aviation fuels are vulnerable to auto-ignition at temperatures as low as 210 °C to 300 °C,
depending on the air-fuel ratio, at atmospheric pressure. Again it was considered essential to assemble
guidance in how to assess the zonal hazards and this led to the development of guidance [17], which is now
referenced from the systems integrity regulations [6].
3-4 STO-TR-AVT-275
LESSONS FROM AGING AIRCRAFT SYSTEMS AUDITS
understand in-depth issues. However, the activity of sampling does not necessarily require a full aircraft, nor
does it necessarily have to result in the destruction of all or part of the aircraft. Additionally, the sample size,
with an aim to undertake sufficient samples to be representative of the fleet, should be considered [18], [19].
Evidence illustrated that with careful planning it was possible for the necessary systems testing, analysis and
inspection to be largely accommodated within the realms of existing maintenance programmes. The
additional burden of sampling beyond the normal maintenance regime into un-inspected or undisturbed
systems would generally be expected to be relatively small.
Sampling has been identified in MOD structures policy and regulation for many years. However, it is only in
relatively recent years that the true value has been recognised in the wider aerospace community as structural
sampling has providing crucial evidence to continue operating aging aircraft fleets. As the focus has shifted
towards the integrity of systems and lives of aircraft continue to be extended, it is likely that sampling of
systems will also prove to be an invaluable tool. It should be remembered that the use of sampling, within an
age exploration programme [20] is a fundamental aspect of Reliability Centred Maintenance (RCM).
However, systems sampling programmes are still relatively rare. Consequently, when events do occur that
necessitate investigations of systems lifing policy, such as studies conducted into lifing of seals [21], flexible
hoses [22] and control rods [23], definitive evidence of the specific performance of these components can be
difficult to ascertain purely from component replacement or flight safety incident/occurrence report data;
hence sampling programmes are often necessary to supplement the available data. Following this study,
recommendations were accepted to emphasise the potential value and importance of sampling of system
components within the relevant regulations [5], [6].
3.3.3 Identification of Target Systems and Zones with Limited A Priori Information
As already discussed, for some of the UK MOD’s fleets the systems safety information available at the time
was limited and hence a method of identifying where to focus efforts was developed. SAAG Paper 001 [4]
describes the approach employed for the VC10 programme, which used an initial 3-day workshop, facilitated
by a highly-experienced design airworthiness engineer and supported by the UK MOD, design engineers,
maintenance engineers, maintenance instructors, in-service engineering investigators and experienced
aircrew to develop an initial Fault-Tree Analysis (FTA) using a Hull-Loss Model, as a template to ensure
that faults that could produce hazards and single-failure points were identified.
The success of this approach was highly dependent upon having knowledgeable representation from
designers, maintainers and operators (aircrew). Each of these organisations brought an essential element of
understanding and perspective; without any one of the parties the picture would have been incomplete and
the understanding of the implications of a potential systems failure could have been flawed. Consideration of
this approach was recommended in the absence of available information with which to identify target
systems and zones for investigation. SAAG Paper 001 [4] includes far more detail on the approach taken as a
case study which, in the interests of brevity, cannot be reproduced in this chapter. However, this work led to
a wider investigation on the methods of undertaking safety assessments for in-service aircraft.
SAE ARP4761 [24] provided valuable guidance for the processes and methods used to conduct safety
assessments. However, greater guidance for in-service platforms, where redesign is not always a feasible
option was required. This in turn led to the production of guidance on the conduct of Zonal Hazard Analysis
(ZHA) [17], referenced from the regulations [6] and has since been applied to most UK MOD aircraft fleets.
STO-TR-AVT-275 3-5
LESSONS FROM AGING AIRCRAFT SYSTEMS AUDITS
reasonably reliable estimate of the effect on the overall structural integrity of the aircraft can be identified
from a disassembled component. For a system, however, the link between damage found in a component and
the capability of the system to perform its required function or the effect on safety of the aircraft can be
less clear. Equally, the performance of the system may well have degraded without readily identifiable
physical evidence.
Therefore, where condition surveys or sampling programmes are being carried out, detailed functional
checks of the performance of target systems before undertaking survey or sampling were found to provide an
extremely useful yard stick. Such checks can identify degraded performance and assist in the understanding
of the significance of any observations found during survey or sampling. Such an approach can be invaluable
in prioritising recovery actions, following the identification of observations. Once a system is disturbed this
information is potentially lost.
For many systems (such as flying control systems), there will already be detailed functional checks in the
Air Publications (Aps). However, many of the checks will be pass/fail in nature. It is useful to identify
greater resolution than just pass/fail in order to determine whether the system performance is degrading and
whether any observations found are related to a degradation of performance. For systems without extant
functional checks, reference to any certification testing or release to service test schedules can provide a
valuable comparison of performance. In addition, maintenance staff with significant experience on the
aircraft type should be consulted to identify target systems where there are difficulties meeting the required
tolerances. Maintenance data may also show where systems are failing functional checks.
Consequently, the UK MOD / Industry embarked upon a comprehensive five-year EWIS Cradle-to-Grave
(C2G) programme. The programme was aimed at informing the regulation and best practice required for
3-6 STO-TR-AVT-275
LESSONS FROM AGING AIRCRAFT SYSTEMS AUDITS
future EWIS installations and providing the tools necessary to determine and influence EWIS in-service
condition. There are three sub-programmes: the Wiring Interconnect Systems Integrity Framework (WISIF),
the Wiring Insulation Material Deterioration Assessment Programme (WIMDAP) and the EWIS Condition
Surveys (EWISCS) [26]. The WISIF element of the programme has been completed with the development
of an EWIS handbook which captures guidance material necessary to support the EWIS integrity across the
MOD air fleets; this document will be referenced from the relevant regulatory articles. The aim of the
WIMDAP programme was to gain a greater understanding of the insulation properties of the major electrical
wire types used in service. The five-year testing programme of insulation properties of all the major wiring
types in use is now nearing completion. The EWISCS task aim was to undertake EWIS surveys of at least
two aircraft from all major platform types in service. This has now been completed and these surveys have
continued to highlight that EWIS issues start with less than perfect initial build and that the condition of
EWIS installations is further exacerbated by deterioration in service. The key issues identified were
maintenance practices and policy, initial and ongoing training and aircraft acceptance standards.
STO-TR-AVT-275 3-7
LESSONS FROM AGING AIRCRAFT SYSTEMS AUDITS
Understanding aging polymers is an issue far wider than just aerospace and these issues are being tackled in
a range of safety-related applications. However, the problem is often further complicated by insufficient
understanding of the effects of changes in the chemical constituents within the polymer over time and limited
understanding of the operating environment. That said, the complexity of the problem does not lessen the
implications and hence there was a need to gain a greater understanding of the degradation of polymers when
used in safety-related applications, with an aim of developing life assessment methods, where necessary.
Therefore, a programme was launched to develop a lifing framework for safety-related polymer components
(such as fuel seals). The concept was to identify what information was needed to support integrity decisions
for such materials and then to address any shortfalls in the data available. Much of the work has focused
upon the key materials used in polymer fuels seals and in gaining an understanding of the effects of static
and cyclic temperature variations on the compressive stress relaxation properties of the material (i.e., by how
much does its sealing force reduce). This work has now been completed and it is planned to publish it as
guidance material in the near future.
3.5 CONCLUSIONS
The inclusion of systems aspects within the AAA has provided a significant step forward in assuring the
airworthiness of the UK’s aging military aircraft fleets. Many lessons have been identified during these
programmes and the continual-improvement approach to enhance regulation and identify best practice has
been successful. However, there is still much work to be done and continually measuring the effectiveness of
airworthiness assurance methods and implementing improvement is clearly an essential component of any
airworthiness approach.
3.6 ACKNOWLEDGEMENTS
The work described in this chapter would not have been possible without the outstanding support provided
by the aging aircraft industrial and academic support base. Those organisations directly supporting the
programme were (in alphabetical order): AACE, Artis, Aviation Support Consultants, BAE SYSTEMS,
University of Bristol, CableConnect Solutions, Helisac, Musketeer Solutions, QinetiQ, RAE Structures, and
TWI. In addition, this programme has been entirely reliant upon cooperation across a wide range of
organisations within the UK MOD.
Finally, the invaluable contribution to this programme over many years from the late Mr Martin Hepworth,
who sadly passed away on 26 November 2017, is gratefully acknowledged.
3-8 STO-TR-AVT-275
LESSONS FROM AGING AIRCRAFT SYSTEMS AUDITS
3.7 REFERENCES
[1] US National Transportation Safety Board, “Aircraft Accident Report: Aloha Airlines Flight 243,
Boeing 737-200, N73711, 28 April 1988”, NTSB AAR-8903 Final Report 14 June 1988.
[2] US National Transportation Safety Board, “Aircraft Accident Report: In-Flight Breakup over the
Atlantic Ocean Trans World Airlines Flight 800”, NTSB AAR-00/03 Final Report, 17 July 1996.
[3] Transportation Safety Board of Canada, Aviation Report 1998 – A98H0003, (Swissair 111), 1998.
[4] Reed, S.C., “Lessons Identified from Initial Ageing Aircraft Systems Audits and Condition Survey
Programmes”, UK MOD Systems Airworthiness Advisory Group (SAAG) Paper 001, 5 May 2011.
[5] UK Ministry of Defence Military Aviation Authority, “Regulatory Article 5723 Ageing Aircraft
Audit”, Issue 3, 1 December 2014.
[6] UK Ministry of Defence Military Aviation Authority, “Regulatory Article 5721 Systems Integrity
Management”, Issue 6, 31 May 2018.
[7] Hepworth, M., “A Framework for Ageing Aircraft Audits”, UK MOD Ageing Aircraft Programmes
Working Group (AAPWG) Paper 010, June 2014.
[8] Moody, D, and Fairley, K., “MOD Aircraft Electrical Wiring Interconnect Systems Integrity”,
UK MOD Systems Airworthiness Advisory Group (SAAG) Paper 002, Issue C, 20 November 2012.
[9] Hepworth, M., “Continuing Airworthiness Management its Contribution to Identifying Evidence of
Ageing in Aircraft”, UK MOD Ageing Aircraft Programmes Working Group (AAPWG) Paper 013,
June 2017.
[10] Hoskins, M., “Condition Survey – Aircraft Interconnectivity”, UK MOD Systems Airworthiness
Advisory Group (SAAG) Paper 005, 3 November 2011.
[11] Air Safety Week Article, “NTSB Accident and Investigations 1”, 27 September 1999.
[12] UK Ministry of Defence Military Aviation Authority, “Regulatory Article 5720 Structural Integrity
Management”, Issue 6, 31 May 2018.
[13] Reed, S.C. and Holford, D.M., “Guidance for Aircraft Operational Loads Measurement Programmes”,
UK MOD Military Aircraft Structures Airworthiness Advisory Group (MASAAG) Paper 109,
31 May 2007.
[14] Reed, S.C., Terry, G.K. and Perrett, B.H.E., “Guidance for Helicopter Operational Data Recording
Programmes”, UK MOD Military Aircraft Structures Airworthiness Advisory Group (MASAAG)
Paper 120, 27 October 2017.
[15] Haddon-Cave, C., “An Independent Review into the Broader Issues Surrounding the Loss of RAF
Nimrod MR2 Aircraft XV230 in Afghanistan in 2006”, 28 October 2009.
[16] US Department of Transport, “A Review of the Flammability Hazard of Jet A Fuel Vapor in Civil
Transport Aircraft Fuel Tanks”, DOT/FAA/AR-98/26, 1998.
STO-TR-AVT-275 3-9
LESSONS FROM AGING AIRCRAFT SYSTEMS AUDITS
[17] Jones, J.P., and Wilson, M., “Guidance on the Conduct of Aircraft Zonal Hazard Analysis (ZHA)”,
UK MOD Ageing Aircraft Programmes Working Group (AAPWG) Paper 011, Issue 2, 28 September 2016.
[18] British Standard 6002-1 Sampling Procedures for Inspection by Variables (Part 1), 1993.
[19] US MIL-STD-414, “Sampling Procedures and Tables for Inspection by Variables for Percent
Defective”, 1957.
[20] UK Ministry of Defence, JAP(D) 100C-22, “Guide to Developing and Sustaining Preventive
Maintenance Programmes”, 10 December 2014.
[21] Name Redacted, “Review of the Lifing and Maintenance Policy for Aircraft System Seals”,
(Task ASG-ASI 1/07), DE&S(WYT)/366/8/2/CASD, 2007.
[22] Name Redacted and Name Redacted, “Lifing of Flexible Hoses Fitted to UK Military Aircraft”,
TES Task 14212.2, TES(WYT)/366/7/2/M&SG, 2007.
[23] Name Redacted and Name Redacted, “Lifing of Control Rods fitted to UK Military Aircraft”,
DE&S(WYT)/366/2/2/AVI, 2008.
[24] SAE Guidelines and Methods for Conducting the Safety Assessment Process on Civil Airborne
Systems and Equipment, Aerospace Recommended Practice (ARP 4761), 1996-12.
[25] UK Ministry of Defence, Aircraft Wiring Standards and Practices (Aircraft Wiring Husbandry)
General and Technical Information, AP101A-0005-1, 2nd Edition, March 2007.
[26] Fairley, K., UK MOD Defence Science and Technology Laboratory (Dstl), “EWIS Research
Programme”, presented at the Aircraft Airworthiness and Sustainment Conference, Brisbane, Australia,
6 July 2018.
[27] UK Ministry of Defence, “Regulatory Article 1140 Military Air Systems Technical Data Exploitation”,
Issue 3, 10 November 2014.
[28] Taylor, D., “Understanding the Corrosion Threat to Ageing Aircraft”, UK MOD Ageing Aircraft
Programmes Working Group (AAPWG) Paper 012, December 2015.
DISCLAIMER: Content includes material subject to © Crown copyright (2018), Dstl. This material
is licensed under the terms of the Open Government Licence except where otherwise stated. To view
this licence, visit http://www.nationalarchives.gov.uk/doc/open-government-licence/version/3 or
write to the Information Policy Team, The National Archives, Kew, London TW9 4DU, or email:
mailto:psi@nationalarchives.gsi.gov.uk..
3 - 10 STO-TR-AVT-275
Chapter 4 – EUROPEAN UNION AND NATO MILITARY
AIRCRAFT AIRWORTHINESS POLICY
4.1 INTRODUCTION
The principles and arrangements set up by the Convention on International Civil Aviation signed in Chicago
on 7 December 1944 (known as Chicago Convention) [1] and consequently the International Civil Aviation
Organization (ICAO) standards and recommended practices issued as annexes to The Chicago Convention,
are dedicated for civil aviation only. As a result, there was no agreement between European Union (EU)
Member States regarding military (state) aviation regulations, processes and procedures that would enable a
common approach in the field of Airworthiness (AW). Member States operate independent aviation safety
systems and are individually responsible for the regulation of their military (State) aircraft. As a result, all
military airworthiness activities had been conducted and regulated on a national basis. Some harmonisation
was achieved only at an individual project level and repeated for each new project. These multinational
programs demonstrated that this situation generated many shortcomings and was a primary cause for delay
and additional cost [1].
The MAWA Forum was tasked to harmonise military airworthiness requirements across Europe. To achieve
this goal, the participating Member States (MS) agreed to the implementation of the following Roadmap:
1) Common regulatory framework;
2) Common certification processes;
3) Common approach to organisational approvals;
4) Common certification/design codes;
5) Common approach to preservation of airworthiness;
6) Arrangements for recognition; and
7) Formation of a European Military Joint Airworthiness Authorities Organisation (EMJAAO).
In delivering the first 6 objectives, the EDA MAWA Forum has developed agreed-upon cooperation
principles, and harmonised airworthiness requirements and procedures. EDA’s participating MS are now
strongly encouraged to implement these principles, requirements and procedures (Table 4-1) into their
national regulations.
STO-TR-AVT-275 4-1
EUROPEAN UNION AND NATO
MILITARY AIRCRAFT AIRWORTHINESS POLICY
Table 4-1: List of Current MAWA Forum Airworthiness Documents and Requirements.
There is still no agreement between participating MS for the 7th goal – formation of EMJAAO. This
objective is not acceptable for some participating MS because of the concern that it could tread upon their
independence in the national defence domain.
Because the basic European Military Airworthiness System documents were already developed, a new
strategic objective was defined. On 5 December 2017 the EDA approved the Document for the Steering
Board (Nr SB Doc 2017/28) “Military Airworthiness Objectives for the Next Harmonisation Phase” [4].
“The Military Airworthiness Objectives in support of the EU Global Strategy” were published in the Annex
to this document. These objectives are grouped into three areas:
1) Development of a harmonised European Military Airworthiness System:
a) Permanent arrangements for the continuing development of common airworthiness policies and
maintenance of the European Military Airworthiness Requirements (EMARs) and related
documents;
4-2 STO-TR-AVT-275
EUROPEAN UNION AND NATO
MILITARY AIRCRAFT AIRWORTHINESS POLICY
During the process of developing coordinated a NATO Total Aviation Approach, Member Nations decided
to prepare the principal document regarding military airworthiness. The basic aim of this document was to
establish a robust framework that ensures airworthiness of aeronautical products, parts and appliances based
on the principles of economy of effort, cooperation and interoperability. On 10 July 2013, the North Atlantic
Council endorsed the document C-M(2013)0035 MC 0601/1 “NATO Airworthiness Policy” [5]. This policy
applies to NATO and its Member and Partner Nations.
STO-TR-AVT-275 4-3
EUROPEAN UNION AND NATO
MILITARY AIRCRAFT AIRWORTHINESS POLICY
a responsibility to provide airworthy products to their aircrew, ground crew, passengers, and to
third parties”.
The above statement was summarized particularly in paragraphs 9 and 10 of the NAWP:
“9. All aeronautical products, parts and appliances provided on behalf of NATO shall be:
a) Certified as airworthy by a NATO recognized airworthiness authority;
b) Properly controlled in accordance with approved continued airworthiness provisions; and
c) Operated and maintained in accordance with approved continuing airworthiness
provisions.
10. All work associated with the airworthiness process shall be performed by authorised individuals
employing approved processes within organizations accredited/approved by a NATO recognized
airworthiness authority.”
There is one exception in the policy: for any aeronautical products, parts and appliances provided on behalf
of NATO that do not have a NATO recognized authority. Paragraph 15 states that only for exceptional
circumstances may Supreme Headquarters Allied Power Europe (SACEUR) waive the requirements by a
reasoned decision and for limited duration, reserving the right to accept risks in order to meet urgent
operational needs.
There are three established stakeholders who fulfil of NAWP principles: NATO Aviation Committee (AVC),
NATO Airworthiness Executive (NAE) and Airworthiness Advisory Group (AWAG).
In order to execute the airworthiness policy additional key documents were prepared [6], [7], [8]:
• NATO Airworthiness Policy (NAWP) Implementation Plan (NAWP IP);
• AWAG Action Plan (AP); and
• NATO Recognition Process (NRP).
4-4 STO-TR-AVT-275
EUROPEAN UNION AND NATO
MILITARY AIRCRAFT AIRWORTHINESS POLICY
The NATO Recognition Process assessment consists principally of nine steps, shown in Table 4-2.
Stakeholders for these steps of NRP are NAE, AWAG and AVC [8].
The status of the NATO recognition process is presented in Table 4-3 [9].
Table 4-3: NATO Recognition Process Status Updated for September 2018.
procedural issue
Hungarian MAA √ In progress √
French MAA(s) In
√ √ √ √ √ √ √
DSAE+DGA progress
Canadian MAA In
√ √ √ √ √
progress
Norwegian MAA In
√ √ √ √ √
progress
Spanish MAA
√ In progress √
(DGAM)
STO-TR-AVT-275 4-5
EUROPEAN UNION AND NATO
MILITARY AIRCRAFT AIRWORTHINESS POLICY
The NATO military authorities determined that an Airborne Early Warning (AEW) capability would provide
the key to meeting the challenge. The operational requirement for the NATO AEW system stressed the need
to detect small cross-section, high speed intruder aircraft at long range. The need to detect maritime surface
targets was also specified to account for the geographical regions in which the AEW aircraft would be
required to operate. The inherent mobility and flexibility of the system, especially for control function, were
also foreseen by NATO planners as providing air, maritime, and land force commanders with an enhanced
Command and Control (C2) capability [10].
In December 1978, NATO approved acquisition of 18 aircraft based on the US Air Force (USAF) Airborne
Warning and Control System (AWACS) to be operated as an Alliance-owned Airborne Early Warning
system (Figure 4-1).
4.4.2 NATO Airborne Early Warning and Control (NAEW&C) Airworthiness System
The NE-3A was introduced to service in NATO in 1982. The aircraft were registered by Luxembourg with
the understanding that the NATO aircraft configuration and Airworthiness certifications would mirror the
USAF E-3, that NATO aircraft would be maintained in accordance with regulations adapted from USAF,
and that the USAF Airworthiness process would provide the requisite certifications for the NATO
programme. This airworthiness concept, where the NE-3A aircraft mirrored the certified USAF E-3
configuration, and the USAF certification was accepted by NATO as sufficient, with NATO providing
individual aircraft airworthiness certifications (Military Airworthiness Certificates – MCAs), was used from
program inception until the early 2000s [11].
4-6 STO-TR-AVT-275
EUROPEAN UNION AND NATO
MILITARY AIRCRAFT AIRWORTHINESS POLICY
In 2008, NATO pursued a unique mission systems modernization program, NATO Mid-Term (NMT), which
had significant changes to the NE-3A baseline configuration [11]. In years 2007 – 2009 the NMT
configuration was further upgraded to provide self defence capability: Large Aircraft Infrared
Countermeasures (LAIRCM). This modification was assessed as “reportable” (per MIL-HDBK-516) [12].
With these changes in configuration, the USAF communicated that it was no longer able to support the
NE-3A Airworthiness Authority, and that to maintain the appropriate level of airworthiness certification
(to include NMT-appropriate Military Type Certificates and Military Certificates of Airworthiness) NATO
needed to establish its own Airworthiness capability [11].
The Force Command Commander recognized the challenge of the unique NE-3A configuration and his
responsibilities as the Airworthiness Authority. Consequently in 2008(9) Force Command and US Agent
called for an Integrated Product Team (IPT) to review and establish Airworthiness assurance process based
on USAF E-3 best practices [12].
Meanwhile in the years 2008 – 2011, Software (SW) and Hardware (HW) modifications continued with
approval by separate, independent Configuration Control Boards. The modifications included: HW
modifications to improve mission computing performance (configuration control lead by NATO AEW&C
Programme Management Agency – NAPMA), capabilities developed for the USAF fleet and implemented
by Force Command (configuration control lead by Force Command) and mission system SW updates
developed at the Main Operating Base (configuration control lead by NE-3A Component) [12]. There was
no formal mechanism to cross check impact of changes on the fleet.
In December 2011, the NAPMO (NATO AEW&C Programme Management Organisation) Board of
Directors (BOD) approved Airworthiness Policy based on IPT study recommendation i.e., split airworthiness
responsibility between Force Command and NAPMA General Manager (NAPMA GM) established Chief
Engineer’s Office to support and execute Operational Safety Suitability and Effectiveness (OSS&E) and AW
checklist process [12].
In 2012 the BOD took the first step of approving an Airworthiness concept via a Memorandum of
Understanding for Airworthiness and Configuration Management (AW MOU), that made the NAPMA GM
the Technical Airworthiness Authority (TAA) and the Force Commander the Operational Airworthiness
Authority (OAA) [13]. This document requires Airworthiness Review Board (ARB), defines Configuration
Management Responsibilities including Single Configuration Control Board (SCCB) and defines Technical
and Operational Airworthiness responsibilities.
An Airworthiness Review Board (ARB) was established to assist the NAPMO Nations on airworthiness
matters related to the NE-3A fleet [14]. The Terms of Reference for the ARB [15], including membership, was
developed jointly by the OAA and TAA for NAPMO BOD approval. The ARB responsibilities include [13]:
• Conducting an annual review of the airworthiness programmes and procedures established by the
OAA and TAA and reporting to the NAPMO BOD on its observations and concerns; and
• Providing independent advice to the NAPMO BOD regarding the information provided
semi-annually by the OAA and TAA.
STO-TR-AVT-275 4-7
EUROPEAN UNION AND NATO
MILITARY AIRCRAFT AIRWORTHINESS POLICY
• Maintaining and operating the NE-3A fleet and related assets in accordance with the certified
baseline (continuing airworthiness);
• Supporting the TAA in the implementation of technical airworthiness procedures and criteria; and
• Keeping the NAPMA GM or his delegated representative informed on all common airworthiness
issues by exchanging information concerning the NE-3A fleet in their respective area of
responsibility; and informing the NAPMO BOD on a semi-annual basis of the operational
airworthiness of the NE-3A aircraft fleet [13].
The Single Configuration Control Board (SCCB) was established to manage the NE-3A certified
configuration baseline. The SCCB is co-chaired by the OAA and the TAA and consist of representatives
from Force Command, the NE-3A Component, NAPMA and other NE-3A stakeholder organisations
approved for membership by the Participants. All proposed changes to the certified configuration baseline of
the NE-3A system, including temporary modifications, must be reviewed by the SCCB.
The AW Memorandum is associated with an Implementing Agreement (IA) [16] that defines the agreed
hardware, software and documentation elements describing the currently fielded NE-3A system
configuration, together with the available airworthiness assurance evidence, applicable on initial
establishment of the TAA; and outlines the key responsibilities of the OAA and the TAA.
The TAA maintains the configuration baseline, as approved through the SCCB. The TAA has authority to
approve changes to the certified configuration baseline, including temporary modifications of the NE-3A
system and for related configuration controlled fielded assets.
4-8 STO-TR-AVT-275
EUROPEAN UNION AND NATO
MILITARY AIRCRAFT AIRWORTHINESS POLICY
The TAA is responsible for regulating the design, manufacture, maintenance and material support for the
NE-3A fleet. The Technical Airworthiness of the NE-3A system is assured and documented through the use
of the OSS&E AW process. The NAPMA Chief Engineer oversees the OSS&E AW process for all proposed
changes to the aircraft [16]. The TAA ensures that the conditions under which the aircraft design was
certified, including its specifications, manufacturing and configuration data, remain valid. In the event of any
deviation from the applicable certification basis he defines any necessary corrective measures. The TAA also
sets requirements to ensure the airworthiness of the NE-3A during its service life [16].
The TAA supports the OAA by providing engineering assistance for aircraft incidents and accidents and
reviews repair instructions and/or flight limitations together with the Original Equipment Manufacturer
(OEM). The support also includes the provision of access to NAPMA contractor support, OEM and support
agencies where applicable [16].
The OAA has authority to approve changes for operations and support as defined in the implementing
documents that do not affect the certified configuration baseline for the fielded NE-3A system and for related
configuration controlled fielded assets.
When there is a compelling operational need to employ an NE-3A aircraft for which it is infeasible or
impractical to comply with the certified configuration baseline, the OAA may issue a Special Flight Release
(SFR) after consultation with the TAA to ensure technical risks are understood.
The OAA supports the OSS&E Process by assigning a Change Project Officer, as required, to coordinate the
OSS&E AW Process for experimental changes as well as changes for operations and support; and by
providing focal point expertise when requested by the TAA. The OAA is responsible for regulating the
operations, operational personnel and operational facilities of the NE-3A fleet.
The OAA supports the TAA by providing expertise and data to support the establishment of the baseline
configuration of the NE-3A. He/She also reviews current waivers with the TAA and provides access to
Operation and Sustainment contracted support and support agencies where applicable. The OAA also assists
in the development of new working relationships with support agencies.
The OAA fulfils the responsibilities of the Operating Nation for Accident Investigation, and initiates
investigations into accidents and occurrences involving NE-3A aircraft.
The joint responsibilities of the TAA and OAA include coordination between the Authorities, Safety
Management (on the principles outlined in ICAO Doc 9859 and MIL-STD-882), Configuration Management
(co-chairing the SCCB), Change Management [16].
4.5 SUMMARY
The basic Airworthiness (AW) principles are common for each and every stakeholder of military aviation.
However, the detailed approaches, and the ways of implementing them in the particular systems differ between
Nations and users. The European Defence Agency has initiated an activity aiming for development of a
common (commonly understood by all EU Nations) system of AW requirements, and further on regulations,
for European military aviation. The EDA Military Airworthiness Authority Forum (EDA MAWA Forum) was
established to achieve this goal, and a set of European Military Airworthiness Requirements was developed
under the Forum umbrella. The importance of having common approach to the AW issues in military domain is
very well understood by all national military AW authorities (or their equivalents) and a process of
harmonisation, bilateral and multilateral recognition is continuously in progress.
NATO missions and operations are supported by aircraft which are owned, leased, rented or chartered by
NATO Member, Partners or other contributing Nations. Airworthiness certification is a sovereign responsibility
STO-TR-AVT-275 4-9
EUROPEAN UNION AND NATO
MILITARY AIRCRAFT AIRWORTHINESS POLICY
of aircraft originating country; however, it is assumed that all aircraft contributing to NATO or NATO led
missions and operations are airworthy, as certified by an appropriate national Authority. NATO affiliated
assets, such as NATO E-3A (NE-3A), C-17 fleets and Alliance Ground Surveillance (AGS) system, have also
to comply with this requirement. There is no single “appropriate Authority” for these kinds of assets. The C-17s
(Strategic Airlift Capability – SAC) are registered in Hungarian (HUN) register of military aircraft and are
under authority of HUN Military Airworthiness Authority. The AGS systems are registered in the Italian (ITA)
Military Aircraft Registry and are under authority of Directorate of Air Armaments and Airworthiness (ITA
MAA). The NE-3A aircraft are registered in Luxemburg register, however, there is no single Military
Airworthiness Authority for this fleet. The NE-3A airworthiness issues are addressed in intentionally developed
(and signed) memorandum of understanding. To address such “complicated” (from AW perspective)
environment, NATO has developed the NATO Airworthiness Policy (NAWP) and NAWP Implementation
Plan, where a recognition process between stockholders: NATO – National MAA(s) is a crucial element.
Looking at the subject under discussion from a Polish perspective, it is necessary to mention that according to
Polish military regulations, the Chief Engineer of Military Aviation (GIWL) is a Polish NMAA. The GIWL,
based on authorisation granted from the Ministry of National Defence, establishes rules and regulations of
military airworthiness, which are common for aviation of all Polish military services: Air Force, Army and
Navy. The Polish military airworthiness system is built on decades of operational experience of military
aviation in peacetime and wartime conditions. The AW system is comprehensive and dedicated to safety in the
military environment. Presently there is no full compliance of the Polish AW system with requirements
provided in EMARs. Some differences are caused by Polish military aircraft operation and maintenance
philosophy derived from the past experiences and harmonisation with a different philosophy of aircraft design
(post-Soviet), which formed the majority of the fleet. However, along with the aircraft replacement process,
where several “western designed” aircraft (F-16, C-130, C-295, M-346, G550, B737) have already been
implemented into the fleet, the AW regulations are also changing. Poland has signed the agreement in principle
of The European Harmonised Military Airworthiness Basic Framework Document (BFD) [17]; therefore, in
order to fulfil the obligation for harmonisation of European military airworthiness regulations, bilateral and
multilateral recognition, the GIWL and special Attorney of Ministry of National Defence for Military Aviation
Authority are responsible for updates of Polish AW regulations to be compliant with EMARs.
4.6 REFERENCES
[1] International Civil Aviation Organization, Convention on International Civil Aviation. Doc. 7300/9.
[2] European Defence Agency, Document “Roadmap on an EU-wide Forum for Military Airworthiness
Authorities (MAWA)” EDA Document No. 2008/39.
[3] European Defence Agency, Defence Ministers’ Political Declaration Regarding the Timely
Development and Implementation of the European Military Airworthiness Requirements”, Document
No. 2009/36.
[4] European Defence Agency, “Military Airworthiness Objectives for the Next Harmonisation Phase”.
Document No. SB Doc 2017/28.
[5] North Atlantic Council, “NATO Airworthiness Policy”, Document No. C-M(2013)0035 MC 0601/1.
[6] North Atlantic Council, “NATO Airworthiness Policy (NAWP) Implementation Plan (NAWP IP)”,
Document No. C-M(2016)0034 (INV).
4 - 10 STO-TR-AVT-275
EUROPEAN UNION AND NATO
MILITARY AIRCRAFT AIRWORTHINESS POLICY
[8] Aviation Committee (AVC), “Refined NATO Recognition Process (NRP)”, Document No.
AC/92-D(2018)0009.
[9] NATO Aviation Committee, “Airworthiness Advisory Group Decision Sheet”, Document
No. AC/92(AW)DS(2018)0001.
[10] NATO AEW&C Programme Management Agency, “NAEW & C Programme – The Onset”,
http://www.napma.nato.int/awacs/9.html (accessed 2018.09.06).
[11] Headquarters NATO AEW&C Force, “Force Order 3.13-5, “NATO E-3A Fleet Operational
Airworthiness”, dated 14 June 2018.
[12] NATO AEW&C Programme Management Agency, “NATO E-3A Technical Airworthiness”, prepared
for: Special Airworthiness Review Board Meeting 8 May 2018.
[13] NATO AEW&C Programme Management Agency, Headquarters NATO AEW&C, “Memorandum of
Understanding between the NATO Airborne Early Warning and Control (NAEW&C) Programme
Management Organisation (NAPMO) and Supreme Headquarters Allied Powers Europe (SHAPE)
Concerning Airworthiness and Configuration Management Responsibilities for the NATO AEW&C
E-3A Aircraft”, dated 4 June 2012.
[15] NATO AEW&C Programme Management Agency, Headquarters NATO AEW&C, “Terms of
Reference for the NE-3A Airworthiness Review Board,” dated 12 December 2012.
[16] NATO AEW&C Programme Management Agency, Headquarters NATO AEW&C, “Implementing
Agreement for the Memorandum of Understanding between the NAPMO and SHAPE concerning
Airworthiness and Configuration Management Responsibilities for the NATO AEW&C E-3A
Aircraft,” dated 4 June 2012.
[17] MAWA Forum, “The European Harmonised Military Airworthiness Basic Framework Document
(BFD)”, Ed. 2.0, 23 May 2013.
STO-TR-AVT-275 4 - 11
EUROPEAN UNION AND NATO
MILITARY AIRCRAFT AIRWORTHINESS POLICY
4 - 12 STO-TR-AVT-275
Chapter 5 – EUROPEAN MILITARY AIRWORTHINESS
MAINTENANCE REQUIREMENTS
5.1 INTRODUCTION
A complete set of harmonized European Military Airworthiness Requirements (EMARs) has been
developed under the European Defence Agency Military Airworthiness Authority Forum (EDA MAWA
Forum) umbrella. In parallel, the Australian Department of Defence, introduced the Aviation Safety
Regulations (DASRs) in 2016. The DASRs align with the EMARs and both sets of airworthiness
requirements contain specific provisions for maintenance of military products (aircraft, engines and
propellers) and components including software.
Responsibility for the oversight and realization of any Aircraft Structural Integrity Program (ASIP) [1],
Avionics Integrity Program (AVIP) [2], Mechanical Equipment and Subsystems Integrity Program
(MECSIP) [3], Engine Structural Integrity Program (ENSIP) [4], Propulsion Structural Integrity Program
(PSIP) [5], Weapon System Integrity Program (WSIP) [6] or their equivalents is spread among the
Military Airworthiness Authority (MAA), Military Design Organisation or Military Type Certificate
Holder (MTCH), Aircraft Maintenance Organisation (AMO) and Continuing Airworthiness Management
Organisation (CAMO).
According to EMAR M [7] and DASR M [8] provisions, maintenance of each military aircraft shall be
organised in accordance with the Aircraft Maintenance Programme (AMP). Thus any ASIP, AVIP,
MECSIP, ENSIP, PSIP and WSIP or analogues must be a part of well-developed AMP.
It is worth mentioning that in contrast to their civilian equivalents (i.e., Part-M), EMAR M allows
OO to contract/task so-called external CAMO for the management of the continuing airworthiness of the
military aircraft.
These provisions may lead to several more or less complex maintenance arrangements but due to
limitation of this chapter length, only three of them are described below, in brief, and are presented
schematically in Table 5-1.
STO-TR-AVT-275 5-1
EUROPEAN MILITARY
AIRWORTHINESS MAINTENANCE REQUIREMENTS
Symbol Process
Contracting/Tasking
Coordination of maintenance activities
5.2.1 Case 1
The OO is responsible for the continuing airworthiness of the military aircraft it operates and:
• Is approved in accordance with EMAR M Subpart G; and
• Is approved in accordance with EMAR 145.
Contract/Tasking Arrangement(s) between OO, internal CAMO and internal AMO are not required. Internal
CAMO is responsible for proper coordination of all maintenance activities with internal EMAR 145 AMO.
Internal AMO is responsible for accomplishing all maintenance activities requested by Internal CAMO.
5.2.2 Case 2
The OO is responsible for the continuing airworthiness of the aircraft it operates and:
• Is approved in accordance with EMAR M Subpart G; and
• Contracts/tasks directly external EMAR 145 AMO.
5-2 STO-TR-AVT-275
EUROPEAN MILITARY
AIRWORTHINESS MAINTENANCE REQUIREMENTS
5.2.3 Case 3
The OO is responsible for the continuing airworthiness of the aircraft it operates and:
• Contracts/tasks external CAMO directly; and
• Contracts/tasks external EMAR 145 AMO directly.
Contract/Tasking Arrangement between OO and external CAMO and between OO and external EMAR 145
AMO are required. External CAMO is responsible for proper coordination of all maintenance activities with
external EMAR 145 AMO. External AMO is responsible for accomplishing all maintenance activities
contracted/tasked by OO.
Reference [9] provides information about splitting up responsibility for the execution of ASIP elements
among the approved design, maintenance and continuing airworthiness management organisations from the
Defence Aviation Safety Regulations (DASR) perspective. Keeping the above in mind and in light of
EMAR’s provisions, responsibilities for the execution of any Integrity Process (IP) 1 elements should be
distributed among the approved organisation as follows:
• Military Type Certificate Holder (MTCH) or EMAR 21/J Military Design Organisation (MDO)
responsibility;
• Establish certification basis;
• Establish integrity program;
• Ensure access to required type design data;
• Conduct structural assessments and tests (including occurrence reporting);
• Establish Loads/Environment Spectra Survey (L/ESS) and Individual Aircraft Tracking (IAT)
systems or Operational Loads Monitoring (OLM) program, where applicable;
• Maintain design/initial/continued airworthiness aspects of the integrity programme; and
• Provide occurrence reporting.
1
ASIP, AVIP, MECSIP, ENSIP, PSIP, WSIP.
STO-TR-AVT-275 5-3
EUROPEAN MILITARY
AIRWORTHINESS MAINTENANCE REQUIREMENTS
In addition, the MAA also has some responsibilities in relation to any Integrity Program, which result from
Aircraft Continuing Airworthiness Monitoring (ACAM) provisions of the EMAR M (EMAR M.B.303
refers) [7]. ACAM puts obligations on the MAA to create and execute an airworthiness survey programme
for monitoring the airworthiness status of the aircraft fleet on its military register, based on 14 Key Risk
Elements (KREs). The KREs that are directly related to an Integrity Programme are:
• A.1 – Type design and changes to type design e.g., information on materials and processes and on
methods of manufacture and assembly of the product;
• A.2 – Airworthiness limitations e.g., Critical Design Configuration Control Limitations (CDCCL)
for the fuel tank safety;
• A.3 – Airworthiness Directives;
• B.6 – Defect management;
• B.7 – Symmetry check;
• C.1 – Aircraft Maintenance Programme;
• C.2 – Component control;
• C.3 – Repairs; and
• C.4 – Records.
The EMAR 145 AMO may issue an authorisation to its employees to provide certification of maintenance for
specialised services e.g., NDT, composite repairs or Arms, Munitions and Pyrotechnic Systems services on the
basis of appropriate competence, training and experience in accordance with a procedure(s) contained in the
Maintenance Organisation Exposition. Table 5-2 presents a release to service of NDT tasks scheme.
EMAR 145 Certifying Staff (CS) NDT Personnel Release Procedure for
AMO Class with EMAR 66 Qualification and an NDT Task
Military Aircraft Certification System
Maintenance Licence
(MAML) Required
5-4 STO-TR-AVT-275
EUROPEAN MILITARY
AIRWORTHINESS MAINTENANCE REQUIREMENTS
EMAR 145 Certifying Staff (CS) NDT Personnel Release Procedure for
AMO Class with EMAR 66 Qualification and an NDT Task
Military Aircraft Certification System
Maintenance Licence
(MAML) Required
Due to the nature of ASIP, AVIP, MECSIP, ENSIP, PSIP or WSIP related maintenance activities, they
should be classified as specialised services and in order to provide such maintenance services, in accordance
with the current EMAR 145 provisions, an organisation should be either:
• EMAR 145 Class A approval rating – an EMAR 145 (AMO holding an A approval rating on a
particular aircraft type must have in its approved scope of work ASIP, AVIP, MECSIP and/or WSIP
maintenance tasks for this aircraft type. This organisation needs to have EMAR 66 certifying staff
with MAMLs and ASIP, AVIP, MECSIP and/or WSIP personnel qualified in accordance with
standards approved by MAA. In this case the ASIP, AVIP, MECSIP and/or WSIP inspector
performs his/her task and signs the work order. The aircraft is released by appropriately qualified
B1, B2 or C certifying staff under the AMO’s A rating; or
• EMAR 145 Class B approval rating – an EMAR B AMO holding a B approval rating on a particular
engine type must have in its approved scope of work ENSIP and/or PSIP maintenance tasks for this
engine type. This organisation needs to have ENSIP and/or PSIP personnel qualified in accordance
with standards approved by MAA. In this case the ENSIP and/or PSIP certifying staff performs and
releases the ENSIP and/or PSIP task on an EMAR Form 1 or on a release to service document
defined by the AMO in the Maintenance Organisation Exposition; or
• Non-approved organisation working under quality system of contracting/tasking of EMAR 145
AMO Class A or B, when appropriate. The contracted/tasked non-approved organisation needs to
have ASIP, AVIP, MECSIP, ENSIP, PSIP or WSIP staff qualified in accordance with standards
acceptable to MAA In this case the staff of contracted non-approved organisation performs and
releases the ASIP, AVIP, MECSIP, ENSIP, PSIP or WSIP task on a release to service document
defined by the contracting/tasking EMAR 145 AMO.
STO-TR-AVT-275 5-5
EUROPEAN MILITARY
AIRWORTHINESS MAINTENANCE REQUIREMENTS
the Maintenance Review Board (MRB) report or equivalent report where applicable, the Maintenance
Planning Document (MPD), the relevant chapters of the maintenance manual or any other maintenance data
containing information on scheduling. Furthermore, an OO’s AMP should also take into account any
maintenance data containing information on scheduling for components. Typical sources for AMP
development are shown on Figure 5-1.
EMAR M makes it possible to use others than MRB processes for AMP development. Practically it may lead
to use two alternative approaches: Reliability-Centered Maintenance (RCM) [11] or Continuous
Improvement Preventive Maintenance, S4000P [12]. Detailed discussions about these approaches are outside
the scope of this chapter and therefore only a cursory comparison is shown in Table 5-3.
5-6 STO-TR-AVT-275
EUROPEAN MILITARY
AIRWORTHINESS MAINTENANCE REQUIREMENTS
Other structure
Standard zonal
SSI analysis
Economy
Mission
analysis
analysis
Safety
?
S4000P
structure analysis
Maintenance rel.
L/HIRF analysis
Enhanced zonal
Other structure
Standard zonal
analysis incl.
SSI analysis
Operation
Economy
Ecology
Mission
analysis
analysis
Safety
Law/
MSG-3
L/HIRF analysis
Enhanced zonal
Other structure
Standard zonal
SSI analysis
Operation
Economy
analysis
analysis
analysis
Safety
EMAR M (M.A.302(f) refers) [7] requires that any AMP should contain a Reliability Programme (RP)
unless otherwise specified by MAA. It is a substantial difference in comparison with Part-M provisions
(Part-M: M.A.302(f) refers), which requires a RP to be included in the AMP only for complex
motor-powered aircraft (CMPA) [14], when the AMP is based on Maintenance Steering Group (MSG) logic
or on Condition Monitoring.
RP’s ultimate aim is to ensure that the all AMP tasks are effective and their frequency is appropriate. The RP
has two basic functions:
• RP is understood as an element of statistical processing of reliability data, providing summary
information on the reliability of the entire aircraft fleet and its components, thus reflecting the
effectiveness of the AMP being implemented, providing appropriate ways to monitor the
effectiveness of the AMP; and
• RP provides timely and important technical information that allows reliability to be maintained at
the required and acceptable level through changes to the AMP or ways of its implementation. The
result may be the extension or limitation of maintenance activities, as well as the re-extension or
addition of maintenance activities.
STO-TR-AVT-275 5-7
EUROPEAN MILITARY
AIRWORTHINESS MAINTENANCE REQUIREMENTS
The RP enables continuous control and analysis of maintenance activities, thereby revealing unfavourable
trends in the maintenance process and effecting the effective corrective actions. These actions should take
into account the economic aspects, in particular, the cost of achieving required reliability level.
The reliability of the aircraft, aircraft systems, components, engines and Auxiliary Power Units (APU) are
monitored through comparison with established reliability quantitative parameter alarm levels. This system
prevents the development of unfavourable tendencies as a result of corrective actions in due time.
The resulting reliability quantitative parameters and the corresponding alarm levels cannot be compared with
reliability parameters and alarm values provided by manufacturers or other aviation organisations without
prior checking of the compatibility of the applied analytical methods.
In an RP alarm, values definitions can be different for a product (aircraft, engines, propeller) vs. components,
as illustrated in the Table 5-4.
5-8 STO-TR-AVT-275
EUROPEAN MILITARY
AIRWORTHINESS MAINTENANCE REQUIREMENTS
The Alarm Levels (AL) could be defined in various ways from the simple to very complex e.g., such as:
AL = Mean of RQP ±2 * Standard Deviation of RQP (5-1)
The objective of the occurrence reporting is the prevention of accidents and incidents by means of reporting,
collecting, storing, protection and dissemination of the occurrence related information only, and not to assign
blame or responsibility on the reporting persons.
The details about occurrence reporting will be included in EMAD 20-20, which is currently under
development. It is expected that in case of aircraft technical occurrences related with the structure, the
following occurrences should be reported:
• Damage to a Principal Structural Element (PSE);
• Damage to or defect of a structural element exceeding allowed tolerances whose failure could:
reduce the structural stiffness, result in the liberation of items of mass or jeopardise proper operation
of systems; and
• Loss of any part of the aircraft structure in flight.
STO-TR-AVT-275 5-9
EUROPEAN MILITARY
AIRWORTHINESS MAINTENANCE REQUIREMENTS
Table 5-5: EMAR’s Occurrence Reporting Requirements. Source: Author’s own study.
The UK Civil Aviation Authority (UK CAA) oversees approximately 1,400 production, maintenance,
continuing airworthiness and maintenance training organisations. Each year, it performs over 4,000 audits
of these organisations and detects in their course several thousand findings, of which about 14% concern
the organisation, including about 1/3 of non-compliances related to the ineffective quality/compliance
system [15]. Of the three major sources of non-compliance with a high and moderate impact on safety
identified by the UK CAA, unfortunately, the quality system is put on the outsider’s position [15] – see
Figure 5-2.
The Polish Civil Aviation Authority (PL CAA) oversees about 60 Part-145 maintenance organisations
and about 90 continuing airworthiness management organisations. Annually, it conducts several hundred
audits in these organisations, during which it detects on average 6 – 10 non-compliances per audit [16]. The
main sources of non-compliance identified by PL CAA during certification audits and audits as part of
ongoing oversight are presented in Figure 5-3, from which it can be seen that the main sources of
non-compliance should be sought in the organisation exposition, personnel, quality system and aircraft
maintenance programme.
5 - 10 STO-TR-AVT-275
EUROPEAN MILITARY
AIRWORTHINESS MAINTENANCE REQUIREMENTS
Figure 5-2: The Three Principal Sources of Non-Compliance with High and Moderate Safety
Impacts Identified by the UK CAA During Audits. Source: Author’s own study based on Ref. [15].
Figure 5-3: The Principal Sources of Non-Compliance Identified by PL CAA During Audits in
Supervised Part-145 and CAMO Organisations. Source: Author’s own study based on Ref. [16].
5.6 CONCLUSIONS
As indicated in the summary of Chapter 4.5, the current airworthiness system of military aircraft in Poland is
not fully compliant with EMAR, although it provides equivalent level of safety.
The gradual replacement of military aircraft fleet in Poland, including in particular the acquisition of new
Military Commercial Derivative Aircraft (MCDA) i.e., G550 and B737-800 has necessitated a gradual
STO-TR-AVT-275 5 - 11
EUROPEAN MILITARY
AIRWORTHINESS MAINTENANCE REQUIREMENTS
change in Polish military airworthiness regulations towards fulfilling NATO commitments and taking full
advantage from a harmonized approach to airworthiness in a complementary sense to EASA and
non-European military airworthiness authorities.
The challenge does not necessarily lie in the general understanding of EMAR requirements, but in the
detailed implementation of distinct requirements in Polish military airworthiness regulations, in particular
EMAR 145 and EMAR M planned for the near future.
5.7 REFERENCES
[1] MIL-STD-1530D, Department of Defense Standard Practice. Aircraft Structural Integrity Program,
13 October 2016.
[2] MIL-STD-1796A, Department of Defense Standard Practice. Avionics Integrity Program (AVIP),
13 October 2011.
[3] MIL-STD-1798C, Department of Defense Standard Practice. Mechanical Equipment and Subsystems
Integrity Program, 8 August 2013.
[4] MIL-HDBK-1783B, Department of Defense Handbook. Engine Structural Integrity Program (ENSIP),
22 September 2004.
[5] MIL-STD-3024, Department of Defense Standard Practice. Propulsion System Integrity Program
(PSIP), 13 July 2015.
[6] MIL-HDBK-515, Department of Defense Handbook. Weapon System Integrity Guide (WSIG),
11 October 2002.
[9] Defence Aviation Safety Authority, DASR Aircraft Structural Integrity Guide,
https://www.google.pl/search?sourceid=navclient&hl=pl&ie=UTF-8&rlz=1T4PLXB_plPL658PL659
&q=DASR+aircraft+structural+integrity+guide. Accessed [04.06.2018].
[10] European Union Aviation Safety Agency, FAQ n.19055, Release to Service of NDT tasks by Part-145
Organisations. https://www.easa.europa.eu/faq/19055. Accessed [23.05.2018].
[11] NAVAIR 00-25-403, Guidelines for the Naval Aviation Reliability-Centered Maintenance Process,
US Navy’s Naval Air Systems Command (NAVAIR), 01 August 2011.
[12] S4000P. International Specification for Developing and Continuously Improving Preventive
Maintenance, Aerospace and Defence Industries Association of Europe, Issue No. 2.0, 01 August 2018.
https://en.wikipedia.org/w/index.php?title=Aerospace_and_Defence_Industries_Association_of_Europ
e&action=edit&redlink=1.
[13] Haslam, P. International Specification for Developing Scheduled Maintenance Programs, S1000D User
Forum 2013, Vienna 2013-09-19. https://www.aaig.at/wp-content/uploads/S7-Paul-Haslam-S-Series-
Spec-Day-2013-S4000M.pdf. Accessed [10.09.2018].
5 - 12 STO-TR-AVT-275
EUROPEAN MILITARY
AIRWORTHINESS MAINTENANCE REQUIREMENTS
[14] Commission Regulation (EU) No 1321/2014 of 26 November 2014 on the Continuing Airworthiness
of Aircraft and Aeronautical Products, Parts and Appliances, and on the Approval of Organisations and
Personnel Involved in These Tasks.
[15] UK Civil Aviation Authority, Airworthiness Analysis, Corporate Aviation Seminar 29 June 2016.
https://www.caa.co.uk/uploadedFiles/CAA/Content/Standard_Content/Commercial_industry/Aircraft/
Airworthiness/Seminars/Corporate_aviation_June_2016/FWM20160629_08_Airworthiness%20Analy
sis.pdf. Accessed [23.05.2018].
[16] Mieciek, K. Statystyki oraz wnioski z nich płynące w kontekście zagrożeń występujących w
organizacjach obsługowych (AMO) oraz zarządzających ciągłą zdatnością do lotu (CAMO) będących
pod nadzorem ULC (in Polish), http://www.ulc.gov.pl/_download/bezpieczenstow_lotow/konferencje/
2015/konferencja_bezpieczenstwa_10_2015/Wska%C5%BAniki_organizacji_AMO_i_CAMO_krzysz
tof_Mieciek.pdf. Accessed [23.05.2018].
STO-TR-AVT-275 5 - 13
EUROPEAN MILITARY
AIRWORTHINESS MAINTENANCE REQUIREMENTS
5 - 14 STO-TR-AVT-275
Part 2: AIRCRAFT STRUCTURAL SYSTEMS
STO-TR-AVT-275 Part 2 - i
Part 2 - ii STO-TR-AVT-275
Chapter 6 – AIRWORTHINESS OF STRUCTURES
(UNITED STATES)
Charles A. Babish IV
Air Force Life Cycle Management Center
UNITED STATES
6.1 INTRODUCTION
The United States Air Force (USAF) established the Aircraft Structural Integrity Program (ASIP) in November
1958 in response to in-flight structural failures resulting in five destroyed B-47 aircraft from March through
April 1958 [1]. Four of the B-47 losses were attributed to fatigue, which led to a probabilistic approach
for establishing the aircraft service life capability. This was called the “safe-life” approach, and it relied
upon the results of a laboratory test of a full-scale airframe subjected to loading that simulated the operational
service environment of the aircraft. The USAF established the safe life of the aircraft by dividing the number
of successfully test simulated flight hours by a scatter factor. The intent of the factor was to account for
aircraft-to-aircraft variation in materials and manufacturing quality. The USAF believed the process to be
sufficient to preclude in-service structural failures attributable to fatigue. The safe life approach was the basis
for all new designs during the 1960s and was also used to establish the safe life of earlier designs that were
subjected to a fatigue test. Losses of an F-111 in December 1969 and an F-5 in April 1970 [1], each far short of
their qualified safe life, demonstrated that the safe life approach had shortcomings. The safe life approach
allowed the use of low ductility materials operating at high stresses, which resulted in designs that were
intolerant to manufacturing and service-induced defects. The aircraft failures arising from the deficiencies of
the safe life approach demanded a fundamental change be made in the design, qualification, and inspection of
aircraft. The damage tolerance approach emerged as the candidate chosen for this change.
Developers of the damage tolerance approach recognized that an aircraft’s structure is subject to a wide
range of initial quality from manufacturing processes as well as from service-induced damage during
operations and maintenance. They also recognized that the aircraft structure had to be inspectable. To ensure
the aircraft operates safely in the presence of anomalies, the USAF requires the structure to tolerate the
defects for some inspection-free period of service usage. The damage tolerance approach provides the USAF
a safety limit for each critical area in the aircraft. The safety limit is the time, in flight hours, required for a
crack to grow from either an assumed initial flaw size, or the inspectable flaw size, to a critical size.
Inspections are scheduled to occur at a time equal to one-half the determined safety limit. The USAF used
the damage tolerance approach to upgrade the structural integrity of several operational aircraft in the early
1970s including the F-111, C-5A, and F-4. The success of these efforts convinced the USAF that damage
tolerance should be the structural safety basis for all future designs. In December 1975, the USAF formally
integrated the damage tolerance approach into the ASIP. During the 1970s and 1980s, the USAF performed
a damage tolerance assessment on every major aircraft weapon system to develop inspection or modification
programs necessary to maintain operational safety. The results of the damage tolerance assessments were
incorporated into USAF Technical Orders which established maintenance requirements to maintain
structural integrity and to control risk to an acceptable level.
STO-TR-AVT-275 5-1
AIRWORTHINESS OF STRUCTURES (UNITED STATES)
1) The flight limits (Mach, altitude, rates, accelerations, etc.) that were used to establish the
initial airworthiness certification are documented in the flight manual Technical Order (TOs) or
similar product;
2) The maintenance requirements to maintain airworthiness are established and documented in TOs,
depot work specifications, and other products as needed;
3) A force management method is developed and implemented prior to initiation of operations to:
determine actual usage for the fleet and each aircraft, adjust planned inspection intervals for each
individual aircraft as needed, update analysis as required, perform surveillance inspections, and
revise the maintenance requirements as needed; and
4) Finally, a Service Life Limit (SLL) is established and documented in the flight
certificate/authorization.
5-2 STO-TR-AVT-275
AIRWORTHINESS OF STRUCTURES (UNITED STATES)
2) Implement modifications based on service, analysis and test results to increase the SLL as required; and
3) Document the revised SLL in each aircraft’s flight authorization when modifications (if required)
are implemented.
1.E-03
1.E-05
1.E-06
1.E-07
1.E-08
1945
1950
1955
1960
1965
1970
1975
1980
1985
1990
1995
2000
2005
2010
2015
2020
6.2 REFERENCES
[1] ASC-TR-2010-5002, “Threats to Aircraft Structural Safety, Including a Compendium of Selected
Structural Accidents/Incidents, March 2010.
[2] MIL-STD-1530D, Change 1, “Department of Defense Standard Practice – Aircraft Structural Integrity
Program (ASIP)”, October 2016.
STO-TR-AVT-275 5-3
AIRWORTHINESS OF STRUCTURES (UNITED STATES)
5-4 STO-TR-AVT-275
Chapter 7 – APPROACH TO STRUCTURAL
RISK ANALYSIS (USAF)
In a fracture mechanics approach to probabilistic risk analysis, failure is defined as the probability that the
maximum stress encountered in a flight will produce a stress intensity factor at an FCL that exceeds the
fracture toughness of the material. Figure 7-1 is a simplified schematic of how a distribution of initial flaw
sizes is projected to grow along a deterministic crack growth curve as a function of time. At time T1, some
proportion of this distribution is greater than the critical crack size, i.e., the crack size at which the fracture
toughness is exceeded (which may or may not be deterministic as shown here). This proportion translates
into probability of failure at T1.
Figure 7-1: Schematic for Projecting Crack Size Distribution as a Function of Flight Hours [1].
STO-TR-AVT-275 7-1
APPROACH TO STRUCTURAL RISK ANALYSIS (USAF)LYNNE DEMMERY
associated with material properties, usage, inspection, and repair. Program inputs included material and
geometry information, aircraft usage, Non-Destructive Evaluation (NDE) information, and repair
capabilities. It should be noted that many current USAF structural risk analyses are based on the original
Lincoln and PROF methods, but are not limited to them.
Table 7-1 lists typical inputs that are used in a risk analysis as well as possible sources of the data. The
following sections will explain each input in detail.
7-2 STO-TR-AVT-275
APPROACH TO STRUCTURAL RISK ANALYSIS (USAF)LYNNE DEMMERY
where:
σ = the reference stress;
β = a geometry correction factor that depends on the size and shape of the crack; and
a = the relevant crack size dimension, usually crack depth or surface crack length.
Geometry correction factors (β) are obtained from the crack growth model and a normalized stress intensity
factor, α(a) or “alpha”, is defined as:
𝐾𝐾(𝑎𝑎)
𝛼𝛼(𝑎𝑎) = = 𝛽𝛽√𝜋𝜋𝜋𝜋 (7-2)
𝜎𝜎
Risk analysis software generally requires the input of the stress intensity factor in this normalized form using
a data table of crack length and corresponding alpha. Since this form is independent of stress, the same alpha
curve theoretically represents any aircraft usage.
A common method for determining an initial crack size distribution is with the Equivalent Initial Flaw Size
(EIFS) method as outlined in the 1989 USAF Durability Design Handbook [3]. According to the Handbook,
the EIFS is a hypothetical crack that is meant to capture the equivalent effect of actual material
imperfections such as random scratches, microscopic flaws, or other damage types. An EIFS distribution is
determined by taking a population of crack sizes with known usage and using the crack growth curve to
determine their sizes at time 0.
EIFS data should come from teardown data of the FCL of interest, when possible. See Ref. [2] for methods
of determining the EIFS distribution from teardown data. EIFS data from other sources should only be used
as a last resort and should, at minimum, be of the same material and geometry (such as an edge detail or
hole detail). The material form and manufacturing process can also have a significant effect on the
EIFS distribution.
STO-TR-AVT-275 7-3
APPROACH TO STRUCTURAL RISK ANALYSIS (USAF)LYNNE DEMMERY
It is also possible to use a probabilistic crack growth analysis that varies crack growth rate and geometry
parameters. Due to the increased complexity of the analysis, Monte Carlo and other sampling methods
are recommended.
The PROF manual describes a procedure for fitting a Gumbel distribution to exceedance data [1]. First, the
exceedance curve must first be converted to the distribution function, Fall(σ):
Fall(σi) = 1 – λ(σi) / λ(σthr) (7-4)
where λ(σi) is the number of peak stresses per unit time exceeding σi and λ(σthr) is the number of
exceedances per unit time of the stress threshold. Let n represent the average number of stress peaks per
flight greater than threshold:
n = (# of peaks in spectrum) / (# of flights in spectrum)
Then, the cumulative distribution of the maximum stress per flight is estimated by:
n (7-5)
Fmax(σi) = [1 – λ(σi) / λ(σthr)]
A least squares fit of the (σi , ln{-ln[H(σi)]}) data pairs will yield estimates of -l/A and B/A. To ensure that
the fit is acceptable at the high stress levels of most influence in the hazard computation, only the highest
stress ranges in the data should be used in determining the least squares fits. It might be noted that B is the
stress that is exceeded in 63 percent of the flights and A is proportional to the steepness of the exceedance
probability vs. stress curve. The larger the value of A, the flatter the exceedance probability curve (resulting
in a larger probability of large maximum stress peaks in a flight).
A practical approach to estimating A and B is to vary these parameters until an acceptable fit is obtained for
the probability of exceeding the high stress levels which drive POF calculations. Engineering judgement is
required when selecting the appropriate fit. It is easily possible to have a large estimate of A that can result in
a probability of encountering a practically impossible stress. If these artificially high stress estimates are
found to drive up the POF, a Weibull exceedance curve may be more suitable.
7-4 STO-TR-AVT-275
APPROACH TO STRUCTURAL RISK ANALYSIS (USAF)LYNNE DEMMERY
Structures Bulletin EN-SB-08-012, Rev. D, defines the recommended Non-Destructive Inspection (NDI)
capability flaw sizes for USAF aircraft structures when no other supporting data is available [4].
Customarily, the capability of an inspection system is quantified by a90, the crack length at which the POD is
0.90. A full POD curve can be developed using the a50 and a90 values from EN-SB-08-012, Rev. D.
An analytical approach for computing cumulative or single flight probability of failure is based on the
concept of conditional expectation [8]. In this method, probability of failure is expressed analytically in
terms of one random variable, known as the control variable, while all others are held constant. Then, the
overall probability is computed by integrating over the other random variables. In probabilistic terms, this is
taking the expectation of the conditional probability over the other random variables. This approach enables
formulation of failure probability as a tractable, low-dimensional integral, while still taking proper account of
maximum stress per flight.
First, consider the cumulative probability of failure up to flight t. This represents the distribution of
time-to-failure as a function of flights. The control variable will be maximum stress per flight, so the
probability of failure conditional on initial crack size and fracture toughness is defined here first. The
cumulative probability of failure during the first t flights is equal to one minus the probability of surviving
STO-TR-AVT-275 7-5
APPROACH TO STRUCTURAL RISK ANALYSIS (USAF)LYNNE DEMMERY
the first t flights. Given the initial crack size a0 and fracture toughness Kc, the probability of surviving flight t
is simply the probability that the maximum stress for that flight will be less than Kc/α:
(7-7)
where H is the cumulative maximum stress distribution, and 𝛼𝛼(𝑎𝑎0 , 𝑡𝑡)is the normalized stress intensity factor
for the crack at time t obtained from a damage tolerance analysis and deterministic crack growth curve. For a
given a0 and Kc, the only random variable in the problem is maximum stress per flight. Since the stress for
each flight is independent, failure on each flight is independent (given a0 and Kc). As a result, the probability
of surviving the first t flights, conditional on a0 and Kc, can be expressed as:
(7-8)
(7-9)
Equation (7-1) to (7-9) gives the cumulative failure probability conditional on particular values of initial
crack size and fracture toughness. In order to introduce these as random variables, take the expectation of
Equation (7-9), which is a double integration:
(7-10)
where f(a0) and g(Kc) are the PDFs of the initial crack size and fracture toughness, respectively.
Next, consider the single flight failure probability. The quantity of interest is the probability of failure during
flight t, given that the aircraft survives the first t-1 flights. This requires a conditional probability, written in
conceptual terms as:
survival to t-1 AND failure at t
P(failure at t|survival to t-1) =
survival to t-1
(7-11)
This could be computed efficiently in terms of Equation (7-10) by performing the numerical integration on
two integrands simultaneously: with and without flight t in the product. This would allow one to obtain both
t −1
Kc
POF(t) and POF(t-1) without having to re-compute ∏ H α (a , i ) at each integration point.
i =1 0
SFPOF here is formulated as a post-processing operation on the POF. Therefore, calculating the POF for the
entire period of interest also allows calculation of the SFPOF. Note that versions of PROF released prior to
2015 use a different SFPOF formulation that does not account for prior failures and may produce overly
conservative results.
7-6 STO-TR-AVT-275
APPROACH TO STRUCTURAL RISK ANALYSIS (USAF)LYNNE DEMMERY
The USAF ASIP guidelines and standards reflect an emphasis on risk analyses of aircraft structural integrity.
However, the more generalized MIL-STD-882E document does not explicitly relate its probability levels to
the singular FCL SFPOF values computed using risk analysis software. To provide guidance in the
consistent use of MIL-STD-882E for airworthiness issues, the Air Force Life Cycle Management Center
(AFLCMC) recently issued Airworthiness Bulletin 150 [10]. This bulletin lists ranges of probability of
aircraft loss per flight hour and the corresponding risk acceptance authority. The risk assessment matrix is
reproduced in Table 7-2.
STO-TR-AVT-275 7-7
APPROACH TO STRUCTURAL RISK ANALYSIS (USAF)LYNNE DEMMERY
7.5 SUMMARY
Probabilistic risk analysis for aircraft structures is a powerful tool that can take into account geometric,
material, and usage variabilities. In addition, risk analyses can be useful for comparing different inspection
intervals and methods. The effects of different repair options may also be explored to determine a reasonable
trade-off between repair types and their direct effect on probability of failure. Although the risk analysis
methods are rigorous, the analyst must keep in mind the quality and veracity of the input data.
Over-conservatism can lead to costly additional and unnecessary inspections, while under-conservatism
could lead to catastrophic failure. The analyst should also keep in mind what variables are not included in the
risk analysis; unknowns are often the downfall of an otherwise acceptable analysis.
7.6 REFERENCES
[1] Miedlar, P., Berens, A., Hovey, P., Boehnlein, T., and Loomis, J., “Aging Aircraft Risk Analysis
Update”, Dayton, OH: 2005.
[2] Domyancic, L.C., “Methods of Determining Equivalent Initial Flaw Size (EIFS) Distributions
Containing Suspended Data”, AIAA SciTech Forum, Kissimmee, FL, 2015.
[3] Manning, S.D., and Yang, J.N., “USAF Durability Design Handbook: Guidelines for the Analysis and
Design of Durable Aircraft”, Wright-Patterson Air Force Base, 1989.
[4] EN-SB-08-012, Rev. D, “In-Service Inspection Crack Size Assumptions for Metallic Structures”,
AFLCMC/EZ, April 2018.
[5] Lincoln, J.W., “Risk Assessment of an Aging Military Aircraft”, Journal of Aircraft, Vol. 22, No. 8,
1985, pp. 687-691.
7-8 STO-TR-AVT-275
APPROACH TO STRUCTURAL RISK ANALYSIS (USAF)LYNNE DEMMERY
[6] Lincoln, J.W., “Method for Computation of Structural Failure Probability for an Aircraft”,
ASD-TR 80-5035, July 1980.
[7] Domyancic, L., McFarland, J., Cardinal, J., and Burnside, O., “Review of Methods for Single Flight
Probability of Failure”, Southwest Research Institute White Paper, March 2011.
[8] Haldar, A., and Mahadevan, S., “Variation Reduction Techniques”, Probability, Reliability and
Statistical Methods in Engineering Design, John Wiley & Sons, Inc., 2000.
[9] MIL-STD-882E, “Department of Defense Standard Practice for System Safety”, 11 May 2012.
[10] “Airworthiness Risk Assessment and Acceptance,” Bulletin AWB-150, USAF AFLCMC,
13 September 2017.
STO-TR-AVT-275 7-9
APPROACH TO STRUCTURAL RISK ANALYSIS (USAF)LYNNE DEMMERY
7 - 10 STO-TR-AVT-275
Chapter 8 – CONTINUING AIRWORTHINESS
OF STRUCTURES (CANADA)
8.1 INTRODUCTION
This chapter presents a brief history of the Royal Canadian Air Force (RCAF) continuing airworthiness
(CAW) process on aircraft structures. It also provides the latest RCAF references related to the CAW on
aircraft structures, as well as the lessons learned and future challenges.
Under the provisions of the Aeronautics Act, the Canadian Minister of National Defence (MND) is responsible
for military aviation, including foreign aircraft within Canada. Within the Department of National Defence
(DND), the Technical Airworthiness Authority (TAA) is responsible for the regulation of the technical
airworthiness aspects of design, manufacture, maintenance and material support of aeronautical products and
the determination of the airworthiness acceptability of those products prior to operational service. The technical
airworthiness program, which is outlined in the Technical Airworthiness Manual (TAM) [1], consists of three
primary elements:
• Initial Airworthiness,
• Continuing Airworthiness, and
• Disposal.
A key element of the DND’s CAW requirement is for an In-Service Monitoring Program (ISMP). As a
minimum, the ISMP covers the monitoring activities for airworthiness directives, service bulletins, flight
safety occurrences, other operator experiences and aircraft usage and condition monitoring. The Aircraft
Structural Integrity Program (ASIP) falls under the ISMP.
Starting with the procurement of the CH146 helicopter (Bell 412) in 1994, the DND no longer required the
inclusion of the aircraft design engineering data with the acquisition of the aircraft. As such, the information
required under Tasks I, II and III of the MIL-STD-1530 was either very limited or not available to the RCAF.
Around the same time the DND started to formalize their technical airworthiness program and the assembly
of the Technical Airworthiness Manual. As it was understood that the DND would have very limited
influence on and access to the design information (i.e., design analysis and full-scale testing), it was decided
to NOT mandate the MIL-STD-1530 in its entirety. Rather, it was decided to focus on those tasks for which
the DND had control and would affect the CAW of the aircraft. This was brought under the In-service
Monitoring Program chapter of Part 3 – Continuing Airworthiness of the TAM. Several annexes were
included to cover the various systems of the aircraft. One of these annexes is the Aircraft Structural Integrity
Monitoring Requirements.
STO-TR-AVT-275 8-1
CONTINUING AIRWORTHINESS OF STRUCTURES (CANADA)
The ASIMP, along with providing an overview of the aircraft design and certification activities, outlines the
tasks performed under the first four key elements.
The RCAF has very diverse aircraft fleets with varying operational roles. As such, the level of details
required under the monitoring requirement varies from fleet to fleet. Aircraft for which the structural
integrity is manoeuvre dependent requires a monitoring system that will take into account every manoeuvre
flown. Whereas, aircraft that fly well-defined profile missions, such as troop transport aircraft, may only
require tracking of flight hours and ground-air-ground cycles. Similarly, fleets that follow a civilian
maintenance program will rely more on the Original Equipment Manufacturer (OEM) for structural
condition monitoring and aging aircraft structural assessment as the OEM will be able to pool their
experience from other fleet operators.
The RCAF aircraft fleets have been certified using different airworthiness standards. A summary of the
RCAF aircraft structural lifing policies, including original OEM and current RCAF methods, is presented in
Table 8-1 (updated from a Table from Ref. [3]). Table 8-1 indicates that two major lifing methods, i.e., Safe
Life and Damage-Tolerant, are used across the RCAF fleets, either in a separate or combined manner.
Table 8-1: Summary of RCAF Aircraft Structural Lifing Policies, Original and Current.
8-2 STO-TR-AVT-275
CONTINUING AIRWORTHINESS OF STRUCTURES (CANADA)
STO-TR-AVT-275 8-3
CONTINUING AIRWORTHINESS OF STRUCTURES (CANADA)
At the time of the gap analysis, the Aging Aircraft Structural Assessment requirement was not part of the
TAM. The requirement for an Aging Aircraft Assessment (AAA) was introduced under the aircraft usage
and condition monitoring requirement in June 2012 with release Change 6 of the TAM. At that time it was to
apply to aircraft structures and systems. However, at Change 6 of the TAM, the Annexes describing the
monitoring requirement for the Mechanical Systems, Propulsion Systems, and Aircraft Electrical Wiring
Interconnection Systems had not been populated. As all of the AAAs activities at DND were related to the
aircraft structure, Change 7 of the TAM only included the detail requirements specified for an Aging Aircraft
Structural Assessment (AASA).
For fleets that have well-defined repetitive mission profile, such as a passenger/troop transport aircraft, the
requirement can be simplified to the collection of flight hour, pressure cycles and ground-air-ground cycles
resulting in a very simple SUM program. However, for aircraft that have heavily manoeuvre based missions,
a more robust and detail SUM program is required. In this case a manoeuvre recognition algorithm based
SUM system is required that can account for the fatigue life consumption for safety-of-flight critical
components. To make thing more interesting a hybrid system may be required for aircraft that have distinct
dual roll such as the RCAF CC150, which has two configurations. One configuration is solely as a passenger
aircraft and second configuration as an air-to-air refueller. In the air-to-air configuration, all refuelling
exercises must be tracked and the associated fatigue life damage accounted for.
All of the RCAF legacy fleets have a TAM compliant SUM program, while the newly acquired fleets are in
the process of meeting the TAM SUM requirements.
8-4 STO-TR-AVT-275
CONTINUING AIRWORTHINESS OF STRUCTURES (CANADA)
1) There shall be a program implemented to monitor structural repairs and inspection records.
The program shall apply at least to primary structure (and dynamic components for helicopters);
2) The structural condition monitoring program shall include long-term impact assessment of structural
repairs at or near critical locations and other repairs;
3) The structural condition monitoring program shall include review of inspection records to determine
trends in the frequency of fatigue damage in specific locations;
4) The structural condition monitoring program shall include review of service difficulty reports (or
equivalent) to detect any difficulty that could impact the long-term aircraft structural integrity; and
5) A process shall be implemented to re-evaluate the maintenance program (including corrosion
prevention and control procedures) as a result of detrimental trends identified through structural
condition monitoring.
For the RCAF fleets meeting the above requirement is not difficult when tracking fleet-wide modifications
or non-standard repairs, as these are often very well documented due to the nature of the design change.
What is difficult to track are those repairs that are covered by Standard Repair Manual (SRM). Often the
record set for these repairs have very limited information. Routinely, the only comment available in the
maintenance record is that the damage was repaired in accordance to the repair ‘XX’ of the SRM. Little, if
any, information is available on size of damage, making analysis of repair interaction and damage trends
difficult. As a result many of the RCAF fleets are not fully compliant to the TAM SCM requirements.
The requirement for an Aging Aircraft Assessment (AAA) was initially introduced under the aircraft usage
and condition monitoring requirement in June 2012 with release Change 6 of the TAM. At that time it was to
apply to aircraft structures and systems. However, at Change 6 of the TAM the Annexes describing the
monitoring requirement for the Mechanical Systems, Propulsion Systems, and Aircraft Electrical Wiring
Interconnection Systems had not been populated. As all of the AAAs activities at DND were related to the
aircraft structure, Change 7 of the TAM only included the detail requirements specified for an Aging Aircraft
Structural Assessment (AASA).
The AASA are to be carried out no later than 15 years after the production of the first aircraft delivered to the
RCAF. As of Dec 2017, all of the RCAF fleets that have passed the 15 year mark have undergone an AASA.
Similar to the SUM and SCM requirements, the approaches taken to perform the AASA has differed based
on the fleet and the Aircraft Structural Integrity Program (ASIP) for that fleet. For the fleets that
pronominally follow a civilian role, such as the CC144 (Bombardier Challenger CL601/604 variants) and
CC150 (Airbus A310), AASA relied heavily on the work done by the OEM under the Federal Aviation
Administration (FAA) mandated program for Aging Airplane Safety [4]. For the predominately military role
fleets a combination on physical audit and review of the structural maintenance records were performed as
part of the AASA.
The level of effort to perform the AASAs was largely dependent on the strength of the ASIP for the fleet in
question. Fleets that had a strong ASIP required little effort to perform the AASA as there was already very
good documentation on the actual usage and physical condition of the airframe. Overall, the RCAF fleets are
STO-TR-AVT-275 8-5
CONTINUING AIRWORTHINESS OF STRUCTURES (CANADA)
well maintained and no significant issues were raised during the AASAs. This is largely due to having a
healthy ASIPs for the RCAF fleets. This is not to say that some of the aircraft did not have corrosion issues,
but these were already being monitored and addressed as part of the ASIP.
8.9 REFERENCES
[1] Department of National Defence of Canada, Technical Airworthiness Manual (TAM), Document No.
C-05-005-001/AG-001, 2015, Change 7, 23 November 2015.
[2] USA Department of Defense Standard Practice, Aircraft Structural Integrity Program (ASIP),
MIL-STD-1530D, 31 Aug 2016.
[3] Liao, M., Renaud, G. and Bombardier, Y., “Application of Quantitative Risk Analysis in Support of
Continuing Airworthiness Management”, The 28th Symposium of International Committee on
Aeronautical Fatigue and Structural Integrity (ICAF 2015), Helsinki, Finland, June 2015.
[4] US Department of Transportation, Federal Aviation Administration, 14 CFR Parts 119, 121, 129, 135,
and 183, Aging Airplane Safety; Final Rule, 2 February 2005.
8-6 STO-TR-AVT-275
Chapter 9 – AUSTRALIAN STRUCTURAL
INTEGRITY APPROACH
9.1 INTRODUCTION
A Joint Directive signed by Secretary, Department of Defence and Chief of the Defence Force established
the Defence Aviation Safety Framework as of 30 Sep 2016. The framework required the implementation of a
credible and defensible aviation safety framework that recognises and supports compliance with statutory
safety obligations. Where appropriate the framework was aligned with International Civil Aviation
Organisation (ICAO) principles and European Military Airworthiness Requirements (EMAR). As at
01 Jan 19, the ADF fully transitioned to the Defence Aviation Safety Regulation (DASR) [1].
Under the previous Technical Regulations, the Australian Defence Force (ADF) managed Aircraft Structural
Integrity (ASI) via a MIL-STD-1530 Aircraft Structural Integrity Program (ASIP), documented within
an Authority approved Aircraft Structural Integrity Management Plan (ASIMP). Under the DASR, the
ASIP construct was preserved for existing weapon systems and established as Acceptable Means of
Compliance (AMC) for new acquisitions (as per DEFLOGMAN Part 2, Volume 10, Chapter 18 and
Airworthiness Design Requirements Manual (ADRM) Section 3 Chapter 12 [2]). Responsibility for the
execution of ASIP elements is divested amongst the Part 21 Initial Airworthiness, Part M Continuing
Airworthiness and Part 145 Maintenance regulations, supplemented by ADRM Section 3 Chapter 12
Aircraft Structural Integrity.
As per MIL-STD-1530D [3], the ASIP is composed of the following primary tasks:
• Task I (Design Information). Task I consists of establishing those criteria and other requirements
which must be applied during design to ensure the overall program goals will be met;
• Task II (Design Analyses and Development Testing). Task II consists of the characterisation of
the environment in which the aircraft must operate, the testing of materials, components, and
assemblies, and the analyses of the aircraft design;
• Task III (Full-Scale Testing). Task III consists of laboratory and flight tests of the aircraft structure
to assist in determining the structural adequacy of the analysis and design;
• Task IV (Certification and Force Management Development). Task IV consists of the analyses
that lead to certification of the aircraft structure as well as the development of the processes and
procedures that will be used to manage force operations when the aircraft enters the inventory; and
• Task V (Force Management Execution). Task V consists of the execution of the processes and
procedures developed under Task IV to ensure structural integrity throughout the life of each
individual aircraft. This task may involve revisiting elements of earlier tasks, particularly if the
service life requirement is extended, if the operational usage is different than the design spectrum, or
if the aircraft is modified.
Figure 9-1 demonstrates the effectiveness of the ASIP approach, implemented in ADF airworthiness
regulations since 1994.
STO-TR-AVT-275 9-1
AUSTRALIAN STRUCTURAL INTEGRITY APPROACH
ACRONYMS
ASMS Aviation Safety Management System DASM Defence Aviation Safety Manual
AVRM Aviation Risk Management DASP Defence Aviation Safety Program
DASA Defence Aviation Safety Authority DASR Defence Aviation Safety Regulation
9-2 STO-TR-AVT-275
AUSTRALIAN STRUCTURAL INTEGRITY APPROACH
As per the ADRM it is essential for Initial Airworthiness of new aircraft types that flight and ground loads tests,
as well as static and durability tests compliment certification by analysis. During Initial Airworthiness
activities, an Aircraft Structural Integrity Documentation Package (ASIDP) is required to summarise the final
analysis and other relevant structures information that will provide the basis for ADF type certification, and
describe important characteristics, limitations and capabilities information for in-service management agencies.
Aircraft procured by the ADF are designed and certified to a wide range of standards such as UK and US
military standards and civil Federal Aviation Regulations (FARs) and Joint Aviation Requirements (JARs).
The ADF therefore accepts a range of structural design standards for type certification and continuing
airworthiness. Seldom, however, is structural verification and certification undertaken with ADF
Configuration, Role and Environment (CRE) in mind. The ADF must therefore, at the time of acquisition,
identify the structural design standard applied during aircraft development; assess the adequacy of that
standard and the adequacy of structural verification and certification for ADF CRE.
An Aircraft Structural Integrity Management Plan (ASIMP) is established during Initial Airworthiness
activities to document the ASIP and sources of the in-service data pack required for ASI management
activities. The ASIMP is a proactive plan that addresses the ASIP activities required to be undertaken over
the short, medium and long-term of an aircraft type’s service life with particular emphasis on the next three
to five years. Throughout the sustainment phase, the ASIMP will reflect the activities necessary to assure
that structural integrity is maintained at the level established at introduction to service.
The aircraft Maintenance Organisation is responsible for the following ASIP elements:
• Execute structural management policy in accordance with Aircraft Maintenance Program;
• Support operational loads measurement program;
• Support collection of structural (usage and condition monitoring) data, including occurrence
reporting for critical structure; and
• Support aging aircraft structural assessments (aircraft inspections).
STO-TR-AVT-275 9-3
AUSTRALIAN STRUCTURAL INTEGRITY APPROACH
Outputs from the ASI Assessment Cycle are used to ensure that inspection intervals, SLLs and CRTs from
initial type certification remain valid. This information is used to verify the structural Life of Type (LOT)
and remaining economic life of aircraft fleets relative to Planned Withdrawal Date (PWD). The systems
which underpin the ASI Assessment Cycle are documented in the ASIMP for each aircraft type.
9-4 STO-TR-AVT-275
AUSTRALIAN STRUCTURAL INTEGRITY APPROACH
retired aircraft and component where possible in lieu of traditional teardowns and where a Full-Scale
Durability Test (FSDT) is judged unnecessary to achieve the required life extension.
An enhanced teardown differs from traditional full-scale certification testing in at least two key aspects:
• The test article must be a retired fleet asset with the structural degradation arising from service usage
‘locked in’ for identification through further testing and forensics; and
• Limitations in loads fidelity are accepted from the onset if testing remains fit for purpose and can be
simplified and/or accelerated as a result.
The primary benefit of the enhanced teardown approach over conventional teardowns or age exploration
inspections of high time airframes [3] the improved probability of detection of production and in-service
defects. Correctly sized for fatigue, a structure designed to either ‘Crack Initiation’ or ‘Slow Growth’
requirements should not exhibit detectable evidence of fatigue crack growth within one service lifetime. That
is, fatigue cracking in well-designed and certified structure will remain below conventional Non-Destructive
Testing (NDT) detection thresholds and should not be identified in-service. These thresholds are typically of
the order of 1 – 6 mm (0.04 – 0.25 inches) length for surface breaking anomalies when using conventional
eddy current inspection methods in aluminium alloys [11], [12]. The advantage of the enhanced teardown
concept is that, through structural experimentation post airframe retirement, flaw sizes can be grown to and
beyond detectable limits, including to failure of the structure. Optimum evidence for forensic analysis,
Quantitative Fractography (QF) and subsequent use in Aircraft Structural Integrity (ASI) management is
therefore developed from an enhanced teardown. Further it is good practice to investigate the robustness of
an aircraft ASIP through mid-life teardown inspections. As Australia has shown the enhanced teardown
delivers the most benefit in a relatively economical fashion.
9.5 REFERENCES
[1] Defence Aviation Safety Regulation (DASR), Release 10 October 2019. http://www.defence.gov.au/
DASP/Docs/Manuals/8000-011/DASRWeb/.
[2] Australian Air Publication, Airworthiness Design Requirements Manual (ADRM), AAP 7001.054,
13 November 2019. https://www.defence.gov.au/DASP/Docs/Manuals/7001054/ADRMWeb/.
[3] Department of Defense Standard Practice, Aircraft Structural Integrity Program (ASIP),
MIL-STD-1530D, CHANGE 1, 13 October 2016. http://quicksearch.dla.mil/Analyse/ImageRedirector.
aspx?token=5743923.36952.
[4] Defence Aviation Safety Authority, Introduction to Defence Aviation Safety Guidebook, March 2018.
https://www.defence.gov.au/DASP/Docs/Media/IntroductiontoDefenceAviationSafetyGuidebookFeb20
19.pdf
STO-TR-AVT-275 9-5
AUSTRALIAN STRUCTURAL INTEGRITY APPROACH
[5] Molent, L., Dixon, B., Barter, S., White, P., Mills, T., Maxfield, K., Swanton, G., and Main, B., (2009)
Enhanced Teardown of Ex-Service F/A-18A/B/C/D Centre Fuselages, In: Proceedings of the
25th International Conference on Aeronautical Fatigue (ICAF) Rotterdam, Netherlands, pp. 123-143.
[6] Molent, L., Barer, S.A., Dixon, B. and Swanton, G., (2018) Outcomes from the Fatigue Testing of
Seventeen Centre Fuselage Structures, International Journal of Fatigue, Vol 111, pp. 220-232.
[7] Fitzgibbon, J. (2008) Defence Expertise Saves $400M on Hornet Upgrade, Department of Defence
Media Release, 4 September 2008.
[8] Main, B., Muller, K., Konak, M., Jones, M., Sudhakar, S., and Barter, S., (2020) Evaluation of a
PC-9/A Wing Main Spar with Misdrills Using Enhanced Teardown at Resonance. In: A.
Niepokolczycki and J. Komorowski (Eds.): ICAF 2019 – Structural Integrity in the Age of Additive
Manufacturing, pp. 874-888.
[9] Van Blaricum, T., Ord, D., Koning, R. and Hobson, M., (2004) Damage Enhancement Testing of RAAF
F-111C Wing A15-5, DSTO-TR-1460, Defence Science and Technology, Melbourne.
[10] Grooteman, F. and Bos, M., (2015) Advanced Life Assessment of the Lead Crack Configuration of the
RNLAF F-16 Wing Damage Enhancement Test, NLR-TP-2015-140, National Aerospace Laboratory
(NLR), Netherlands.
[11] Reams, R.H. and Babish, C.A. IV, (2013) In-Service Inspection Flaw Assumptions for
Metallic Structure, EN-SB-08-012 Rev C, Airforce Life Cycle Management Centre (AFLCMC),
Wright-Patterson Air Force Base (WPAFB) Ohio, USA.
[12] Australian Air Publication. (1999) Non Destructive Testing General Procedures AAP 7002.043-36,
AL 35, 14 May.
9-6 STO-TR-AVT-275
Chapter 11 – IMPLEMENTATION OF THE AIRCRAFT
STRUCTURAL INTEGRITY PROGRAM PRINCIPLES
FOR PZL-130 “ORLIK” TC-II TRAINER AIRCRAFT
Piotr Reymer
Air Force Institute of Technology
POLAND
11.1 INTRODUCTION
The Air Force Institute of Technology (AFIT) has been involved in the SEWST (Introduction of Damage
Tolerance Based Operation, SEWST) research program for the PZL-130 “Orlik” TC-II Polish trainer aircraft.
The aim of the program was to establish a new fatigue life limit for the modernized structure (refurbished
fuselage and brand-new set of wings) as well as to introduce a new maintenance system incorporating the
Aircraft Structural Integrity Program (ASIP) principles. The SEWST program was led by EADS PZL-Okęcie,
the manufacturer of the aircraft, however AFIT was responsible for most of the research activities.
The SEWST program was carried out throughout five years (2010 – 2014). The program was contracted by
the Polish MOD (Ministry of Defence) and led by EADS PZL-Okęcie, the manufacturer of the aircraft. The
AFIT was the main subcontractor and was therefore responsible for most research activities. The total cost of
the program was 7 million USD, whereas the estimated savings after introduction of the program are around
70 million USD.
The SEWST program consisted of several activities among which the most important were:
• Operational Load Monitoring;
• Load Spectrum Development;
• Full-Scale Fatigue Test (FSFT) (including Non-Destructive Inspection (NDI) program);
• Teardown Inspection (TI); and
• Individual Aircraft Tracking (IAT) Program.
The main task of the structural part of the program was to carry out the FSFT. In order to achieve this goal a
set of subtasks had to be performed. Beginning with the definition of the flight profile of the PZL-130
aircraft operated in the Polish Air Force [1], followed by the Operational Load Monitoring Program (OLM),
the load spectrum for the FSFT was defined [2]. During the OLM as well as the FSFT, two aircraft were
instrumented with sets of over 100 strain gauges. The aircraft used in the FSFT was a TC-I version
withdrawn from service, overhauled to TC-II version by reinforcing the fuselage and installing a brand-new
pair of modernized wings. The TC-II version wings have a larger span as well as wing area. The overhaul
carried out on the FSFT test specimen is identical to the modernization which all the TC-II aircraft undergo.
STO-TR-AVT-275 11 - 1
IMPLEMENTATION OF THE
AIRCRAFT STRUCTURAL INTEGRITY PROGRAM
PRINCIPLES FOR PZL-130 “ORLIK” TC-II TRAINER AIRCRAFT
11 - 2 STO-TR-AVT-275
IMPLEMENTATION OF THE
AIRCRAFT STRUCTURAL INTEGRITY PROGRAM
PRINCIPLES FOR PZL-130 “ORLIK” TC-II TRAINER AIRCRAFT
The wing bending moment was measured in three sections of each wing. For each section, four strain gauges
were installed (two on the upper and lower flange of the front spar and two on the upper and lower flange of
the rear spar). Fuselage bending moment was measured in two sections (front near Frame no 1 and in the rear
near Frame no 9), in each section the bending moment was measured by means of strain gauges installed on
the stringers. Moreover, in the rear fuselage sections, two rosettes were installed on the sheeting in order to
measure tail section torque. The empennage loads were measured with strain gauges installed in both
horizontal and vertical stabilizers root sections. Finally, landing loads (vertical load and bending) were
measured by sensors installed both on main landing gear as well as the support structure.
During the test flight program, 19 carefully planned flights were carried out, which consisted of selected
flight exercises and additional maneuvers (like spins) in order to gather necessary and representative load.
The load monitoring system was calibrated before flights by applying known loads to the structure and
measuring the corresponding strain signals.
In addition to the above flight test program, several additional flights carried out by military pilots after the
aircraft was handed over to the Air Force Base, allowing the comparison of test flight severity with actual
service flights.
The assumed load blocks were to represent four main flight components: taxiing, flight, landing and buffeting.
Flight and buffeting loads were developed directly using strain data gathered during flights. Since landings
performed during the experimental flight program were very smooth, which is usually not the case, a decision
STO-TR-AVT-275 11 - 3
IMPLEMENTATION OF THE
AIRCRAFT STRUCTURAL INTEGRITY PROGRAM
PRINCIPLES FOR PZL-130 “ORLIK” TC-II TRAINER AIRCRAFT
was made to develop landing loads separately according to known literature [3]. Similarly, taxiing loads
showed very low amplitudes, and due to limitations in number of load lines within a load block (discussed in
following paragraph), could not be introduced directly. Instead, three types of artificial taxiing loads were
developed: full stop, left and right turn. These were randomly placed within the load block. The load levels
were chosen based upon the actual loads measured during experimental flights.
The time schedule of the FSFT and technical capabilities of the loading system imposed some limitations for
the load block. Preliminary estimations showed that a ratio of about 120 load lines per hour of flight,
(the planned 200 Simulated Flight Hours (SFH) were to have 24 000 load lines in one block), had to be
achieved in order to finish the test within the contracted time schedule. Since the sampling frequency of the
measured data was approximately 100 Hz, (400 Hz for buffeting loads), the amount of collected data was too
high. In order to achieve estimated ratio, the gathered data had to be filtered.
The level of filtration was chosen using iterative methods. Basing on the preliminary assumptions, resulting
from defined flight profiles in the Polish Armed Forces, representing 194 flights within the load block, flights
were filtered and put together to be close to the number of lines estimated. Since further reduction was planned,
and there was possibility to increase the loading frequency for buffeting loads when only empennage actuators
operate, this initial number of load lines was within acceptable range. As such, the preliminary load block
consisted of approximately 30 000 load lines.
The load block contained 194 flights each composed of: initial taxiing, flight with or without buffeting, landing
and final taxiing. Components were filtered and prepared separately. Hence it was necessary to verify load
sequences after load block assembly in order to eliminate load lines, for which load levels for all 20 actuators
either differed by less than 5%, or three sequential loads for all actuators were either ascending or descending.
Since the test specimen was not fully constrained within the test rig, it was necessary to assure that for all load
lines the resultant forces will be in static balance. Additionally, loads exerted on the structure were checked
along with the g factor to check if the resultant mass of the aircraft remained within a reasonable range (empty
airplane and fuel).
After performing all of the above steps, the final load block consisted of approximately 26 000 load lines: that
is acceptable due to the possibility of accelerating the test during buffeting loads.
The FSFT started in autumn 2011 and was contracted for 36000 Simulated Flight Hours (SFH). The load
spectrum used in the test represented 200 SFH, including landing and taxiing loads. The loading frequency,
achieved after initial tuning of the system, was around 0.5 Hz, which for the load spectrum consisting of 24 000
load lines allowed accomplishment of 1000 SFH every 15 working days, excluding shutdowns for NDT
inspections and holidays.
11 - 4 STO-TR-AVT-275
IMPLEMENTATION OF THE
AIRCRAFT STRUCTURAL INTEGRITY PROGRAM
PRINCIPLES FOR PZL-130 “ORLIK” TC-II TRAINER AIRCRAFT
Figure 11-3: FSFT Design and Actual Test Rig for the PLZ-13 Orlik Aircraft.
During the entire test, the structure was monitored using various NDT methods ranging from simple visual
inspection to eddy current or ultrasonic testing. According to tripartite consultations between EADS, AFIT
and VZLU, the non-destructive tests were divided into three levels.
Level 1 was based on simple NDI techniques and was carried out every 1000 ± 50 SFH. This inspection was
executed by VZLU in accordance with NDI instructions delivered by ITWL. The estimated time for
inspection was 2 man hours. No modifications to the test specimen or test rig were necessary for this level.
Level 2 inspection required using advanced NDI equipment thus additional access to the structure was
needed. This inspection was conducted every 5000 ± 1500 SFH for between 0 and 30000 SFH and the
time interval was shortened to 3000 SFH ± 1500 SFH for 30000 to 36000 SFH. VZLU also conducted
this level of inspection in accordance with NDI instructions delivered by the ITWL although first inspections
were conducted in the presence of ITWL qualified personnel. Estimated time for this inspection was
2 working days.
Level 3, detailed NDI inspection, included dismounting of wings and fuselage. Four major inspections were
planned throughout the test every 10 000 ± 2500 SFH. ITWL was responsible for this level of inspection in
accordance with NDI instructions delivered by ITWL. Estimated time for inspection was 5 working days.
A reporting routine, valid for all levels of inspection, was developed. It included detailed documentation of
found damage including photographic documentation and detailed definition of location, as well as methods
used to find and verify it. It allowed precise identifying of damage and its location, as well as enabled
tracking of it throughout the test to verify its severity and impact on structure integrity. Reports were
developed by VZLU and delivered to ITWL for interpretation and storage.
Figure 11-4 shows the test progress. Blue diamonds on the graph represent NDI inspections whereas the
thick red line defines the contracted test limit. Despite initial tuning problems, the test course was rather
stable, which allowed it to be finalized close to the estimated deadline. The test ended at 35056 SFH due to
critical lower wing spar damage, for which repair was both difficult to carry out and economically unjustified
due to its being so close to the FSFT limit.
Damage found during inspections was located in different structural parts of the aircraft. Most of the damage
was found in the wing area both in the right (137 findings, 35% of total findings) and left (124 findings, 32%
of total findings) wing parts. A total of 16 damages which account for 27% of the total findings were
detected in the fuselage area. Total of 21 damages were found in the horizontal stabilizer area and only
STO-TR-AVT-275 11 - 5
IMPLEMENTATION OF THE
AIRCRAFT STRUCTURAL INTEGRITY PROGRAM
PRINCIPLES FOR PZL-130 “ORLIK” TC-II TRAINER AIRCRAFT
4 damages in the landing gears. The damage also differed by the type, e.g., cracked rivets (32%), ERW
(Electro Resistance Welded joints failure, 27%) removed material or structure (22%), removed material or
rivet head (10%), and cracked screw (8%).
Figure 11-5 presents two graphs depicting the location where damage was found by percentage as well as a
chart describing the types of damage found.
Figure 11-5: Distribution of Damage Found Throughout the Structure and by Type of Damage.
Rivets were one of the more common failing items during the test. During inspections, some damages were
found in the early stages, when a crack was propagating within a rivet; some were found only when the rivet
had failed due to fast damage development. The left side of the Figure 11-6 shows a group of rivets located
on the fuselage left front side. Damaged rivets were found during 29 500 SFH Level 2 inspection. Additional
rivets were defined as being damaged on the right side of the fuselage during 33 000 SFH Level 1 inspection.
11 - 6 STO-TR-AVT-275
IMPLEMENTATION OF THE
AIRCRAFT STRUCTURAL INTEGRITY PROGRAM
PRINCIPLES FOR PZL-130 “ORLIK” TC-II TRAINER AIRCRAFT
The right side of Figure 11-6 shows the final damage number 818, that caused the test to terminate. The
critical damage occurred at 35056 SFH although several smaller damages (failed rivets and small cracks)
were detected in this region from 33 000 SFH.
Teardown Inspection (TI) of the PZL-130 TC-II after FSFT was carried out by disassembling the aircraft
structure down to separate parts [6] and performing versatile NDI in order to both examine the elements
which were inaccessible during the test, as well as to verify damages which were reported during the test [6].
Figure 11-7 shows the process of disassembling the wings. After additional analysis of the wing’s outer
surface using the MAUS V automated system, the skin sheeting was carefully removed by drilling out the
rivets thus allowing access to the inner wing structure.
STO-TR-AVT-275 11 - 7
IMPLEMENTATION OF THE
AIRCRAFT STRUCTURAL INTEGRITY PROGRAM
PRINCIPLES FOR PZL-130 “ORLIK” TC-II TRAINER AIRCRAFT
During TI, many new damages were found, the majority of which were located in the wings, since the wing
inner structure was inaccessible during the test. Like the results of NDI during FSFT, the damages were of
different types, ranging from failed rivets, through partial cracks, up to elements cracked completely through.
After, the TI selected cracked elements (both with through and closed cracks) were examined with
Philips XL30 Scanning Electron Microscope in order to identify fatigue markers, which were introduced in the
modified load spectrum [2], [7], [8]. A total of 23 samples were examined from various damages found in the
wing-fuselage connection area. During the study, a magnification range of 10-500000x was used. Nucleation,
fatigue fracture and final overload areas were clearly visible on the samples crack surfaces and fatigue fracture
areas of coupons tested before FSFT [2], [4]. Although the FSFT aircraft samples showed similarities, it was
not possible to define the fatigue marker bands caused by the reordering of the load spectrum.
A dedicated software was developed for analyzing and maintaining captured data. For each recorded flight,
the intensity factor, as well as the number of equivalent flight hours, are calculated. The intensity factor is
defined as the proportion of equivalent and actual flight hours for a single flight or set of flights. The
technical issue for the intensity factor is to compare the real load spectrum accumulated by an aircraft
structure during a flight with the defined spectrum used for the FSFT. This comparison is calculated based
on fatigue damage caused by both spectra. The fatigue damage is calculated according to S-N curve
methodology and uses the S-N curve defined in Ref. [9].
11.7 REFERENCES
[1] Leski, A., Reymer, P., and Kurdelski, M., “Development of Load Spectrum for Full-Scale Fatigue Test
of a Trainer Aircraft”. ICAF 2011 Structural Integrity: Influence of Efficiency and Green Imperatives:
Proceedings of the 26th Symposium of the International Committee on Aeronautical Fatigue and
Structural Integrity (ed. J. Komorowski), Montreal, Canada, 1 – 3 June 2011, pp. 573-584.
[2] Leski, A., Dragan, K., and Reymer, P., “Full-Scale Fatigue Test of PZL-130 Orlik Structure”.
Proceedings of the 27th Symposium of the International Committee on Aeronautical Fatigue and
Structural Integrity (ed. A. Brot), Jerusalem, Israel, 5 – June 2013, pp. 199-208.
[3] Department of Transportation Federal Aviation Administration, Report No. AFS-120-73-2 “Fatigue
Evaluation of Wing and Associated Structure on Small Airplanes”, Washington DC, May 1973.
[4] Anderson, A., and Parker, R.G.R., “Full-Scale Fatigue Test of the Pilatus PC9/A Trainer Aircraft”, 20th
ICAF Symposium: Structural Integrity for the Next Millennium, Bellevue, Washington USA, 1999.
[5] Molent, L., Barter, S.A., White, P., and Dixon, B., “Damage Tolerance Demonstration Testing for the
Australian F/A-18”. International Journal of Fatigue 31, pp. 1031-1038, 2009.
[6] Handbook for Best Practice in Teardown of Aircraft Structures, Technical Paper TR-AER-1-2005,
The Technical Co-operation Program, April 2005.
11 - 8 STO-TR-AVT-275
IMPLEMENTATION OF THE
AIRCRAFT STRUCTURAL INTEGRITY PROGRAM
PRINCIPLES FOR PZL-130 “ORLIK” TC-II TRAINER AIRCRAFT
[7] Barter, A.A., Molent, L., and Wanhill, R.J.H., “Marker Loads for Quantitative Fractography of Fatigue
Cracks in Aerospace Alloys”, 25th ICAF Symposium: Rotterdam, Netherlands, 2009.
[8] Leski, A., Kłysz, S., Lisiecki, J., Gmurczy, G., Reymer, P., Bochenek, D., and Zasada D., “Introduction of
Fatigue Markers in Full-Scale Fatigue Test of an Aircraft Structure”, Fatigue of Aircraft Structures, 2011.
[9] Federal Aviation Administration, Advisory Circular 23-13A, Fatigue, Fail-Safe, and Damage Tolerance
Evaluation of Metallic Structure for Normal, Utility, Acrobatic, and Commuter Category Airplanes,
US Department of Transportation, September 29, 2005.
STO-TR-AVT-275 11 - 9
IMPLEMENTATION OF THE
AIRCRAFT STRUCTURAL INTEGRITY PROGRAM
PRINCIPLES FOR PZL-130 “ORLIK” TC-II TRAINER AIRCRAFT
11 - 10 STO-TR-AVT-275
Chapter 12 – SU-22 FIGHTER-BOMBER AIRCRAFT SERVICE
LIFE EXTENSION PROGRAMME SUPPORTED BY
OPERATIONAL LOAD MONITORING
IMPLEMENTATION
Artur Kurnyta
Air Force Institute of Technology, Warsaw
POLAND
12.1 INTRODUCTION
In-service usage monitoring is an important airworthiness requirement for maintaining aircraft safety.
Fatigue load spectra change during operational use, especially for military aircraft, due to different tactical
requirements, new missions, advanced aircraft configurations (e.g., higher masses) and different
environments [1]. During recent decades, load and usage monitoring has grown from a simple g counter to a
highly complex and intelligent on-board system with the ability to monitor various locations in real time.
The modern approach to aircraft utilization is inseparably connected with the concept of material fatigue.
More than 100 years ago, at the beginning of aviation, hazards associated with material fatigue were poorly
recognized. The aviation industry in 1930 – 1940, when the first fully metal aircraft structures were
developed, simply didn’t deal with the material fatigue phenomenon. At that time, planes were designed
focusing mainly on the problem of static strength of the structure. In the following years, the development of
aircraft structures due to World War II, as well as the rapid evolution of public air transportation, caused
increased awareness of the phenomenon of material fatigue. From that point, combination of static strength
and fatigue life was taken into account during aircraft design [2], [3].
The development of aircraft design and operation techniques continues to this day, contributing to
an increase in safety and extending the life of aircraft. Unfortunately, the main “milestones” [4]
in consciousness regarding the phenomenon of structural fatigue, occurred after fatal aircraft crashes, e.g.,
de Havilland Comet in1954, F-111 in 1969, Boeing 707 in 1977 and Boeing 737 in 1988.
The rational use of an aircraft, i.e., use in accordance with its tactical and technical characteristics, is possible
only in the case of a functioning operating system, which ensures the optimization of aircraft usage. The
quality of the operating system is mainly determined by the maintenance and repair methods, as they contain
a set of regulations specifying the scope and frequency of service, control the level of reliability and
technical condition of the aircraft. Regarding the above, Individual Aircraft Tracking (IAT) by use of an
Operational Load Monitoring (OLM) system is essential to ensuring the implementation of the maintenance
and utilization strategy [5]. Moreover, OLM allows the determination of significant life extension potential
for the aging aircraft, which, after performing the necessary repairs and modifications, will enable their
further operation according to their technical condition [6], [7].
The case study of the Su-22 “Fitter” aircraft OLM system implementation is presented in this paper as a part
of the latest Service Life Extension Program (SLWP).
STO-TR-AVT-275 12 - 1
FIGHTER-BOMBER AIRCRAFT SERVICE
LIFE EXTENSION PROGRAMME SUPPORTED BY
OPERATIONAL LOAD MONITORING IMPLEMENTATION
which were prolonged in 1998 to 2000 FH / 3000 landings in 20 years for Su-22M4 and 2000 FH / 4000
landings in 20 years for Su-22UM3K. For both versions, the first general overhaul was scheduled after 800
FH in 12 years, and inter-renovation overhaul for 700 FH / 10 years. After the expiration of these resources,
the operational use of those aircrafts was carried out on the basis of the provisions of the operational bulletin,
which assumed a phased service life extension every two years or 200 FH as part of the maintenance.
The additional diagnostic and control works were carried out, determining the final service life of 3000 FH /
3000 landings in 30 years for Su-22M4 and 3000 FH / 4000 landings in 30 years for Su-22UM3K.
However, in 2014 the Polish MOD decided to continue the operation of the aging Su-22 “Fitter” for another
10 years, after successful confirmation of structural integrity by a Structural Service Life Extension Program
(SLWP). The SLWP included [8]:
• Ground and flight tests for a separate aircraft;
• Flight Data Recorder (FDR) archives analysis for the period of 5 years of operational use of the
whole fleet;
• Full-Scale Durability Test (FSDT) of a retired aircraft; and
• Operational Loads Monitoring (OLM) system implementation for all Su-22UM3K.
The double-seated trainer version was of special interest because it accumulated significantly more
landing-take-off cycles in their lifetime than the combat variant, and is threatened with experiencing more
severe loading spectra, both during flight and landing.
Ground and flight tests were performed to determine total loads on wings, and fuselage as well as landing
gear elements. During ground static tests, as well as 10 flights with a total duration of 8 hours and
25 landings (including 15 touch-and-go landings), the load spectra were obtained by measuring actual strains
in selected Critical Points (CP) from more than 40 strain gauges. The sensors were installed in locations that
enabled easy determination of the total wing bending moment acting in the wing attachment points as well as
the variable-sweep pivots. In addition, strain gauges mounted on the fuselage enable determination of the
total fuselage bending moments. A third group of sensors was installed on elements of the landing gear and
the landing gear attachment points. Other aircraft regions were omitted as the program focused on landing
gear and wing structure.
Thanks to the flight data records kept at the institute, fleet-averaged, as well as individual service profiles
were developed for the Su-22s. The M4 and UM3K variants were compared. The individual profiles are now
used in IAT activities that is a part of the SLWP. For the presented extension program, data from the
previous five years were used (2009 – 2013), as the service profile changed significantly through the years.
A period of five years was considered to be sufficient for obtaining a usage profile, and in addition,
a statistically significant number of flight hours had been recorded. The usage profiles were based on the
vertical acceleration (g-force) signal Nz, whereas the usage due to landings was calculated based on the
product of mass and vertical acceleration m × Nz. The rainflow method of cycle classification was used for
cycle calculation. The calculated usage cycle values have been distributed into bins of 0.25Nz width, taking
the maximum value in the range as the bin value. The usage profiles were calculated for an average 1000 FH
period. To determine the landing load profile, recorded landing histograms were created. Landing mass at
landing was determined based on the assumed empty aircraft weight (taking the aircraft variant into
account). This baseline was than corrected by adding the recorded fuel remnant. The installed armament
weight was not taken into account, due to lack of data.
Data from the ground and flight test as well as from FDR archives analysis were used to determine the load
distribution and correlation between flight parameters (mainly vertical overload) and strain prior to the
Full-Scale Durability Test (FSDT). The FSDT was conducted on a taken out of service aircraft which had
already experienced 3127 landings and 1583 Flight Hours between 1985 and 2003. The test was designed to
12 - 2 STO-TR-AVT-275
FIGHTER-BOMBER AIRCRAFT SERVICE
LIFE EXTENSION PROGRAMME SUPPORTED BY
OPERATIONAL LOAD MONITORING IMPLEMENTATION
prove the desired 3200 EFH (Equivalent Flight Hours), durability of 6500 landings with a safety factor of 4,
as well as the total durability of the aircraft structure. The detailed description of that activity is presented
in Chapter 25.
The main system element is an airborne multi-role recorder with integrated data acquisition, which manages
signal acquisition and archiving from a permanently bonded strain gauges sensor network and accelerometer.
Periodical analysis of aircraft structural loading for each aircraft is performed off-line on the Ground Station,
and suitable conclusions are drawn regarding further operational usage. A block diagram of the OLM system
is presented in Figure 12-1. The Data Acquisition Unit is an SSR-500 multi-role recorder manufactured by
Curtiss Wright Control Avionics & Electronics, with adequate analogue and digital user modules to acquire
data from 8 strain gauges, 11 parameters from Built-in Test for supervising the operation of the device,
13 GPS parameters and a 3-axis accelerometer. One user slot is filled with a dummy module presently, and is
ready for use in any future upgrade for a chosen type of sensor and measurement task, as the recorder has a
modular architecture and can be reconfigured as needed (Figure 12-2). The MT-1 is a custom-made and
tested additional unit for strain gauges bridge completion as well as for shunt calibration of strain gauge
measuring channels of SSR-500 DAU. A DC to DC converter is implemented to stabilize voltage from the
on-board DC power supply, to prevent under- or over-voltage during powering of other installed equipment.
The main data source for OLM is an integrated network of 8 strain gauges, mounted on predefined structural
critical points for monitoring the operational loading both during flight and landing to determine load
exceedances, e.g., harsh landings. A 3-axis, wide bandwidth Microelectromechanical System (MEMS)
accelerometer is used for load determination. Additionally, parameters like power-up and time-on counter,
internal voltage values on buses for e.g., sensor excitation and digital components, error counter, chassis and
internal temperature and chassis status, as well as GPS navigation data are acquired.
STO-TR-AVT-275 12 - 3
FIGHTER-BOMBER AIRCRAFT SERVICE
LIFE EXTENSION PROGRAMME SUPPORTED BY
OPERATIONAL LOAD MONITORING IMPLEMENTATION
A pair of strain gauges (T1 and T2 on the Left Hand (LH) side, T5 and T6 on the Right Hand (RH) side) are
bonded on the bottom and top flanges of the wing main spar, near the wing main attachment point.
Measurement from those channels are used during flight to determine loading in certain manoeuvres and
aircraft configurations. For exceedances during landings, one sensor is installed on the main landing gear strut
(LH side – T3, RH side T7) to measure longitudinal loads, and second sensor (LH side – T4, RH side T8) is
bonded on the gear-wing connection to measure lateral loads. The sensors are installed symmetrically on both
sides of the aircraft to be sensitive for asymmetrical loading as well as for redundancy (Figure 12-3).
The accelerometer is located near the centre of gravity point on the top of the fuselage. Main components
like DAQ SSR-500, MT-1 and DC-DC converter are located on top of the first technical hatch of the aircraft.
12 - 4 STO-TR-AVT-275
FIGHTER-BOMBER AIRCRAFT SERVICE
LIFE EXTENSION PROGRAMME SUPPORTED BY
OPERATIONAL LOAD MONITORING IMPLEMENTATION
Figure 12-4: Loads from Strain Gauges T1 and T4 During Typical Parts of the Flight.
STO-TR-AVT-275 12 - 5
FIGHTER-BOMBER AIRCRAFT SERVICE
LIFE EXTENSION PROGRAMME SUPPORTED BY
OPERATIONAL LOAD MONITORING IMPLEMENTATION
Figure 12-4 shows changes on T1 (LH side top flange of main spar) and T4 (LH side main landing
gear – wing connection point). It can be seen that particular flight phases are easily distinguishable even only
by strain gauges. Correlation and synchronization of the data from OLM and on-board data recorder give
enough information for individual tracking of a particular aircraft.
The main “added value” of having an OLM system implemented on-board the aircraft is the ability to detect
exceedances of operational use both during flight and landing. Those exceedances can have a meaningful
influence on fatigue parameters, like fatigue usage, operational intensity factor and equivalent-to-real flight
hours ratio. Moreover, parameters e.g., time of acceleration during start, detachment speed, speed during
closing the landing gear, touchdown speed, touchdown g-level, maximum peak-to-peak strain value during
touchdown. Based on OLM data, one should be able to detect e.g., harsh or asymmetric landing and quantify its
influence on aircraft usage [10].
12.5 SUMMARY
Operational load monitoring is a next step toward operational usage optimization of the aircraft. The need for
an accurate and reliable fatigue usage monitoring system is of increasing importance to ensure the safe and
economical utilization of aircraft, especially regarding aging aircraft, which are now expected to last much
longer than first envisaged. At the moment, implemented OLM allows for counting and determination of
operational load exceeds both in the air and during landings for the twin seated version of the Su-22 aircraft.
Overloads can have significant influence on airframe strength due to high mechanical strain occurrence.
Data analysis based on the Flight Data Recorder enables specifying usage profile values as operating
intensity factors and equivalent flight hours for each monitored aircraft.
Activities in progress within the SLWP program are now focused on parallel utilization of the direct strain
gauge data for the aircrafts equipped with the OLM system for individual tracking and determination of
usage profile values as operating intensity factors and equivalent flight hours for each monitored aircraft.
12.6 REFERENCES
[1] Dilger, R., Hickethier, H., and Greenhalgh, M.D., “Eurofighter a Safe Life Aircraft in the Age of
Damage Tolerance”, International Journal of Fatigue 31, pp.1017-1023, 2009.
[2] National Research Council, “Aging of US Air Force Aircraft”, National Research Council, Washington
DC, ADA330900, January 1997.
[3] Schutz, W., “Fatigue Life Prediction for Aircraft Structures and Materials”, AGARD-LS-62,
ICAF-Doc-693. Advisory Group for Aerospace Research and Development, France, 1973.
[4] Wanhill, R.J.H., “Milestone Case Histories in Aircraft Structural Integrity”, NLR TP 2002 521.
National Aerospace Laboratory NLR, 2002.
[5] Molent, L, and Aktepe B.P., “Review of Fatigue Monitoring of Agile Military Aircraft”, Fatigue
and Fracture of Engineering Materials & Structures 23(9): pp. 767-785, 2000.
[6] Grandt, A., “Fundamentals of Structural Integrity: Damage Tolerance and Nondestructive Evaluation”,
Wiley-Interscience, 2004.
[7] Newcamp, J., Verhagen, W., and Curran, R., “Aging Military Aircraft Landscape: A Case for End-of-Life
Fleet Optimization”, 8th European Workshop on Structural Health Monitoring, Bilbao, 2016.
12 - 6 STO-TR-AVT-275
FIGHTER-BOMBER AIRCRAFT SERVICE
LIFE EXTENSION PROGRAMME SUPPORTED BY
OPERATIONAL LOAD MONITORING IMPLEMENTATION
[8] Kurdelski, M., Reymer, P., Stefaniuk, M., and Kurnyta, A., 2017 “Service Life Extension Program
Based on Operational Load Monitoring System and Durability Test of the Ageing Fighter-Bomber
Jet”, Proceedings of 29th ICAF Symposium, Nagoja, Japan.
[9] Kurnyta, A., Zieliński, W., Reymer, P., and Dziendzikowski, M., “Operational Load Monitoring
System Implementation for Su-22UM3K Aging Aircraft”, proceedings of the 11th International
Workshop on Structural Health Monitoring, Stanford University, Stanford, CA, September 12 – 14,
2017;
[10] United States Department of Defense. MIL-STD-1530C: Aircraft Structural Integrity Program
(ASIP), 2005.
STO-TR-AVT-275 12 - 7
FIGHTER-BOMBER AIRCRAFT SERVICE
LIFE EXTENSION PROGRAMME SUPPORTED BY
OPERATIONAL LOAD MONITORING IMPLEMENTATION
12 - 8 STO-TR-AVT-275
Chapter 13 – AIRCRAFT STRUCTURAL INTEGRITY
MANAGEMENT ACTIVITIES IN FINLAND
Ilpo Paukkeri
Finnish Defence Forces Logistic Command
FINLAND
13.1 INTRODUCTION
This chapter provides a brief review on aircraft structural integrity management activities in Finland from the
Finnish Defence Forces (FDF) viewpoint.
The national approach to aircraft structural integrity management was initiated by the Finnish Air Force
(FINAF) during the late 1990s and early 2000s, following the identification of the severity of the operational
usage compared to the design spectra. Aircraft Structural Integrity Program (ASIP) was implemented for the
FINAF F/A-18C/D Hornet fleet according to MIL-STD-1530 [1]. Lately, the ASIP approach has also
introduced for the FINAF BAE Hawk.
The FDF ASIP has evolved as a primary means to manage structural integrity in order to maintain safety and
availability of aging systems and ensure economic service life of the fleets. From the fleet management
perspective, effective ASIP processes provide data for the following objectives:
• Allow required number of aircraft to remain in service until the planned out-of-service date;
• Keep maximum number of aircraft available, and enable prediction of necessary repairs or
modifications;
• Instruct deployment of the fleet in terms of severity of usage; and
• Instruct decommissioning of the fleet beginning from the most stressed individuals.
The FIMAA has delegated authority to FDF Logistics Command (FDFLOGCOM) to act as a Type
Certification Holder (TCH) and Continuing Airworthiness Management Organization (CAMO) for
the Finnish military aircraft, with the following tasks:
• Develop airworthiness criteria, and processes and means to show compliance with the airworthiness
criteria;
• Develop and sustain a maintenance program and a structural integrity plan for each type; and
• Maintain a system for data collection, assessment, analysis and reporting related to failures and
defects that pose risks for maintaining airworthiness of the aircraft.
STO-TR-AVT-275 13 - 1
AIRCRAFT STRUCTURAL INTEGRITY
MANAGEMENT ACTIVITIES IN FINLAND
The FDF’s TCH does not have an organic design capability and is typically supported by OEMs and the
domestic partners.
As Finland has its unique operating environment, tasks and usage spectrum, the FDF must have its own
capability to:
• Assess effects of the usage spectrum on the service life and structural integrity of individual
airframe and systems;
• Implement changes and repairs needed to the structure due to differences in actual usage vs. design
usage spectrum; and
• Adjust maintenance requirements and intervals to maintain optimal availability with minimized
costs.
Furthermore, the national capability should enable re-certification of the fleet (component) when extension
of the service life is deemed necessary.
Regarding new fleet acquisition programs, a requirement for comprehensive structural integrity management
capability places a request for the OEM to provide a sufficient data package along with a need for
contractual agreements to enable access to engineering data when the need arises during the sustainment.
A requirement for an independent design capability necessitates the most complete set of engineering data,
including material and process specifications, numerical models (CFD, FEM) and fatigue test reports with
test spectrum and load data package.
Task IV (Certification and Force Management Development) involves the activities to develop the basis for
Task V. The list of activities (not exhaustive) includes:
• Identify the certification basis and requirements: design standards, certification criteria, etc. and
related guidance material, as well as verification compliance documents;
• Develop systems and processes to perform Operational Loads Measuring (OLM) and Individual
Aircraft Tracking (IAT);
• Establish a structural maintenance plan for each of the critical structural components/locations;
• Define the national policy for preventive structural inspections, modifications and repairs;
• Estimate fatigue lives for the critical structural components/locations based on the data from
monitoring/tracking systems;
• Decide on the primary measures of structural life consumption for fleet management and
maintenance scheduling;
• Evaluate and maintain the structural inspection program in order to ensure sufficient safety and
availability levels with minimized costs; and
• Decide on additional measures especially concerning structural management of aging aircraft,
e.g., fleet leader inspections and structural teardown programs as the two most common concepts,
which provide also means to evaluate effectiveness of ASIP.
13 - 2 STO-TR-AVT-275
AIRCRAFT STRUCTURAL INTEGRITY
MANAGEMENT ACTIVITIES IN FINLAND
Task V (Force Management Execution) embodies the actual fleet monitoring and structural integrity
management tasks, based on collecting usage data by the HOLM and IAT systems and assessing identified
usage changes and damage findings from structural inspections (including information from other operators).
The usage data collected by the HOLM and IAT systems is further utilized to determine a baseline operating
spectrum for fatigue analysis, to adjust timing and intervals for structural inspections and modifications,
as well as to detect other types of events that affect the structural service life.
All the relevant information and actions are documented in the ASIP book, compiled in accordance with the
MIL-STD-1530 standard and updated on regular basis. The ASIP book consists of two volumes: Volume I
describes the principles of integrity management, identifies the design and development data, and determines
life targets and limits, whereas Volume II includes the actual, detailed management plan for the fleet.
The effective ASIP execution is based on continuous collaboration and communication within the FDF and
strategic partners. The ASIP status is communicated regularly to engineers and managers responsible for the
fleet management. The status reports include decided service life goals and applied fatigue indices, fleet
fatigue overview and trends, projections for end of life, and operational guidance for the individual aircraft
based on the fatigue indices.
An overall effectiveness of ASIP is evaluated regularly and an action plan is prepared based on the
evaluation in order to improve the identified focus areas. The evaluation is applied for each critical structural
detail and then individual evaluations are compiled to get the overall ASIP status. The evaluation addresses
at least the following essential questions:
• What is the actual level of implementation of the ASIP activities?
• Do the activities contribute to achievement of the service life goal?
• What is the operational risk level now / projected for the end of life?
• Does the certification represent the current perception?
STO-TR-AVT-275 13 - 3
AIRCRAFT STRUCTURAL INTEGRITY
MANAGEMENT ACTIVITIES IN FINLAND
As the F/A-18 ASIP already addresses most of the topics, the objective is to document the best practices
experienced with F/A-18 and define policies applicable to other FDF fleets as well.
The current lifing practices for F/A-18C/D and BAE Hawk Mk. 51/51/66 form a generally good basis to be
followed. Those lifing practices are based on activities within Finnish industry, as OEMs were not required
to establish ASIP programs per MIL-STD-1530 or per any other similar standard.
For other FDF aircraft, OEM practices, especially in those aircraft certified in the Federal Aviation
Regulations (FAR)/Certification Specifications (CS) 25 category, may be similar but unknown. There are no
ASIP programs.
Aircraft certified in the FAR 23 category generally do not follow lifing policy principles at all. The same
applies for rotorcraft airframes in the FAR 27 category.
The R&D activity has been carried out as subsequent multi-annual programs conducted by the FDF,
involving a collaborative network of industrial partner companies, universities and the VTT Technical
Research Centre of Finland (VTT). The main areas of R&D activity during the 2010s with some highlights
are briefly described in Sections 13.6 and 13.7.
Operational loads and structural stresses for the aircraft has been determined by use of nationally developed
capabilities for Computational Fluid Dynamics (CFD) and Finite Element Analysis (FEA) supported by
flight simulation capability. Examples of recent activities include:
• Enhancing the CFD model of the Hawk and applying the model to predict aerodynamic load
distributions for the tail plane in different flight conditions, to be utilized in a domestic fatigue test
for a component;
• Modelling the flow field around NH90 fuselage and other developments in helicopter flow
simulation model and computations;
• Further CFD method development especially related to time-accurate simulations and fluid-structure
interactions;
• Determining load spectra for F/A-18 major components and interfaces, in order to perform more
accurate fatigue life assessments for critical structures in FINAF usage; and
• Developing crack growth analysis capability in order to implement damage tolerance approach; for
that purpose, studying limit load cases for FINAF F/A-18 to determine critical crack sizes,
necessitating evaluation of aerodynamic limit load distributions for the whole aircraft in critical
design load cases.
OLM and IAT systems provide the fundamental means for collecting actual operational usage data for
structural integrity management. National R&D activities have been carried out on this area during the last
two decades, featured by development of OLM and IAT systems for the F/A-18 aircraft.
13 - 4 STO-TR-AVT-275
AIRCRAFT STRUCTURAL INTEGRITY
MANAGEMENT ACTIVITIES IN FINLAND
The systems currently applied for the structural integrity management of the FINAF F/A-18 Hornet are
briefly described below:
• SAFE (Structural Appraisal of Fatigue Effects) is the OEM’s fatigue tracking system that measures
strains in the structural locations considered the most critical and calculates Fatigue Life
Expenditure (FLE) for those locations. The SAFE system provides IAT capability and primary
measures for monitoring fatigue life consumption of the FINAF F/A-18 aircraft [2].
• The HOLM (Hornet Operational Loads Measurement) programme has been running since 2006 with
the goal of quantifying the effects of operational usage on the structure of F/A-18 Hornet aircraft. The
programme was initiated by the FINAF while the SAFE system was considered incomplete for the
FINAF life management purposes in terms of sampling rates and coverage of the structure. The
programme is based on two F/A-18C aircraft equipped with relatively high number of strain gauges
and a few accelerometers, measuring loads on various structural locations during actual operation in
squadron usage and providing an ability to predict fatigue damage accumulation on various structural
locations. The data captured from the two HOLM aircrafts is available as input data for different types
of service life assessments and for verification of numerical models (CFD, FEM) [2].
• The parameter-based fatigue life analysis is an IAT method developed by Patria to meet the demand
to evaluate structural life consumption of the whole fleet without any modifications to the aircraft
systems. The analysis makes use of the flight parameter data, stored by standard aircraft systems,
and artificial neural networks that have been trained by the measured strain gauge data from the
HOLM flights. The neural networks are utilized to generate virtual strain gauge data, which allows
calculating predicted fatigue lives for each individual aircraft [2].
Non-Destructive Inspection (NDI) technologies have been investigated nationally and new methods
implemented to monitor the integrity of structural components of the FINAF aircraft. As the damage
tolerance approach has been introduced, a need has emerged to quantify the reliability and performance of
each NDI method in order to validate the NDI-based decisions, such as determination of inspection intervals.
For this objective, the statistical method called Probability of Detection (POD) has been applied in a couple
of studies, increasing understanding of aspects related to performance of NDI and realizing the importance of
POD assessment [2].
National capability for designing and performing structural repairs and modifications has involved
investigating fatigue characteristics on different materials and components by experimental testing,
developing processes of carrying out repairs on composite materials and structures [2].
The scope of the FISIF is to assess the structural integrity of the F/A-18 aircraft and constituent parts. The
scope of the CREDP is to investigate, develop, and expand repair techniques for structural components of the
F/A-18 by conducting research and engineering tasks.
Participation in the FISIF and CREDP has been extremely valuable for national structural integrity
management. Direct data exchange with other operators has provided the FDF with information on
inspection findings, new inspection methods, fatigue tracking data, structural repairs and modifications. The
received data have promoted flight safety and enabled the FDF to plan and implement structural
refurbishment programmes needed for securing adequate structural service life.
STO-TR-AVT-275 13 - 5
AIRCRAFT STRUCTURAL INTEGRITY
MANAGEMENT ACTIVITIES IN FINLAND
13.7 CONCLUSION
The FDF ASIP has evolved as a primary means to manage structural integrity in order to maintain safety and
availability of aging systems and ensure economic service life of the fleets.
Although initiated during the sustainment phase, the FDF ASIP is intended to cover all the tasks defined in
the MIL-STD-1530. Design and development (Tasks I-III) information is to be identified as far as possible
and incorporated into the program, to be assessed when the need arises during the sustainment.
The overall effectiveness of ASIP is evaluated regularly, and an action plan prepared in order to improve the
identified focus areas. As regards lessons learned, the FDF experience suggests early initiation of ASIP,
particularly establishing national certification and lifing policies and validating the structural life of aircraft
as exactly as possible. An important action is to perform POD assessment to validate NDI-based decisions.
Furthermore, an effective ASIP execution necessitates continuing alertness within the organization,
in addition to capable analysis tools and available repair capabilities.
Continuous and collaborative R&D programs focused on aircraft structural integrity management have been
considered crucial to develop in-country engineering and design capability on implementing ASIP.
With fairly limited R&D and engineering resources within the FDF and its domestic partners, focusing on
the most effective and efficient activity is important.
Regarding new acquisition programs, a requirement for comprehensive structural integrity management
capability places a request for the OEM to provide a sufficient data package and a need for contractual
agreements to enable access to engineering data when the need arises during the sustainment.
International co-operation is highly acknowledged by the FDF, particularly the information exchange and
collaborative actions within the F/A-18 user community that has greatly improved understanding and
knowledge within the FDF and domestic partners. Using the common standard base has helped to share
information and adopt ASIP practices from other users.
Increasing structural issues with the aging aircraft pose risks and challenges to meet requirements for
operational availability and economic service life. Major concerns include crack findings on fracture critical
structures for which there is no “full life” modification available and findings on structures that allow very
small critical crack sizes. In addition, exceedance of certified safe life for some components may necessitate
either significant re-certification activities or introduction of new structural inspections to further add to the
increasing workload. These are examples of structural issues that justify the necessity to maintain a
responsive and effective ASIP through the remaining service life.
13.8 REFERENCES
[1] MIL-STD-1530, Department of Defence Standard Practice: Aircraft Structural Integrity Program
(ASIP), 1 November 2005. Accessible via internet: http://everyspec.com/MIL-STD/MIL-STD-1500-
1599/download.php?spec=MIL-STD-1530.023416.pdf.
[2] Viitanen, T., and Siljander, A. (Eds.), A Review of Aeronautical Fatigue Investigations in Finland
March 2015 ‒ March 2017, Presented at the 35th Conference of the International Committee on
Aeronautical Fatigue and Structural Integrity (ICAF), Nagoya, Japan, 5 ‒ 6 June 2017. Available at:
https://www.vtt.fi/inf/julkaisut/muut/2017/ICAF_Doc2433_Finland_Review_2017.pdf.
13 - 6 STO-TR-AVT-275
Part 3: AIRCRAFT PROPULSION SYSTEMS
STO-TR-AVT-275 Part 3 - i
Part 3 - ii STO-TR-AVT-275
Chapter 14 – AIRCRAFT PROPULSION
SYSTEMS (UNITED STATES)
Paul Bascom
United States Air Force, Life Cycle Management Center
UNITED STATES
14.1 INTRODUCTION
This section will address the United States Air Force’s (USAF) approach to propulsion continued
airworthiness as expressed in our fielded engine life management approach. The methods and practices used
to assure engines are being supported with the proper technical activities to provide the service life for the
required safety levels and maintenance costs will be discussed. Engine life management concepts have been
developed and have evolved within the propulsion community to provide a roadmap of activities that must
be considered for the support of the engines in the field. The Propulsion System Integrity Program (PSIP), as
described in MIL-STD-3024, offers a methodical systems engineering approach to the full life cycle
management of propulsion systems and the key elements of sustainment and airworthiness will be described.
This focused effort also highlighted additional overall factors affecting engine durability, specifically the
roles of design, manufacturing, maintenance, and operations. Design factors identified were failure modes
attributed to low cycle fatigue, wear, crack propagation, creep/stress rupture, dynamics, thermal
environment, materials, and usage. Manufacturing contributors were the roles of quality, inspection,
misassembly, and methods/techniques. Maintenance factors were items such as spares, tech data,
STO-TR-AVT-275 14 - 1
AIRCRAFT PROPULSION SYSTEMS (UNITED STATES)
inspections, repairs, and training. The area of operations included usage, environment, foreign object
damage, tactics, and parts tracking.
ENSIP broke down the propulsion life cycle into five sequentially ordered tasks:
1) Task 1: Design Information;
2) Task 2: Design Analysis, Material Characterization & Development Tests;
3) Task 3: Component and Core Engine Tests;
4) Task 4: Ground and Flight Engine Tests; and
5) Task 5: Engine Life Management.
ENSIP is an organized and disciplined approach to the structural design, analysis, development, production,
and life management of gas turbine engines with the goal of ensuring engine structural safety, increased
service readiness, and reduced life cycle costs. The application of ENSIP introduced two concepts:
1) The introduction of fracture mechanics based design and life management system (damage
tolerance); and
2) The use of Accelerated Mission Testing (AMT) to assess durability. It should be noted that not all
failure modes are equally exercised during an AMT.
For engines entering development since 1978, damage tolerance has been incorporated into the acquisition
program as prescribed by ENSIP, with existing engines receiving a Durability And Damage Tolerance
Assessment (DADTA). Damage tolerance concepts are used in ENSIP to define critical part inspection
requirements:
• Where to inspect;
• When to inspect; and
• How to inspect.
This enhances safety, reliability, and quality assurance through defining intelligent inspection requirements.
The first four tasks are a building block approach to produce a knowledge base supporting the sustainment or
life management task where closed loop force management techniques are utilized to address actual usage.
14 - 2 STO-TR-AVT-275
AIRCRAFT PROPULSION SYSTEMS (UNITED STATES)
scheduled to place an early emphasis on the definition of proper design criteria, to perform detailed structural
analyses before hardware is fabricated, and to conduct component verification testing at realistic operational
loads. The overall objective of the mechanical integrity program is to enhance engine safety and reliability
by improving component durability, damage tolerance, and quality. An important by-product of this program
is the reduced cost of ownership that is realized when the production engine system is deployed for use.
In recognition of and response to this overarching approach, a new document emerged to address all aspects
of the propulsion system called PSIP. It was also decided that the significance of this effort warranted
elevation back to a military standard which permitted it to be contractually referenced by the DOD. This
approach successfully unifies all the propulsion elements of ENSIP, MECSIP, and AVIP into one governing
document describing the system integrity at the whole engine level. Additionally, recognition of the
increasing interdependence and integration of the propulsion system within the overall weapons system is
addressed across all phases of development with emphasis on a rigorously controlled Interface Control
Document (ICD). The ICD is a formalized agreement between propulsion system provider and the weapon
system contractor establishing physical and functional definitions and responsibilities. This can range from
routine definition of a common coordinates system (x, y, and z planes) to incoming inlet distortion patterns at
the aircraft interface plane.
The five pillar construct originally envisioned for turbomachinery now encompasses the controls and
subsystems components, bringing rigor and consistency across the whole spectrum of life cycle
management. The urgency of embracing this approach became apparent with the progressing complexity of
current and future propulsion systems and the ever increasing integration and dependency among previously
segregated hardware entities. This unified approach harnesses the following elements across the entire
propulsion system:
• Design based on mission usage profiles and expected service life exposure;
• Durability capabilities that account for failure modes such as low cycle fatigue, high cycle fatigue,
and creep;
• Adherence to the component classification process describing safety-critical, mission-critical,
durability-critical, and durability non-critical categories;
• For safety-critical and mission-critical designations, appropriate damage tolerance compliance
actions are defined;
• Materials and process characterization;
STO-TR-AVT-275 14 - 3
AIRCRAFT PROPULSION SYSTEMS (UNITED STATES)
• Definition of the internal and external environments, including aerothermal and vibratory
influences; and
• Demonstration of robustness is accomplished by engine testing and AMT.
The activities listed, along with many other elements of the engine development process, form an essential
building block of knowledge when transitioning to full production status and sustainment.
The goal of an LTF program is to accelerate operationally relevant exposure on a series of engine assets by
prioritizing their insertion into flying aircraft. Careful coordination with the using command is essential to
ensure their support of this objective while minimizing the impact on mission readiness. The intent is to use
this select population to encounter any failure modes and buy reaction time for the larger fleet with spares,
repair techniques, or redesigns if necessary.
Embedded within an LTF program are scheduled ACIs, where the engines are completely disassembled and
methodically inspected. Non-destructive inspection techniques are used to assess the wear and durability
characteristics of the hardware and then documented for the program office engineers. This progressive
approach that accompanies the flying accumulation provides the engineering community with quantifiable
data on wear out modes, hot section distress progression, and other leading health indicators on the
individual components.
Engines are designed to an initial set of expected aircraft missions in the development phase of acquisition.
As the aircraft are fielded and get operational experience it’s essential to update the mission usage attributes
and issue component life updates reflecting actual usage. Personnel from the government program office and
engine OEM interview pilots, acquire and review significant engine parameters from either on-board or
offboard sources, to create new mission profiles. Additionally, these new profiles are used to generate new
mission definitions for conducting any subsequent AMTs, ensuring any exposure testing has the greatest
relevance to the fleet experience.
When the most suitable remedy for a discovered safety-related shortfall is a redesign, that component
becomes a candidate for CIP to fund the necessary engineering activities for redesign and test. CIP is an
14 - 4 STO-TR-AVT-275
AIRCRAFT PROPULSION SYSTEMS (UNITED STATES)
ongoing program for engines within the Air Force portfolio to execute design solutions and designate engine
assets for full-scale AMTs as required to evaluate the effectiveness of the changes.
Safety-related events are governed by a process described by propulsion’s best practice titled “Aircraft Gas
Turbine Engine Flight Safety Risk Management Process”. This document describes the methods and practices
used to actively track flight safety issues and apply sound engineering principles to eliminate or significantly
reduce the technical risk of encountering engine failures. Figure 14-1 provides an overview of the process
while the rates and consequences are governed by Engine Related Loss Of Aircraft (ERLOA) and Non-
Recoverable In-Flight Shut-Down (NRIFSD) metrics. Program office engineers track the progress real time
with thrice yearly senior engineering leadership reviews for all the engines within the Air Force portfolio.
14.6 SUMMARY
Continuing airworthiness is best ensured when rigorous systems are in place to establish baseline capabilities
of gas turbine engines and then manage and monitor their actual usage. Unplanned and unexpected exposure
or failures are swiftly identified, assessed, and management plans laid in place as remedies are developed.
Robust monitoring of the fleet and agile responses with sound systems engineering principles have proven to
be effective at keeping the US Air Force fleet always postured to be safe, effective, and ready.
14.7 REFERENCES
[1] Cassidy, F., Vukelich, S., and Sammons, J., AIAA 89-2464 “Engine Life Maturation Process”,
AIAA/ASME/SAE/ASEE 25th Joint Propulsion Conference, 1989.
[2] Holmes, R., RTO Technical Report 28 “Recommended Practices for Monitoring Gas Turbine Engine
Life Consumption”, Chapter 2, 2000.
[3] Department of Defense Handbook, Engine Structural Integrity Program (ENSIP), MIL-HDBK-1783B
w/change 2, 2004.
STO-TR-AVT-275 14 - 5
AIRCRAFT PROPULSION SYSTEMS (UNITED STATES)
14 - 6 STO-TR-AVT-275
Chapter 15 – US ARMY PROPULSION –
CONTINUING AIRWORTHINESS
Gary Kellogg
US Army Combat Capabilities Development Command
UNITED STATES
This paper will share the US Army’s approach in maintaining airworthiness of these older systems from
a propulsion perspective. The methods used in establishing initial airworthiness of drive systems will
be discussed as well as the ongoing efforts to sustain those systems through additional testing and
usage spectrum updates. The two primary turboshaft engines used in these platforms, T55-GA-714A and
T700-GE-701D engines, are military qualified engines with ongoing programs to maintain their
airworthiness as well. Additional information regarding the T700 life management will be provided to share
how the US Army has evolved its philosophy toward using current fleet usage instead of prescribed mission
profiles. The benefit of other Army aviation enterprise programs that monitor new part manufacturing as
well as component condition throughout its life cycle are described.
The “Basis for Airworthiness” includes analysis and testing during product development. Many propulsion
components are deemed ‘on condition’, whereas the component may remain in service an indefinite period of
time as long as it meets inspection requirements. Some components have a prescribed retirement life
(aka, life limit) due to a limitation in capability, such as Low Cycle Fatigue (LCF). These life limits are
established to maintain safe operation of the component based on data and analyses available at that time. As
the fleet matures into production and fielding, the basis for those life determinations must be reviewed
periodically. Several potential changes that impact the life limit may have occurred, including material
processing, changes in design, and analysis tool revisions, as well as field experience and mission requirements.
STO-TR-AVT-275 15 - 1
US ARMY PROPULSION – CONTINUING AIRWORTHINESS
15 - 2 STO-TR-AVT-275
US ARMY PROPULSION – CONTINUING AIRWORTHINESS
Rotorcraft transmissions are designed to the drive system Maximum Continuous Power (MCP) rating
specified for that particular platform. The components are sized for operating at MCP 100% of the time.
Stress analysis of gear shafts verifies that static strength requirements are met and peak fatigue stresses are
below the material endurance limit. Gearbox bench testing is conducted to demonstrate that gear teeth have
infinite life in tooth bending. Bearings are required to have a minimum rolling contact fatigue life of
4500 hours at 75% of the drive system MCP rating or cubic mean power. Gearbox endurance and overstress
testing is conducted to ensure there are no wear or other failure modes that would limit the service life of
gearbox assembly (as well as to verify the gearbox capability to operate for some limited power excursions
above MCP). This results in most gearbox components having an unlimited fatigue life which allows for
maintenance based ‘on condition’ and hence components are only removed for cause. US Army rotorcraft
gearboxes also include debris detection, temperature, and/or vibration sensors to identify excessive wear or
other impending failure modes before airworthiness is affected. Many drive system components also have
required periodic maintenance inspections, servicing, and/or scheduled overhaul intervals to ensure safe
operation. Barring any appreciable increase to the drive system MCP rating or other system performance
requirements, this overall approach results in continued airworthiness throughout the service life of aircraft.
There are other drive system components that are not designed by torque alone. For example, the critical
loading and design of main and tail rotor drive shafts and some gearbox housings is also impacted by aircraft
flight loads, and their fatigue life may not be unlimited. These components are subject to fatigue testing as
part of the original design and qualification. The fatigue life analysis includes conservative assumptions
regarding material strength, aircraft usage, and flight loads which are used to compute a retirement time with
six nines of reliability. To maintain this level of reliability, it is important to periodically reassess aircraft
usage. Assumptions used to develop original usage spectra can become outdated due to changing mission
requirements, training, aircraft upgrades, and piloting techniques. US Army Fatigue Life Management, as
defined in the Aeronautical Design Standard Handbook ADS-79E-HDBK, Condition Based Maintenance
System [2] for US Army Aircraft includes detailed guidance for updating aircraft usage. This involves using
regime recognition tools, health and usage monitoring data, and pilot surveys. Partial updates to the usage
spectrum are also conducted that focus on certain damaging regimes in an effort to evaluate published
retirement time for specific components. Any potential airworthiness concerns for these components in an
aging fleet are mitigated by the original design and qualification requirements, field maintenance, periodic
updates to usage spectra and component retirement lives.
The T55 engine operating limits are defined in the aircraft operator’s manual and the maintenance
procedures required to ensure proper engine operation are defined in the maintenance manuals. Daily
pre-flight checklists describe actions such as ensuring that the inlet is free of debris, the engine filter status
indicators are clear, there are no oil/fuel leaks, all attachments are secure, etc. In addition, engine overhauls
are performed at 2,800 hour intervals. During an overhaul, the engine is removed from the aircraft and sent
STO-TR-AVT-275 15 - 3
US ARMY PROPULSION – CONTINUING AIRWORTHINESS
to an approved overhaul facility (Corpus Christi Army Depot (CCAD) or Honeywell’s Greer, South Carolina
facility). At the overhaul facility, a complete teardown and rebuild of the engine is performed. Any
component at or near its life limit is replaced, the compressor and turbine blades are replaced and any other
issues addressed.
General issues affecting the fleet are communicated to the field by safety or maintenance information
messages. These messages notify the operators of issues identified by the technical community and describe
the actions required to mitigate these issues.
The T55 engine component cyclic/fatigue life consumption is tracked by cycle counting algorithms that are
part of the Electronic Control Unit (ECU). These algorithms track the life for seven critical rotating
components based on changes of engine rotor speed. Component life cycles are computed for each flight and
are summed with previous counts to establish the total life cycle count. The total life cycle count indicates
the fatigue life consumed. A safe life fatigue limit for each critical component has been established by testing
and analysis. The component is removed from service when the total life cycle count, as determined by the
ECU algorithms, is equal to the established fatigue life limit.
The T700-GE-700 life limits were originally developed by GE Aircraft Engines, using a prescribed mission
mix from the engine model specification. There were a total of 10 missions anticipated for the UH-60A
platform; including troop assault, resupplying units in combat, troop extraction, sling load, and training. These
missions provided the operational input as to how the engine was expected to operate. However, over a period
of time, many changes to that mission mix are noted. Aircraft tend to gain weight over time, which increases
the power required across most of the mission service operational spectrum. Changes in tactics, training and
procedures often impact the duration of missions as well as the ‘mix’ of the missions. Therefore, it is very
important to provide the engine manufacturer with current information regarding how the engine is operated.
During the 1990s, there were various engine data recording devices available to install on aircraft. The US
Navy uses variants of the T700 engine that shares many parts with the T700 engine used by the US Army
and was successful in installing recording devices and obtaining fleet representative data which was screened
and eventually provided to GE. These mission updates were then used by GE as part of the life management
process with eventual updated life limits made available to the US Navy. The US Army was unsuccessful in
installing the same devices due to electromagnetic interference concerns. As a result, the 10 mission mix was
slightly adjusted based on the increased power available of the T700-GE-701C engine; that updated
10 mission mix was the basis for the T700-GE-701C life determinations.
Upon completion of the T700-GE-701C life analysis, the recommended life limits were substantially lower
than expected. One means to assess the accuracy of the newly calculated lives was to perform a validated life
analysis. The CT7 turboprop engine had a similar hot section configuration and significant cumulative
operating experience and a substantial database of field inspections including more than 200 eddy current
inspections. Using the newly updated analysis toolset as well as operational experience from the GE CT7
turboprop engine, GE performed a validated life analysis. The analysis indicated that the predicted lives were
inconsistent with field experience. While there were no known cracking events in the field, the analysis
15 - 4 STO-TR-AVT-275
US ARMY PROPULSION – CONTINUING AIRWORTHINESS
predicted that several events should have occurred. This analysis allowed the US Army to risk manage the fleet,
instead of extensive field removals and their associated impact on readiness and cost. It also emphasized the
substantial benefit of inspecting high time parts and maintaining a database on inspection results.
During the development of the T700-GE-701D, emphasis was placed on understanding the engine’s LCF
capability from a “double hump” cycle perspective, an example of which is shown in Table 15-1. Many
helicopter applications have missions that require a major cycle (i.e., engine start through maximum rated
power) as well as minor cycles. The severity as well as the sheer volume of minor cycles is important to
properly assess as the damage accrued during these minor cycles can be a significant contributor to the
calculated life limits of a component.
Due to this emphasis on the double hump cycle, GE Aircraft Engines performed a finite element model
based LCF Analysis of the T700-GE-701D engine with the double hump cycle and reported a minimum life
for each analyzed component in terms of cycles. The US Army manages components in terms of hours so an
additional assessment was required to convert the cyclic lives to hours to support the logistical capabilities of
the current system. By this time, the US Army had accomplished installing recording devices on aircraft and
was then able to obtain actual fleet data that could be utilized in some fashion during a lifing analysis. The
assessment was performed by the US Army and utilized simplified lifing algorithms developed by US Army
personnel to calculate fatigue damage based on the available fleet data.
Fundamental engineering principles (e.g., stress as a function of speed squared) are used to define a
mathematical relationship that estimates the stress for a given speed range. Information provided by
GE Aircraft Engines in the T700-GE-701D LCF Analysis was utilized with this relationship to define
constants that are unique for each feature and component assessed. With these constants now defined, the
relationship can be used to determine the stress for any other speed range given. The life for a given speed
range can then be calculated and fatigue damage calculated as the fraction of life consumed for the given
speed range. The damage from all identified speed ranges for a flight data record is summed to calculate the
STO-TR-AVT-275 15 - 5
US ARMY PROPULSION – CONTINUING AIRWORTHINESS
mission life in hours. For a set of flight data, the same process is followed to determine individual mission
lives and damage. The individual mission damage and individual mission duration is summed for all flights
of the data set and used to calculate the component life limit in hours.
These algorithms described above are packaged in a single analysis tool which is able to reduce the size of
actual flight data files to rotor speed peak/valley points by eliminating duplicate and intermediate points. The
tool then performs rainflow cycle counting and the damage summation for all peak/valley points of the
recorded mission. These algorithms enable the US Army to no longer rely on mix of anticipated missions, as
was the case with the 10 mission mix, but on the actual missions being flown by the warfighter today.
The initial life limits published for the T700-GE-701D engine utilized the aforementioned US Army
“Cycles-to-Hours” lifing algorithms to assess roughly 21,000 flight records of T700-GE-701C and -701D
usage. Due to the infancy of the -701D fleet at the time, a decision was made to leverage the -701C usage
data and apply correction factors to the -701C calculated life developed from statistical analysis of the
competing fleets. The correction factors applied to differences in average mission duration and power turbine
speed for various airframes.
Several years later, this assessment was performed again but with a much larger data set of flight data
records and additional airframes that had been upgraded with recording equipment since the initial analysis.
Roughly 90,000 flight records compiled across three airframes were assessed for the revised life limits and
due to this analysis, increased volume correction factors no longer needed to be applied. Additionally, newly
developed analysis tools by US Army personnel enabled the data set to be quickly filtered, segregated, and
statistically analyzed to give the US Army a better understanding into actual engine usage and the
similarities and differences inherent between airframes. Finally, the typical removal cause and previously
established removal limits were considered to ensure that the new life limits were supportable from a
logistics perspective. These analysis tools developed during the lifing assessments of the T700-GE-701D
engine have now been adapted and used for other engine families in the US Army inventory.
Lastly, the T700 Product Office has taken a proactive approach to remove the T700-GE-700 and -701
models from inventory as well as reducing the number of 701C models. The T700-GE-701D will be the
single T700 configuration. The initiative was pursued from a logistical and cost perspective but
provides many airworthiness and safety benefits as well. The T700-GE-701D has benefitted from
almost 40 years of improvement and has an In-Flight Shut-down (IFSD) rate of less than 0.1 failures per
100,000 flight hours. Therefore, the main focus from an engine airworthiness perspective is to maintain
vigilance over changing operating environment, new product quality and adherence to field and depot
maintenance practices and procedures.
15 - 6 STO-TR-AVT-275
US ARMY PROPULSION – CONTINUING AIRWORTHINESS
The PLT program established limited examinations for airworthiness assurance of both Critical Safety Items
(CSI) and non-CSI items that require testing and which are used on Army Aircraft. The inspections are
focused on safety, performance, form, fit and function as well as forward and reverse traceability.
PLT encompasses both non-destructive and destructive testing of CSIs (as well as key component features)
to verify that the parts have the proper characteristics and meet key drawing requirements subsequent to the
manufacturing process.
Future systems are envisioned to take advantage of significant growth in processing capability and analytics
such that life-limited components will be retained in service based on their actual usage. Each serial numbered
part will have its unique retirement life based on how it was operated in the field. Substantial physics-based
analysis and correlation testing will be required to ensure adequate safety margin exists for this approach.
Changes in the US Army’s record keeping and reporting systems will be necessary to accommodate this drastic
departure from current practice that has the same component life for every part. Initiatives to create longer
periods of maintenance free operations will require improved diagnostics and prognostics.
15.3.2 RIMFIRE
As mentioned previously, one of the key knowledge points to properly managing the propulsion systems is
understanding the condition of the hardware once it is returned to the depot. The Reliability IMprovement
through Failure Identification and REporting (RIMFIRE) program was established to document the condition
of T700 engines upon initial disassembly of the engine at depot. Its approach was to accurately record
component condition and any relevant failure mode data, develop a web-based database with a user friendly
interface, conduct engineering assessments and provide reliability reports to identify key opportunities for
component improvement. The intent was to make informed decisions on prioritizing component improvements
to maximize safety of flight, engine time on wing and cost reductions. The program was extended to other
Army engines, drivetrain components, rotor blades and other flight critical components.
The RIMFIRE program has evolved into the predominant source of identifying and quantifying the
component failure mode data for Propulsion components. During the process of inspecting components and
noting their condition, unexpected hardware condition and configurations were documented. The original
program was subsequently expanded to include Special Inspections, whereas specific inspections can be
conducted on a particular problem area. For example:
• The use of incorrect common fasteners installed by field units at a critical location was discovered
during inspections. A special inspection was put in place to quantify the number of occurrences,
which was the basis for a determination on the need for further field action;
• Manufacturing discrepancies have been detected and quantified for a number of engine and gearbox
assemblies. Once again, special inspections were established and statistical analyses performed to
estimate the suspect population size;
• Special inspections were instituted to measure the effect of a shaft surface condition due to a quality
escape. The data was used to determine the expected impact on fielded hardware that justified no
further field action;
STO-TR-AVT-275 15 - 7
US ARMY PROPULSION – CONTINUING AIRWORTHINESS
• Data gathered during engine disassembly was used to correlate the location of component weep
holes to determine is a major source of oil leakage could be attributed to the alignment of those
weep holes; and
• Bearing damage, high cycle fatigue failures, and improper balance grinding are some of the over
34 types of failures that RIMFIRE has detected that could have caused In-Flight Shutdowns (IFSD).
While many subject matter experts focus on maintaining appropriate Low Cycle Fatigue (LCF) life limits on
critical rotating parts as a primary means to maintain continued airworthiness of our aging fleets, the
RIMFIRE program has illustrated that a comprehensive inspection program at the depot is essential in
maintaining airworthiness while minimizing field actions that affect readiness and cost.
One item under consideration is the development of a database that contains the listing of components that
have finite life limits (retirement lives), as well as their final disposition upon retirement. A database that
includes the final inspection results is critical to ensure that our predicted life limits are appropriate as well as
establishing a database for any future lifing validation tasks. This initiative should ensure components are
being inspected at the time when any crack formation is most likely (near their retirement life) to be detected.
15.3.4 Summary
The Commanding General, US Army Aviation and Missile Command, is the US Army’s airworthiness
authority. He has delegated that authority to the Combat Capability Development Command (CCDC),
Aviation and Missile Center’s (AvMC) Aviation Engineering Directorate (AED). The AED develops
airworthiness criteria, standards and methods of compliance so that the necessary analyses and test data can
be generated and approved to provide the basis for airworthiness. Continued airworthiness of the systems
requires cognizance of the material, operational, maintenance and environmental changes inherent to aircraft
systems as well as an ability to determine the airworthiness impact of those changes. Such an endeavour
requires extensive design knowledge of the product, proactive data gathering of the product’s current
condition and an awareness of probable future changes.
15.4 REFERENCES
[1] US Army Regulation 70-62, Airworthiness of Aircraft Systems, Figure 2-3 Continued Airworthiness
Process Flowchart, 11 May 2016.
[2] US Army Aeronautical Design Standard Handbook ADS-79-HDBK, Condition Based Maintenance
System for US Army Aircraft, Appendix A, Fatigue Life Management, 8 Feb 2016.
[3] Army Military Airworthiness Certification Criteria (AMACC), Appendix D, Engines, Aircraft,
Turboshaft, paragraph 3.3.8.3 Low Cycle Fatigue (LCF) Life, 12 March 2019.
15 - 8 STO-TR-AVT-275
Chapter 16 – ENGINE STRUCTURAL
INTEGRITY PROGRAM (ESIP)
P. Trembath
Directorate of Technical Airworthiness and Engineering Support
CANADA
16.1 INTRODUCTION
A number of high profile catastrophic or near-catastrophic instances of uncontained failure of critical engine
components provide graphic illustrations of the need to maintain the structural integrity of engines. Various
studies suggest incremental improvement with successive generations of engine design. Several Royal
Canadian Air Force (RCAF) aircraft types continue to operate with engines that do not benefit from these
improvements.
A causal factor for loss of structural integrity of engines is inaccurate usage tracking due to changing mission
profiles. If increased usage severity is not captured by programmatic review and update, inadvertent
exceedance of life limits may occur; conversely, if usage severity has fallen, then critical parts may be retired
prematurely.
From an aging aircraft perspective, there is increased potential difference between the actual engine usage
and the assumption-based design (or baseline) engine usage as the aircraft ages. Therefore, it is essential to
ensure the engine critical component life limits and associated assumptions are re-assessed throughout the
life of the engine to ensure they are reflective of actual usage [1].
STO-TR-AVT-275 16 - 1
ENGINE STRUCTURAL INTEGRITY PROGRAM (ESIP)
any change to the Statement of Operating Intent (SOI), change to the Aircraft Structural Integrity Program
(ASIP) baseline usage spectrum, or similar change identified by the WSM should trigger a review of usage.
16.2.5 Determine Life Consumption Based on Actual Usage (of Critical Components, as
a Minimum) Accordingly
If it is found that there has been a change to the baseline usage spectrum such that either the life limits or the
calculation of life consumption changes, the results must be incorporated into the approved maintenance
program. If a reduction to life limits is identified, fleet records are to be reviewed to determine if any
in-service components exceed life limits as a result.
In order to facilitate programmatic execution and technical review and assessment of the program,
supporting documentation is also required.
16 - 2 STO-TR-AVT-275
ENGINE STRUCTURAL INTEGRITY PROGRAM (ESIP)
Involvement in Component Improvement Programs (CIP): several RCAF fleets are members of a CIP,
which provides access to and collaboration with other users as well as a forum to identify and address
common problems with OEM and other stakeholder involvement.
IVHM (Integrated Vehicle Health Management): under the ESIP, IVHM systems are not required; any
practicable means of capturing adequate data to ascertain parts life consumption is permitted, and automated
data recording systems are not required.
Several fleets make use of Engine Health Monitoring (EHM) and/or Health and Usage Monitoring System
(HUMS) data for limited trend monitoring and operational or maintenance planning purposes. The use of such
systems is not programmatically driven, and is fleet-specific based to a great extent on OEM offerings; legacy
fleets may not have robust means for automated data recording for the purposes of EHM or usage monitoring.
Several fleets also make use of Oil Analysis (OA) programs: RCAF OA providers standardize against Joint
Oil Analysis Program samples. The use of OA is not programmatically driven (i.e., as a component of an
across-the-board IVHM requirement), and is fleet-specific based on legacy or OEM maintenance program
requirements.
16.4 OUTCOMES
It is not yet possible to ascertain the effectiveness of the ESIP, as compliance with the regulation is not
required until 30 Jun 2019. Review of fleet in-service programs has in several instances not identified a clear
chain of analysis linking RCAF-specific usage rates to parts life consumption. It is understood this does not
necessarily imply a safety concern, but unless such a chain of analysis exists parts life consumption must be
based upon conservative assumptions in order to ensure adequate safety, which may be substantially
sub-optimal economically for the RCAF. Further, whereas reviews of usage rates for engine critical
components have been conducted ad hoc, the ESIP mandates a periodic review. Reassessment of the cyclic
exchange rate (low cycle fatigue major cycles consumed per flying hour) of one RCAF fleet, performed
shortly before the release of ESIP requirements, found that evolutionary changes in role and mission mix had
resulted in significant increases in life consumption rate, on the order of 20 – 30 % per flying hour, since the
last time such an analysis was carried out. This case is representative of the rationale for the ESIP.
STO-TR-AVT-275 16 - 3
ENGINE STRUCTURAL INTEGRITY PROGRAM (ESIP)
16.5 REFERENCES
[1] Technical Airworthiness Authority, TAA Advisory 2016-02e, Engine Structural Integrity Monitoring
Requirements, 5 July 2016.
16 - 4 STO-TR-AVT-275
Chapter 17 – AIRCRAFT ENGINE STRUCTURAL INTEGRITY
MANAGEMENT ACTIVITIES IN FINLAND
Ilpo Paukkeri
Finnish Defence Forces Logistic Command
FINLAND
17.1 INTRODUCTION
This chapter provides a brief review on the approach and activities related to aircraft engine structural
integrity management in Finland, from the Finnish Defence Forces (FDF) viewpoint. In addition to this
chapter, the report includes dedicated chapters for aircraft structures (Chapter 13) and mechanical systems
(Chapter 21).
The FDF have not implemented a systematic, comprehensive approach for aircraft engine structural integrity
management. However, many related activities have been conducted within the FDF and its industrial partners,
especially for the engines of the F/A-18C/D Hornet fighters and the BAE Hawk jet trainers operated by the
Finnish Air Force (FINAF), as well as the NH90 rotorcraft operated by the Finnish Army. In addition to
national activities, there are life management programs established by the OEMs and the FDF’s involvement in
international user groups and programs, which has provided important fora to share information with other
users as well as allowed to participate in collaborative research and development efforts.
Along with engine Maintenance, Repair and Overhaul (MRO), a capability for demanding repairs and
component manufacturing have been developed in country with Patria as a strategic industrial partner.
During in-service operation with the FINAF Hornet and Hawk fleets, a significant number of repairs and
modifications have been developed nationally – with very limited design and certification data. The lacking
data and design capability have been partly patched up by national R&D activities.
The impact of national R&D and repair development efforts have been acknowledged, especially in the
remarkable cost savings achieved. Experimental research has enabled the FDF to extend the recommended
component life or the engine maintenance/overhaul intervals. Nationally developed repair schemes have
been implemented successfully for several life-limited and expensive parts. In some cases, repair
development has been extended to redesign and manufacturing of new parts.
The ENSIP standard is organized into five basic tasks of which Tasks I – IV relate to engine design,
development and testing/certification, whereas Task V is dedicated to engine life management, which is
apparently the most essential from the FDF perspective.
The initial review of ENSIP tasks suggest the following major elements when considering implementation of
the standard:
• Identification of the Tasks I – IV information and engineering data required for the
development/execution of the engine life management (Task V) capability;
• Determination of the component’s life limits;
STO-TR-AVT-275 17 - 1
AIRCRAFT ENGINE STRUCTURAL
INTEGRITY MANAGEMENT ACTIVITIES IN FINLAND
• Monitoring and collecting data on the actual usage (loads / environment and engine performance
parameters);
• Individual engine / component tracking program and life assessment; and
• Engine structural maintenance plan.
The life limits determined for the components, as well as the required maintenance actions (i.e., Engine
structural maintenance plan) are to be reviewed and updated based on the actual monitoring of usage and
continuous inspections/assessments.
An example of recent enhancement of the national monitoring and data collection capability is the engine
health monitoring system implemented in the FINAF Hawk Mk66 fleet. The system measures operational
engine parameters and stores them on-board, to be downloaded after the flight to the embedded life
assessment system which counts the cycles accumulated for each life-limited part.
The development of data analytics is a key technology area in order to exploit the growing amount of
collected data, providing Continuing Airworthiness Management Organization (CAMO) with tools for life
cycle planning, fleet management and maintenance planning/scheduling.
Spare parts’ availability and cost issues have been the usual drivers behind the repair development projects
or the establishment of capabilities in component manufacturing. Besides the application of manufacturing
techniques, some cases have necessitated a design change for a component due to insufficient reliability or
performance.
In addition to developing solutions for urgent problems occurring within the fleet sustainment, Patria is
continually exploring new methods of manufacturing and repairing aircraft parts. Special techniques, such as
the application of thermal coatings and electron beam welding, are widely used today. Patria’s work on 3D
printing (additive manufacturing) technology achieved an important milestone in January 2018 with
successful completion of the first 3D printed part installed and flying in the F/A-18 Hornet. The part was
designed in accordance with the MDOA approval granted to Patria and was manufactured from the Inconel
625 superalloy. MDOA approval refers to Military Design Organization Approval in accordance with
European Military Aviation Requirements (EMARs) and granted by the Finnish Military Aviation Authority
(FMAA).
17 - 2 STO-TR-AVT-275
AIRCRAFT ENGINE STRUCTURAL
INTEGRITY MANAGEMENT ACTIVITIES IN FINLAND
Special emphasis has been on the experimental methods (e.g., creep tests and associated metallographic
investigations) focused on certain turbine blades and the possibility for associated life extension, with the
following examples:
• Analysis of the FINAF F/A-18 Hornet low pressure turbine blades [2];
• Analysis of the FINAF F/A-18 Hornet high pressure turbine blades [2];
• Microstructural degradation of a single-crystal gas turbine blade [3];
• Guideline – turbine blade failures [3]; and
• The effect of volcanic ash on gas turbine blades and vanes [3].
Experimental research support from VTT Technical Research Centre of Finland (VTT) has enabled the FDF
to increase the operational life (flight hours) of the high pressure turbine blades of the FINAF F/A-18
engines. The 10% increase in life yielded approximately 3 million USD in savings to the taxpayers, with a
nearly hundred-fold Return On Investment (ROI).
An analysis of the FINAF F/A-18 Hornet low pressure turbine blades ordered by the FDF from the VTT is
briefly described in the ICAF 2017 review [2] from which some direct quotes are made in the following
paragraphs:
• Three FINAF F/A-18C Hornet Low Pressure Turbine (LPT) blades were received for investigation
that aimed to determine in-service effects such as coating condition, microstructural degradation,
cracking and attachment wear, and also to determine the 3D and cross-section geometry of the
blades. Two of the blades were removed after 2496 EFH (Engine Flight Hour) and 3890 EOT
(Engine Operation Time) after reaching the end of their nominal life. The third blade had seen 1363
EFH and 2130 EOT in an engine that had sucked in snow when the aircraft incidentally veered off
the runway [2].
• The blades were photographed, 3D laser scanned (dimensional geometry), inspected
(stereo microscopy) and then sectioned at selected locations for subsequent metallographic
investigations [2].
• The crystallographic orientation did not show significant deviations from the expected status. No
obvious life-threatening damage in terms of coating condition, thermal blade material degradation or
cracking was observed in any of the inspected LPT blades. The fully served blade included one 0.1
mm deep crack in a location and orientation that was not threatening the blade integrity [2].
• In the blade with the snowbank incident (rapid in-service cooling), there was some limited coating
delamination at the base plate and on the blade root area. The observed modest microstructural
degradation of the blades was not considered to indicate life-threatening damage either [2].
• The dovetail area of the blades investigated showed some wear marks at the contact surfaces and
deformation of the base material under the contact surfaces and some lamellar cracking, but there was
no indication of fatigue cracking developing from these lamellar cracks or from the outer surface of the
waist of the blades. Therefore, a modest extension of operating hours could be justified [2].
STO-TR-AVT-275 17 - 3
AIRCRAFT ENGINE STRUCTURAL
INTEGRITY MANAGEMENT ACTIVITIES IN FINLAND
An analysis of the FINAF F/A-18 Hornet high pressure turbine blades ordered by the FDF from the VTT is
briefly described in the ICAF 2017 review [2] from which some direct quotes are made in the following
paragraphs:
• Three of the FINAF F/A-18C Hornet High Pressure Turbine (HPT) blades flown the extra 10%
were removed from service and subsequently subjected to metallographic investigation to compare
the blade condition after 3160 EOT against the condition of previously studied blades which had
been used up to a nominal life of 2800 EOT. The main effort on damage characterization was
concentrated on overheating on the concave side of the blade, microstructural degradation and depth
of cracks at the leading edge [2].
• The concave side’s overheating did penetrate the blade’s wall in one of the three inspected blades
and thus exceeded the acceptance criteria. Nevertheless, the overheating damage does not seem to
threaten the blade life and can be easily monitored by visual inspection. Only in one blade some
cooling holes at the blade tip were blocked by debris, but this was not believed to be life-threatening
either [2].
• The maximum crack depth at the leading edge after 3160 EOT did not exceed that of previously
assessed blades after nominal 2800 EOT. Also, the gamma prime (γ’) depleted zone inside the
cracks was continuous up to and within the crack tip which suggests sufficiently slow crack growth
to justify the previous decision to extend the blade life from 2800 EOT by 10% [2].
17.5 CONCLUSION
The FDF has identified a need for investigating the implementation of a systematic and comprehensive
approach for aircraft engine structural integrity management, with the intention of avoiding duplication of
effort by examining generally known standards, such as the MIL-STD-1783 (ENSIP). Another standard to
be considered is the MIL-STD-3024 Propulsion System Integrity Program (PSIP) as it provides a
comprehensive system engineering approach addressing structural/mechanical/electrical systems integrity
(see Chapter 14). The establishment of integrity management processes along with associated critical
capabilities is considered crucial to supporting the FDF’s Type Certificate Holder (TCH) and Continuing
Airworthiness Management Organization (CAMO) functions.
The initial review of ENSIP tasks suggests the following major elements when considering implementation
of the standard:
• Identification of the Tasks I – IV information and engineering data required for
development/execution of the engine life management (Task V) capability;
• Determination of the component life limits;
• Monitoring and collecting data on the actual usage (loads/environment and engine performance
parameters);
• Individual engine/component tracking program, life assessment; and
• Engine structural maintenance plan.
Regarding best practices, the experience gained from implementing Aircraft Structural Integrity Program
(ASIP) may provide some applicable practices when considering implementing an Engine Structural
Integrity Program (ENSIP). For example:
• Identify all the information related to the engine design and development phases, and incorporate
them in the ENSIP to be assessed when the need arises during the sustainment;
17 - 4 STO-TR-AVT-275
AIRCRAFT ENGINE STRUCTURAL
INTEGRITY MANAGEMENT ACTIVITIES IN FINLAND
• Develop and maintain an up-to-date ENSIP plan/database for the fleet, documenting all the relevant
information and actions; and
• Evaluate regularly the overall effectiveness of the ENSIP and prepare an action plan to improve the
identified focus areas.
The national activity related to aircraft engine structural integrity is highlighted by the repair development
and R&D achievements enabling significant life cycle cost savings while assisting in securing the required
levels of aircraft availability and flight safety. From that experience, the following can be recommended to
support the aircraft engine structural integrity management program:
• Establish repair development programs and manufacturing capabilities in country, with the potential
to provide feasible and cost-effective solutions for engine component obsolescence/availability
issues;
• Invest in R&D on jet engines to create knowledge and capability within the domestic research
community to be utilized as basis for engine component life assessment/extensions and other
structural integrity management activities; and
• Promote active collaboration with the OEMs and participation in international user groups, which
provides access to other users’ findings, sharing best practices and technology advancements along
with actual collaborative research efforts.
17.6 REFERENCES
[1] MIL-HDBK-1783B, Department of Defence Handbook: Engine Structural Integrity Program (ENSIP),
15 February 2002. Available at: http://everyspec.com/MIL-HDBK/MIL-HDBK-1500-1799/MIL
_HDBK_1783B_1924/.
[2] Viitanen, T., and Siljander, A. (Eds.), A Review of Aeronautical Fatigue Investigations in Finland
March 2015 ‒ March 2017, Presented at the 35th Conference of the International Committee on
Aeronautical Fatigue and Structural Integrity (ICAF), Nagoya, Japan, 5 ‒ 6 June 2017. Available at:
https://www.vtt.fi/inf/julkaisut/muut/2017/ICAF_Doc2433_Finland_Review_2017.pdf.
STO-TR-AVT-275 17 - 5
AIRCRAFT ENGINE STRUCTURAL
INTEGRITY MANAGEMENT ACTIVITIES IN FINLAND
17 - 6 STO-TR-AVT-275
Chapter 18 – AGING OF ENGINE STRUCTURES
Wieslaw Beres
National Research Council (NRC)
CANADA
A thorough understanding of the failure mechanisms affecting gas turbine components is essential if the
failure modes, the safe life, and the life usage of each component are to be accurately determined and safely
monitored [2]. Typical failure mechanisms summarized in Table 18-1 [3] are:
• Low Cycle Fatigue (LCF);
• High Cycle Fatigue (HCF);
• Thermomechanical Fatigue (TCF);
• Creep (C);
• Overstress;
• Corrosion (Cor);
• Hot Corrosion (HC);
• Erosion (ER); and
• Fretting and Wear (WR).
Table 18-1: Failure Modes and Life-Limiting Properties for Turbine Engine Components [3].
STO-TR-AVT-275 18 - 1
AGING OF ENGINE STRUCTURES
Legend:
ER Erosion LCF Low Cycle Fatigue
C Creep HCF High Cycle Fatigue
COR Corrosion HC Hot Corrosion
TMF Thermomechanical Fatigue WR Wear
The ability of a component to resist the effects of these failure mechanisms is a function of the material
properties, the component design and manufacture processes and also the component operating environment.
These features of the component are fixed by the design and application of the engine and cannot be affected
by the method in which the engine is operated or maintained. Conversely, there are external factors which
also have an influence on the rate of component life consumption but which can be reduced during engine
manufacture, operation and maintenance [2]. The external factors affecting life usage rate are:
• Manufacturing methods and material defects;
• Build and maintenance errors;
• Foreign Object Damage (FOD); and
• Limit exceedances during operation. [2]
These four engine damage sources are effectively controlled by engine manufacturers, operators and
maintainers therefore they could be described as avoidable. However in reality, in operational practice, this is
difficult to achieve.
During long-term service in an extreme environment under high stresses and at high temperatures, turbine
engine components such as discs, blades or vanes suffer from cumulative damage, which gradually degrades
their mechanical properties and could lead to a component failure or even loss of engine structural integrity.
18 - 2 STO-TR-AVT-275
AGING OF ENGINE STRUCTURES
metallurgical aging. Depending on the component, alloy composition and service condition, the
plastic damage may be due to creep, Low Cycle Fatigue (LCF) or creep-fatigue interactions.
As internal damage builds up, the resistance of components to deformation under static (creep) or
cyclic (LCF) loading is reduced. Extensive loss of resistance to plastic deformation can lead to
cavitation and internal cracking under creep conditions or ductility exhaustion and crack initiation
under LCF or TMF conditions. Both lead to failure.
A summary of all basic failure modes for gas turbine engine components is shown in Table 18-1. The
primary failure modes that are assessed for life usage purposes are LCF and creep [3].
LCF damage is dominant in bores, bolt hole and fillet areas of compressor and turbine discs. It is caused by
stress cycling associated with engine start-up and shut-down and with RPM excursions during service. While
the driving force for LCF damage accumulation in compressor discs is primarily mechanical, in hot section
areas of the engine it may also be associated with thermal cycling. In addition, TMF damage in vanes and
high-performance blades is often a life-limiting factor in modern gas turbine engines.
The component service temperatures are usually high enough to cause aging of the material, which results in
loss of tensile and creep strength and therefore its load bearing capacity.
Creep damage is usually predominant in the mid-airfoil section of a turbine blade and in some cases in rims
of turbine discs, where the stresses and temperatures may be sufficiently high to cause time-dependent
plastic deformation.
Engine designers expend a great deal of effort to establish life limits for components whose deterioration
may threaten operational safety. The lifing procedures employed to establish life limits vary with
components and the mode of damage. In modern turbines, discs and spacers are normally designed to
withstand LCF and burst due to overspeed as well as creep for the hot section side of the engine. Turbine
blades and vanes are designed to withstand creep as well as thermomechanical fatigue and High Cycle
Fatigue (HCF).
The life limits for these components are initially established based on service estimates of damage
accumulation rates. They are then revised as field experience is accumulated. Typical lifing procedures can
be grouped as follows:
• For rotating components the most common lifing procedure employed to establish safe life limits
follows a time or a cycle to crack initiation criterion, which is factored down to represent
a minimum property component;
• For turbine blades and vanes, life limits may or may not be prescribed by engine designers.
In particular, turbine blades and vanes in aerospace gas turbine engines are seldom lifed because of
the difficulties associated with predicting the service behaviour of complex alloy systems under
operational conditions that vary widely with user practice; and
• In the case of coated components, these difficulties are compounded due to coating-substrate
interactions, which are usually not well understood. A “Life-on-Condition” approach is sometimes
employed, where, for instance, creep growth or untwist of airfoils is measured and distortion limits
are used as retirement criteria.
As indicated, rotating parts such as compressor and turbine discs are usually life-limited due to LCF damage
accumulation. The most common and traditional lifing method employed for these parts follows a
“cycles-to-crack initiation” criterion, where a minimum life capability is defined statistically for simulated
service conditions through extensive coupon testing and component test verification in spin test rigs.
The statistical minimum is usually based on the probability that no more than 1 in 1000 components will
STO-TR-AVT-275 18 - 3
AGING OF ENGINE STRUCTURES
have developed a detectable crack (typically chosen as a crack of 0.8 mm in length). It should be mentioned
that the value “1 in 1000” is chosen arbitrarily. Various engine manufacturers use different values, e.g., 1 in
980, or 1 in 750 components [3].
This approach has been criticised as being overly conservative and costly because in this lifing approach
critical rotating components, such as discs, are usually discarded with significant amounts of useful residual
life. There are two major concerns when a cycles-to-crack-initiation rejection criterion is used to life rotating
parts. The first one is that, by implication, 99.9% of the components will be retired before any detectable
crack has formed and it is possible that these parts may be capable of significantly longer service before they
develop a detectable crack (actually 0.8 mm in length). The second concern is that the discarded components
may be capable of tolerating crack of sizes much greater than the 0.8 mm typically used. This limit actually
reflects the sensitivity and reliability of current Non-Destructive Evaluation (NDE) methods rather than the
mechanical tolerance of a particular part for cracks. For example, crack detectability by dye penetrant
methods depends on crack length, whereas crack depth is the critical factor for lifing purposes [3].
Because many modern components have demonstrated high levels of resistance to crack growth, and modern
materials have different characteristics, alternative lifing procedures based on damage tolerance and the
application of fracture mechanics principles, are now in use. The philosophy behind these alternative lifing
procedures assumes that the component may be capable of continued safe operation during crack growth,
providing that the cracks grow sufficiently slowly during service to allow their growth to be reliably detected
and perhaps even monitored through regularly scheduled inspections.
In their most elementary forms, these alternative damage tolerance based lifing procedures, which are also
known as Life-On-Condition, Retirement-for-Cause or simply fracture mechanics lifing, assume that the
fracture critical locations of a component contain cracks of a size that lie just below the detection limit of the
Non-Destructive Inspection (NDI) technique used to inspect the component. The crack is then assumed to
grow during service in a manner that can be predicted by linear elastic fracture mechanics, or any other
acceptable method, until a predetermined dysfunction limit is reached beyond which the risk of failure due to
rapid crack growth becomes excessive [3].
The accuracy of the process for determining safe component life and the monitoring of its consumption have
a major impact upon the safety of the engine or on its life cycle cost. Engine design-features and operational
environments that undermine safety the most or are very costly, both in direct cost or in maintenance efforts,
must be identified. Only then can managers assess the potential benefit of proposed modifications to the
design or changes to operating practices. Even minor changes to operating practice, which may have little or
no operational impact, can significantly lower life usage rates and bring the benefits of higher availability
and reduced life cycle cost [2], [3].
For Damage Tolerance (DT) based procedures to be successful, supporting methodologies must be
developed. These supporting technologies include:
• Non-destructive inspection;
• Mechanical testing of test coupons and components;
• Structural analysis;
• Mission profile analysis; and
• Condition monitoring of components.
Particularly, extensive material testing involving room and elevated temperature crack growth rate data
must be performed for the application of deterministic and probabilistic fracture mechanics based life
prediction concepts.
18 - 4 STO-TR-AVT-275
AGING OF ENGINE STRUCTURES
Not all components are candidates for life usage management to damage tolerance limits. Two examples are:
• Components made out of material with a short crack propagation life; and
• Components in a situation when the engine parts cannot be inspected.
Few, if any, mature engines that are currently in service were designed with damage tolerance
methodologies, though some may be viable candidates for damage tolerance management.
Basic damage modes, such as Low Cycle Fatigue (LCF), Thermomechanical Fatigue (TMF), creep or
High Cycle Fatigue (HCF) have been identified for all critical components of gas turbines.
Characterization and understanding of these damage modes acting separately are generally
well-developed. However, components in service are subjected to a combination of these damage
mechanisms induced by the effects of mechanical forces, temperature and environment. Therefore the
major challenges in life prediction and life extension of gas turbine engines are to understand the
complexity of multiple damage modes and their interactions. Often, a linear cumulative damage model is
used to count damage induced by fatigue, creep and creep-fatigue interaction, even though there is a lack
of a physical rationale underlying this rule [9].
The fatigue damage accumulation process in structural components can be broken into crack nucleation,
small crack growth and large crack growth stages. A life obtained from material coupon testing supported by
spin rig verification is usually regarded as the “safe life” of a component, which implies that it is defect-free
until the end of its prescribed life. This safe life approach was the sole lifing philosophy until the advent of
fracture mechanics, which provided the theory and methodology to deal with crack growth behaviour in
materials. The lifing methodology that takes into account the period of crack growth is referred to as the
“damage tolerance” philosophy [10], [11], [9], [12].
In general implementation of lifing procedures for critical components, there are key questions that need to
be addressed to practically life a given component. Some of these questions include: “What is the definition
of crack nucleation?” “What are initial cracks?” “What is the crack growth rate?” and also “how can cracks
in structural components be detected before they become critical?” Addressing these questions poses great
challenges to engineering implementation of the lifing procedures. In addition, at the theoretical level other
issues, such as models that can bridge the crack nucleation and crack growth stages, must be addressed.
These include physics-based modelling of fatigue and creep crack nucleation, with their applications to life
STO-TR-AVT-275 18 - 5
AGING OF ENGINE STRUCTURES
prediction of aerospace components [9], [12], [13], [14]. In addition, efforts were made to develop
diagnostics, prognostics and health management systems for gas turbine engines, as described in Refs. [10],
[11], [15], [16].
Damage tolerance concepts for gas turbine engines have emerged due to the limitations of the traditional safe
life design concept, where only 1 in 1000 components is expected to develop a small fatigue crack at the end
of a safe life period, Figure 18-1. The remaining 999 components are needlessly retired in a crack-free
condition with a large amount of potential life still available. Many critical rotating components, such as
turbine discs and spacers, are treated in this wasteful manner.
By applying a safety-by-inspection life cycle management approach, which relies on predictions of crack
growth life and Non-Destructive Evaluation (NDE) of components at overhaul, safety of these components
can be assured and significant cost savings can be realized. This is often referred to as the Damage Tolerance
Based Life Cycle Management (DTLCM) concept, which is illustrated in Figure 18-2. It is shown that at the
end of a Safe Inspection Interval (SII), all components are inspected and then the components in which
cracks indications were not found can be returned to service for another SII. This procedure is repeated until
a crack is found. In this approach the components are retired based on their individual condition. DTLCM
procedures assume that:
1) Flaws exist in manufactured parts;
2) They are located in the fracture critical locations of the components; and
3) Their sizes are just below the detection limit of the NDE technique used to inspect the components
[10], [11].
The flaws are predicted to grow during service at a rate governed by the local stress distribution and
operating environment for the parts in a manner that can be predicted by Linear Elastic Fracture Mechanics
(LEFM). A dysfunction limit is determined, beyond which the risk of failure due to rapid crack growth
becomes excessive. The process of crack growth from an initial size (ai) to a dysfunction crack size (ad) in a
18 - 6 STO-TR-AVT-275
AGING OF ENGINE STRUCTURES
component is usually established numerically, based on the best estimates of the service loads and material
properties. The results generate a Crack Propagation Interval (CPI), which forms the basis for SII:
1/2 CP
Crack size
Crack found
Safe Life Retire
Usage (Cycles)
The US ENSIP damage tolerance approach [17] uses quantitative measures of the maximum crack size that
may be missed during depot level inspection as the starting point for damage tolerance analysis. Probability
Of Detection (POD) information is also considered in the damage tolerance analysis. The initial surface
crack size used in the analyses usually corresponds to 90% POD with 95% confidence (90/95 POD) for NDE
techniques. In addition, the Probability Distribution Function (PDF) of size for cracks missed during
inspection is used as the starting condition for each simulation, which is the initial crack size ai. [10], [11].
Deterministic Fracture Mechanics (DFM) calculations, which represent a worst case scenario, are used to
predict SII. Accordingly, Probabilistic Fracture Mechanics (PFM) is used to quantify risk, by simulating the
consequences of missing a crack during inspection. The PFM approach uses the LEFM principles for
calculating a , but input parameters such as initial crack size, dysfunction crack size, and constants of the
Paris equation are treated as random variables.
A flow diagram of the DFM/PFM methodology is presented in Figure 18-3. These procedures, based on
DFM and PFM, take into account material data, POD and distribution of initial crack sizes missed by
particular NDE techniques. In addition, curves describing the Stress Intensity Factor (SIF) vs. crack size are
used. Simulation of the crack growth from an initial crack size to the dysfunction crack size is performed and
simulation results are fitted with probability distribution functions [10], [11].
Each set of simulated failure data obtained in PFM analyses is evaluated statistically. Two or three parameter
Weibull, lognormal or gamma distributions are used as Probability Distribution Functions (PDF) candidates
which fit the simulated data. Probability-plot-correlation-coefficient plots are used to initially estimate shape
parameters for Weibull or gamma distributions. This determines which specified distribution family provides
the best fit to the simulated failure data. Next, three distributions are fitted to the data using both the least
STO-TR-AVT-275 18 - 7
AGING OF ENGINE STRUCTURES
square and the maximum likelihood methods. Finally, the Anderson-Darling goodness-of-fit test is
performed for fitted PDF curves. This test is a modification of the Kolmogorov-Smirnov (K-S) test which
gives more weight to the tails than the K-S test [18], [19]. In-house software packages are used to perform
disc failure simulations, while MATLAB® and in-house statistical analysis software are used to analyse the
simulation results [10], [11].
Figure 18-3: Flow Diagram of the Damage Tolerance Assessment Algorithm [10], [11].
18 - 8 STO-TR-AVT-275
AGING OF ENGINE STRUCTURES
Both types of analysis, DFM and PFM, require the stress and thermal analyses of components, which are
performed at NRC-Aerospace using commercial Finite Element (FE) packages such as MSC.Nastran,
MSC.Marc and ABAQUS. Meshing for the FE analyses is prepared with MSC.Patran and Femap. Following
linear elastic calculations, elastic-plastic analyses are performed to establish the redistribution of stresses
under operational loads. In other companies and research organizations, different software and different
methods are used to perform component lifing analyses, and damage tolerance analyses, e.g., Refs. [20],
[21], [22], [23], [24], [25], [26] and [27]. Material data is obtained using material testing performed on
service exposed materials using appropriate test methods, e.g., [28].
Some assumptions have to be made in the damage tolerance assessment of old hardware due to a lack of
detailed technical data. In the approach reported here it is assumed that:
• Cracks in the component are large enough to grow according to the Paris relation;
• For 2D calculations, growth of the three-dimensional cracks is governed by the stress intensity
factor at the deepest point;
• Only a one dimensional stress state is driving propagation of the crack (mode I);
• Effect of minor cycles can be neglected;
• Temperature distribution in a component can be inferred from known operational temperatures;
• Only cracks emanating from surfaces are allowed to grow; and
• Effects of compressive residual surface stresses can be neglected.
Conservative safety factors are used during the analysis to compensate for these simplifying assumptions.
In addition, for the practical application of DTLCM in the field, a final safety factor on predicted SII is also
defined. This safety factor is based on confidence in the analysis, material property data, engine usage, and
inspection capability. The value of the safety factor is based on an overall confidence in the entire
DT process, and is typically between 2 and 5 [10], [11].
The Applied Vehicle Technology (AVT) Panel of NATO Science and Technology Organization
(NATO STO) prepared a number of reports and conference proceedings on gas turbine engine lifing, damage
tolerance and required instrumentation, [28], [29], [31], [32], [33].
Using these algorithms, damage tolerance analyses were performed at the NRC-Aerospace for several
components of gas turbine engines operated by Royal Canadian Air Force (RCAF). Deterministic Fracture
Mechanics (DFM) methods were used to assess damage tolerance of a part containing a crack at the detection
limit of the inspection techniques used in the field, while Probabilistic Fracture Mechanics (PFM) methods
were used for assessment of the risks associated with the application of various inspection strategies.
Components from a few engines powering military aircraft, among them J85, Nene X and T56, were assessed,
[10], [11], [34]. The damage tolerance methods developed were transferred to the RCAF airworthiness
authority for military aircraft, who have implemented some of the recommendations for the maintenance and
operation of the military gas turbine engines. The damage tolerance methods developed contributed to the
Canadian Engine Structural Integrity Program (ESIP) requirements promulgated for the RCAF [35], [36].
18.3 REFERENCES
[1] NATO RTO: “Recommended Practices for Monitoring Gas Turbine Engine Life Consumption”,
RTO Technical Report 28, RTO-TR-28, (AC/323(AVT)TP/22), 2000.
[2] Eady, C., “Modes of Gas Turbine Component Life Consumption”, Chapter 4 in Recommended Practices
for Monitoring Gas Turbine Engine Life Consumption, RTO Technical Report 28, RTO-TR-28,
(AC/323(AVT)TP/22), NATO RTO, 2000.
STO-TR-AVT-275 18 - 9
AGING OF ENGINE STRUCTURES
[3] Beres, W., “Mechanics of Materials Failure”, Chapter 5, in Recommended Practices for Monitoring
Gas Turbine Engine Life Consumption, RTO Technical Report 28, RTO-TR-28,
(AC/323(AVT)TP/22), NATO RTO, 2000.
[4] NATO RTO Workshop on Life Management Techniques for Aging Engines; Aging Mechanisms and
Control, Symposium B: Monitoring and Management of Gas Turbine Fleets for Extended Life and
Reduced Costs, Symposium held in Manchester, UK, October 2001, RTO-MP-079(I), 2003.
[5] Beres W., and Koul, A.K., “Damage Tolerance Based Life Assessment for Aging Nene 10 Turbine
Discs”, NATO RTO Workshop on Life Management Techniques for Aging Engines, Manchester, UK,
October 2001, Aging Mechanisms and Control, Symposium B: Monitoring and Management of Gas
Turbine Fleets for Extended Life and Reduced Costs,” RTO-MP-079(I), pp. 14.1-14.16, 2003.
[6] Beres W., and Koul, A.K., “Stress Intensity Factor Calculations for Cracks Emanating from Bolt Holes
in a Jet Engine Compressor Disc”, 23rd International Congress of Aeronautical Sciences, Toronto,
Canada, September 8 – 13, 2002, Paper 0433, pp.1-8.
[7] Immarigeon, J.-P., Beres, W., Au, P., Fahr, A., Wallace, W., Koul, A.K., Patnaik, P., and
Thamburaj, R., “Life Cycle Management Strategies For Aging Engines”, NATO RTO Workshop on
Life Management Techniques for Aging Engines, Manchester, UK, October 2001. “Aging
Mechanisms and Control, Specialists’ Meeting on Life Management Techniques for Ageing Air
Vehicles”, RTO-MP-079(II), pp. 17.1-17.16, 2003.
[8] Immarigeon, J.-P., Koul, A.K., Beres, W., Au, P., Fahr, A., Wallace, W., Patnaik, P., and
Thamburaj, R., “The Aging of Engines: An Operator’s Perspective”, NATO RTO Lecture Series 218
on “Aging Engines, Avionics, Subsystems and Helicopter,” RTO-EN-14, 2000, pp. 2.1-2.20.
[9] Wu, X., Beres W., and Yandt, S., “Challenges in Life Prediction of Gas Turbine Critical Components”,
CASI Canadian Aeronautical and Space Institute Aero 2007 Conference, Toronto, ON, April 23 – 26,
2007, Paper 413, CASI-2007-0016, 2007.
[10] Beres, W., Dudzinski, D., Robertson, S., and Prentis, C., “Damage Tolerance Assessment of Ageing
Gas Turbine Engines through Analyses and Testing,” NATO RTO Symposium AVT-157 “Ensured
Military Platform Availability,” Montreal, Canada, October 13 – 17, 2008, Paper 19, pp. 19-1 to 19-22.
[11] Beres, W., Kearsey, R.M., Wu, X.J., Yang, Q., Dudzinski, D., and Tsang, J., “Research on Advanced
High Temperature Materials and Protective Coatings for New and Legacy Gas Turbine Engines for
Military Platforms”, Design, Modelling, Lifing and Validation of Advanced Materials in Extreme
Military Environments, NATO STO Applied Vehicle Technology Panel (AVT) Symposium held in
Biarritz, France from 15 – 18 October 2012, MP-AVT-187, Paper 6, pp 6.1-6.28.
[12] Wu, X., “Evolution of Life in Metallic Materials – A Holistic Approach to Material Life Prediction”,
NRC, Institute for Aerospace Research, Structures and Materials Performance Laboratory Report
No. AN-SMPL-2007-0070, Ottawa, 2007.
[13] Koul, A.K., Tiku, A., Bhanot, S., and Junking, B., “Improving Component Life Prediction Accuracy
and Reliability through Physics Based Prognosis – A Probabilistic Turbine Blade Case Study”, ASME
Turbo Expo 2008: Power for Land, Sea and Air, June 9 – 13, 2008, Berlin, Germany, GT2008-51526.
[14] Koul, A.K., Tiku, A., Bhanot, S., and Junking, B., “Importance of Physics-Based Prognosis for
Improving Turbine Reliability – RRRA 501KB Gas Turbine Blade Case Study”, Proceedings of
POWER2007 ASME Power 2007, July 17 – 19, 2007, San Antonio, Texas, USA, POWER2007-22064.
18 - 10 STO-TR-AVT-275
AGING OF ENGINE STRUCTURES
[15] Bird, J., Wu, X., Patnaik, P., Dadouche, A., Létourneau, S., and Mrad, N., “A Framework of Prognosis
and Health Management – a Multidisciplinary Approach,” ASME Turbo Expo 2007: Power for Land,
Sea and Air GT2007-27953, May 14 – 17, 2007, Montreal, QC, Canada.
[16] Bird, J., Wu, X., Patnaik, P., Dadouche, A., Létourneau, S., and Mrad, N., “A Multidisciplinary
Approach to Diagnostics Prognosis and Health Management of Military Gas Turbine Engines”,
Ensured Military Platform Availability, NATO-RTO-AVT Symposium AVT-157, Paper No. 27,
Montreal, QC, Canada, October 13 – 17, 2008.
[18] Lawless, J.F., Statistical Models and Methods for Lifetime Data, Wiley, 1982.
[21] Ingraffea, A.R., and Wawrzynek, P., “FE Methods for Linear Elastic Fracture Mechanics”, in
Comprehensive Structural Integrity, Vol. 3: Numerical and Computational Methods, Chapter 3.01,
Elsevier, 2003.
[22] Harter, J.A., AFGROW User’s Manual and Technical Guide: Version 4.0008.12.11, AFRL-VA-WP-
TR-2003, Air Force Research Laboratory, 2003.
[23] The Cornell Fracture Group, Cornell University, FRANC3D Concepts User’s Guide V2.6, 2003.
[24] Southwest Research Institute, DARWIN User’s Guide, San Antonio, TX, 2007.
[25] “User Manual for ZENCRACK 7.5”, Zentech International Ltd, 2007.
[26] Beres, W., Fread, D., Harris, L., Haupt, P., Kappas, J., Olson, R., Reineke, P., Robertson, S., and
Stocks, G., “Critical Components Life Update for Gas Turbine Engines – Case Study of an
International Collaboration”, ASME Turbo Expo 2008: Power for Land, Sea and Air June 9 – 13, 2008,
Berlin, Germany, GT2008-50655.
[27] Hou, J., Dubke, J., Barlow, K., Slater, S., Harris, L., Calcuttawala, S., and Beres, W., “3D Crack
Growth Analysis and its Correlation With Experiments for Critical Turbine Components Under
an International Collaborative Program”, ASME Turbo Expo 2008: Power for Land, Sea and Air,
June 9 – 13, 2008, Berlin, Germany, GT2008-50548.
[28] Standard Test Method for Measuring Fatigue Crack Growth Rates, ASTM Standard E647-05, 2005.
[29] NATO RTO Lecture Series 218 bis Aging Aircraft Fleets: Structural and Other Subsystem Aspects,
NATO RTO-EN-015, 2001.
[30] NATO RTO: “Improving Military Engine Reliability”, NATO Report RTO-TR-AVT-126,
AC 323/TP-289, 2010.
[31] NATO RTO: “Ensured Military Platform Availability”, NATO RTO Meeting Proceedings MP-AVT-157,
AC323/TP-235, 2008.
STO-TR-AVT-275 18 - 11
AGING OF ENGINE STRUCTURES
[32] NATO STO: “Continuing Airworthiness of Ageing Aircraft Systems”, NATO STO Meeting
Proceedings MP-AVT-222, AC323/TP-590, 2014.
[33] “Test Cell and Controls Instrumentation and EHM Technologies for Military Air, Land and Sea
Turbine Engines”, MP-AVT-229, AC323/TP625, 2015.
[34] Beres, W., Dudzinski, D., and Murzionak, A., “Fatigue Crack Growth Rate Evaluation in a Turbine
Disc after Spin Rig Testing”, 12th International Conference on Fracture, Ottawa, ON, July 12 – 17,
2009, Paper 00587, Session T12.006, pp. 1-9.
[35] RCAF: “Engine Structural Integrity Program (ESIP) Requirements”, RCAF AEPM, 2013.
18 - 12 STO-TR-AVT-275
Part 4: AIRCRAFT MECHANICAL SYSTEMS,
AND OTHER SUB-SYSTEMS
STO-TR-AVT-275 Part 4 - i
Part 4 - ii STO-TR-AVT-275
Chapter 19 – AIRCRAFT MECHANICAL
EQUIPMENT AND SUBSYSTEMS
Matthew Bozzuto
United States Air Force
UNITED STATES
19.1 INTRODUCTION
This section will address the United States Air Force’s (USAF) approach to continued airworthiness for
aircraft mechanical equipment and subsystems. The USAF uses a structured systems engineering process as
described in MIL-STD-1798 [1] to manage aircraft mechanical equipment and subsystems. The majority of
program offices execute a Mechanical Equipment and Subsystems Integrity Program (MECSIP). MECSIP is
a cradle-to-grave task-based approach with the goal of achieving the desired level of safety and aircraft
availability at the most economic cost across the life cycle of the weapon system. In addition to MECSIP,
key elements of sustainment and airworthiness make up the current state of subsystems management practice
in the USAF.
Prior to 2005, though the MECSIP standard existed as an available method for subsystems management,
it was not called out as a requirement in USAF directive publications and it lacked sufficient detail for
effective sustainment. In 2005, Air Force Instruction (AFI) 63-101/20-101 began requiring each program
office to execute integrity programs for their aircraft. Only recently developed aircraft from the mid-2000s
to today had a MECSIP requirement in development. In contrast, almost all USAF aircraft in use today had a
requirement for ASIP from the start. The MECSIP requirements are primarily a collection of best practices,
the result of individual program lessons learned on subsystems management before MECSIP was
a requirement.
MECSIP has grown to incorporate the roles of materials, design, manufacturing, testing, usage, personnel,
and repair/overhaul factors, providing sufficient insight into component failures to enhance readiness, safety,
and sustainment cost for the USAF.
MECSIP includes five sequentially ordered tasks, which are summarized below and described in detail in
MIL-STD-1798 [1]:
1) Task I: Preliminary Planning – The purpose of Task I is to scope the tailoring, planning, and
development strategy for applying MECSIP;
2) Task II: Design Information – This task encompasses the efforts required to identify and
understand all technical criteria that apply to the initial design, development, materials,
manufacturing processes, and production planning for each specific system or equipment
STO-TR-AVT-275 19 - 1
AIRCRAFT MECHANICAL EQUIPMENT AND SUBSUSTEMS
application. A key effort within Task II is the classification of components and identification of
appropriate control mechanisms for all safety-critical components and those mission-critical
components which must not fail in service;
3) Task III: Design Analysis and Development Tests – Analyses and development tests support the
Design Control Activity and verify compliance with performance, functional, and integrity
requirements;
4) Task IV: Component Development and System Functional Tests – These tests are intended to
verify the subsystem integrity performance and to validate design verification analysis. Test scope
includes subsystems or individual components, in simulated subsystem installation environments or
during flight and ground testing; and
5) Task V: Sustainment, Force Management – Force management includes those actions necessary
to ensure that the performance, safety, reliability, and durability requirements established in Tasks I
through IV are met and maintained throughout the entire life of the weapon system.
MECSIP addresses the unique characteristics of subsystem components while borrowing some successful
program elements and organization from ASIP. Like ASIP, MECSIP uses a task-based approach and
requires identification and management of safety-critical components to prevent catastrophic failures.
MECSIP also takes it a step further, incorporating a focus on improving management of non-safety-critical
components to reduce aircraft downtime (approximately 40% of aircraft downtime is attributable to
mechanical equipment and subsystems). A second major difference between MECSIP and ASIP is the high
number of failure mode types that are characteristic of subsystem components as compared to crack and
fracture characteristics in aircraft structure. Understanding of the relationship between component usage,
maintenance, overhaul, and cost often lacks the level of precision necessary to prevent unscheduled
maintenance on non-critical components. While Reliability-Centered Maintenance and Condition Based
Maintenance Plus are helping in this area, progress is hindered by a lack of Item Unique Identification
technologies for parts marking and tracking across the USAF enterprise. Historically, the chief focus of
integrity programs has been to identify, manage, and prevent failure of safety-critical parts, defined as those
parts whose proper functioning is essential to safe aircraft operation. MECSIP goes further by positively
impacting aircraft availability and sustainment cost through management tasks for all mechanical equipment
and subsystem components regardless of safety-critical priority.
The initial release of MECSIP guidance was as a military standard, MIL-STD-1798, in 1988. As part of the
evolution of DOD acquisition philosophy, the military standard was revised and converted into a handbook
in 1997, then converted back to a military standard in 2008.
In the mid-1990s, the KC-135 program, along with a few other transport aircraft programs, suffered
component failures that resulted in catastrophic mishaps. In response, the Functional Systems Integrity
Program (FSIP) emerged as a comprehensive integrity process with the objective of proactively ensuring
operational safety, suitability, and effectiveness of aircraft functional systems. The FSIP concept combined
standards from existing, standalone integrity programs, mainly MECSIP and the Avionics Integrity Program
(AVIP), and added sustainment tasks to meet this new purpose.
19 - 2 STO-TR-AVT-275
AIRCRAFT MECHANICAL EQUIPMENT AND SUBSUSTEMS
In the mid-2000s, catastrophic mishaps on T-38s and a B-52 prompted a greater emphasis on sustainment
in the mechanical equipment and subsystems integrity domain. For reference, in 2004 more than 50% of
major programs had no formal subsystems management plan. Those programs with formal plans used five
different types of plan. In 2009, USAF engineering began annually reviewing all aircraft programs against
MIL-STD-1798 regardless of program origin or status.
In addition to reviews, the USAF began hosting an annual conference targeted at creating cross
communication and best practice sharing opportunities among USAF aircraft program offices, industry,
other USAF organizations (using commands, logistics, etc.), and other services to foster growth and
innovation in the MECSIP community.
Management of mechanical equipment and subsystems has continued to evolve over the past decade.
In 2010, MIL-STD-1798B incorporated sustainment tasks as used by transport aircraft FSIPs. In 2013,
MIL-STD-1798C expanded the sustainment requirements for mechanical equipment and subsystems to
encompass aircraft availability and cost in addition to the traditional focus on safety [1].
In 2018, the Air Force clarified and strengthened the requirement in AFI 63-101/20-101 that all programs
must apply and tailor MIL-STD-1798 [4]. Over time, programs are shifting their management processes to
align with MECSIP even though the majority of them originally included no MECSIP requirements. Today,
nearly all USAF programs have a MECSIP manager and MECSIP master plan integrated into their
mechanical equipment and subsystems management processes. Through the review process, programs are
encouraged to comply with the latest revision of MIL-STD-1798 within two years of publication. All new
programs entering development are required to include MECSIP requirements for all phases from
development through sustainment. The class A and B combined mishap rate for manned aircraft due to
mechanical equipment and subsystems causes has steadily declined for more than a decade.
STO-TR-AVT-275 19 - 3
AIRCRAFT MECHANICAL EQUIPMENT AND SUBSUSTEMS
Figure 19-1 shows the steps required for completing the five MECSIP tasks. The MECSIP master plan is
the foundation of each MECSIP, documenting what the program is doing to meet the requirements of
MIL-STD-1798. Updates to the MECSIP master plan occur regularly throughout the life of the system. Plan
updates occur during each task of engineering development. In sustainment, plan updates are precipitated by
various events such as changes in design, usage, or sustainment philosophy, and MIL-STD-1798 revisions.
The MECSIP master plan is tailorable by determining requirement-by-requirement applicability for each
aircraft. Any requirements judged non-applicable must have appropriate justification for exclusion from the
MECSIP plan.
Figure 19-1: Steps Required for Completing the Five MECSIP Tasks.
Effective mechanical equipment and subsystems component management enables programs to optimize
aircraft integrity and availability. The MECSIP tasks classify components according to safety-, mission-, and
cost-criticality. These categories have implications for aircraft safety, usage, and availability, as well as
sustainment cost and maintenance/repair procedures and timelines. Key enablers are engineering data
availability; component tracking infrastructure and monitoring; data collection and analytics capabilities;
field and depot maintenance documentation; and effective reliability analysis.
USAF programs have made significant progress in the past decade in all areas mentioned above by
implementing functional systems integrated database software solutions that provide the program office with
a software layer for managing mechanical equipment and subsystems issues. The database solutions facilitate
effective documentation and management of maintenance data, fleet configurations, and technical requests,
among other capabilities. Currently, there are three major database solutions in use across the USAF, with
varying governance processes. One very successful database solution has grown rapidly recently as a result
of a shared approach for funding upgrades. The aircraft program leaders voluntarily meet monthly and work
ongoing projects funded by program office money. Program offices share the cost of Cybersecurity and
19 - 4 STO-TR-AVT-275
AIRCRAFT MECHANICAL EQUIPMENT AND SUBSUSTEMS
storage. This governance process results in government-owned software modifications being made available
to all participating programs. It has proven to be a very dynamic, responsive, and successful approach to
building robust functional systems integrated database solutions while sharing the cost through collaboration.
In the future, the USAF plans to manage these database applications at the enterprise level through an
ongoing database consolidation, which will make them more accessible to all aircraft programs.
STO-TR-AVT-275 19 - 5
AIRCRAFT MECHANICAL EQUIPMENT AND SUBSUSTEMS
will have limited utility until already installed components are marked, since the unused portion of life for
those components will remain unknown, clouding insight into future events. In the long term, however, the
Air Force would benefit in the safety, aircraft availability, and cost domains for mechanical equipment and
subsystems by implementing a system analogous to the Individual Aircraft Tracking System used to good
effect by the Aircraft Structural Integrity Program.
For example, when a program office identifies an item of concern in a commodity division, they must have a
point of contact in order to pursue a solution to the problem, since it resides outside their control and chain of
command. This is a growth area for the USAF and a point of interest for MECSIP effectiveness. Identifying
a point of contact who is knowledgeable, possesses the appropriate level of authority, and is willing to act is
a critical first step toward solving day-to-day problems efficiently. ASIP has excelled in this area and has
multiple avenues for delineating roles and responsibilities across the USAF, as opposed to MECSIP’s single
focus on program office responsibilities.
19 - 6 STO-TR-AVT-275
AIRCRAFT MECHANICAL EQUIPMENT AND SUBSUSTEMS
The USAF uses a design-based airworthiness process with the following basic characteristics as described in
AFI 62-601 [7]:
1) The airworthiness process is a higher-level process for new aircraft design and reportable (high-risk)
design changes, but is not recurring on a regular basis;
2) The Technical Airworthiness Authority, which is independent of the program office, approves flight
operation of an air system configuration and the conditions of that operation. The Technical
Airworthiness Authority issues Military Type Certificates (MTC) and other flight releases;
3) Once the Airworthiness process is completed and the MTC or other flight release issued, the
program office has Systems Engineering responsibilities (Operational Safety, Suitability, and
Effectiveness (OSS&E)) for the subsystems;
4) The program office assesses airworthiness of minor design changes; and
5) Aircraft program offices manage aircraft subsystems in sustainment. The Technical Airworthiness
Authority retains authority over major design and operation changes during the lifecycle of the
system.
a) Reportable design modifications;
b) Major modifications or changes to flight envelope and planned usage;
c) Hazards identified through mishap investigations; and
d) Discrepancies found during airworthiness audits.
19.7 SUMMARY
The United States Air Force uses a design-based airworthiness process that factors in the intended mission,
usage, and operating environments. Throughout the lifecycle of a system, the program office is responsible
for its operational safety, suitability, and effectiveness. MECSIP is the Air Force’s process for ensuring
integrity of mechanical equipment and subsystems throughout the lifecycle of the aircraft from early
acquisition through disposal. MIL-STD-1798 includes five MECSIP tasks that support airworthiness
certification during the development phases and long-term integrity of the system in sustainment.
19.8 REFERENCES
[1] US Department of Defense, Department of Defense Standard Practice MIL-STD-1798C, Mechanical
Equipment and Subsystems Integrity Program (MECSIP), 8 August 2013.
[3] Kinzig, W., “USAF Strategy for Aging Aircraft Subsystem Research and Development”, NATO
Report RTO-MP-079(II), Life Management Techniques for Ageing Vehicles Meeting (Presented
8-11 October 2001), 2003.
STO-TR-AVT-275 19 - 7
AIRCRAFT MECHANICAL EQUIPMENT AND SUBSUSTEMS
[4] US Air Force, Air Force Instruction (AFI) 63-101/20-101, Integrated Life Cycle Management, 9 May
2017.
[7] US Air Force, Air Force Instruction (AFI) 62-601, USAF Airworthiness, 11 June 2010.
19 - 8 STO-TR-AVT-275
Chapter 20 – AIRCRAFT MECHANICAL SYSTEMS
INTEGRITY PROGRAM IN CANADA
Sean Leithead
Royal Canadian Air Force
CANADA
20.1 INTRODUCTION
This chapter discusses the development and implementation of an Aircraft Mechanical Systems Integrity
Program (MSIP) for Canada’s Royal Canadian Air Force (RCAF) aircraft fleets. Requirements for MSIP
were first formally introduced in Canada’s Department of National Defence’s (DND) Technical
Airworthiness Manual (TAM) [1] in 2015. In 2018, a Technical Airworthiness Authority (TAA)
Advisory [2] was developed to provide further guidance on MSIP implementation to the various RCAF
aircraft fleets. This Advisory was published in February 2019. In general, MSIP covers the entire lifecycle of
an aircraft: acquisition, in-service, and fleet life extensions if applicable.
STO-TR-AVT-275 20 - 1
AIRCRAFT MECHANICAL SYSTEMS INTEGRITY PROGRAM IN CANADA
The maintenance program developed for the new aircraft acquisition must be robust enough, in terms of
aircraft mechanical systems. This includes the identification and tracking of the following:
1) Safety-critical components and systems;
2) Life-limited components;
3) Appropriate overhaul requirements (based on approved System Safety Analysis techniques and
standards);
4) Appropriate inspection requirements for components and aircraft zonal areas; and
5) Corrosion prevention and control programs.
Proper installation of mechanical systems including routing and clearance of lines and fittings should also be
confirmed. To accomplish this, an on-aircraft Enhanced Zonal Analysis Procedure (EZAP) is required to be
carried out by DND personnel prior to 1st aircraft delivery. Any request to waive this requirement must have
prior approval by the TAA.
20 - 2 STO-TR-AVT-275
AIRCRAFT MECHANICAL SYSTEMS INTEGRITY PROGRAM IN CANADA
3) Capability to search effectively for mechanical systems in-service difficulty trends in the fleet’s
Maintenance Electronic Record Keeping System (ERKS);
4) Creation and execution of Corrective Action Plans to improve aircraft safety and airworthiness
related to aircraft mechanical systems; and
5) When practical (such as during an on-aircraft second line periodic inspection), carry out an EZAP if
one was not completed when the fleet was acquired. The TCH should ensure that any modifications
completed during the aircraft fleet’s in-service phase do not compromise the proper installation of
any mechanical systems, including routing and clearance of lines and fittings, and interactions with
electrical wiring.
STO-TR-AVT-275 20 - 3
AIRCRAFT MECHANICAL SYSTEMS INTEGRITY PROGRAM IN CANADA
2) Ensure continuation of the in-service phase aircraft mechanical systems integrity monitoring
program until the end of the new ELE;
3) Review of components that may become obsolete or unsupportable before the new ELE is reached;
4) Identify certification requirements that need to be met in order to support the ELE and then make
findings of compliance against them. In other words, have a design change certification basis for the
ELE; and
5) Consultation with other country users of the same aircraft fleet, if such countries have undergone
a similar ELE extension or have been using the aircraft fleet longer than DND has.
20.6 SUMMARY
Canada’s Mechanical Systems Integrity Program (MSIP) is focused upon reducing the occurrence of
mechanical system in-service difficulties through preventive maintenance and appropriate system health
monitoring activities. The goal is to provide a means to improve aircraft fleet effectiveness and safety. MSIP
is designed to provide flexibility to the varying in-service support contractors of the RCAF aircraft fleets,
while still providing sufficient oversight and guidance.
20.7 REFERENCES
[1] Department of National Defence, Technical Airworthiness Manual (TAM) C-05-005-001/AG-001,
1 Apr 2019.
[2] Department of National Defence, Mechanical Systems Integrity Program TAA Advisory 2019-01,
available at https://www.canada.ca/en/department-national-defence/services/military-airworthiness/tec
hnical-airworthiness-authority-overview/technical-airworthiness-regulatory-documents/technical-airwo
rthiness-authority-advisories/technical-airworthiness-authority-advisory-2019-01.html., 1 Feb 2019.
20 - 4 STO-TR-AVT-275
Chapter 21 – AIRCRAFT MECHANICAL SYSTEMS INTEGRITY
MANAGEMENT ACTIVITIES IN FINLAND
Ilpo Paukkeri
Finnish Defence Forces
FINLAND
21.1 INTRODUCTION
This chapter provides a brief review on the approach and activities related to aircraft mechanical systems
integrity management in Finland, from the Finnish Defence Forces (FDF) viewpoint. In addition to this
chapter, the report includes dedicated chapters for aircraft structures (Chapter 13) and engines (Chapter 17).
The FDF have not implemented a systematic, comprehensive approach for aircraft mechanical systems
integrity management. However, many activities have been put into practice within the FDF and industrial
partners to manage the integrity of the mechanical systems of the current fleets, especially the F/A-18C/D
Hornet fighters and BAE Hawk jet trainers operated by the Finnish Air Force (FINAF), as well as NH90
rotorcraft operated by the Finnish Army.
In-country design and engineering capabilities have been instrumental in the maintenance of the
airworthiness and mission capability of the system through maintenance, repair, overhaul and modification
programs that are tailored for the national requirements and based on the actual use of aircraft. This work has
been done in close co-operation with other users of the same aircraft type by openly sharing information.
The work is being carried out under the supervision of the Finnish Military Aviation Authority (FIMAA)
that sets a regulatory basis for such activities, the requirements for different organizations and approves
these organizations.
The FIMAA has delegated authority to FDF Logistics Command (FDFLOGCOM) to act as a Type
Certification Holder (TCH) and Continuing Airworthiness Management Organization (CAMO) for the
Finnish military aircraft, with the following tasks:
• Develop airworthiness criteria, and processes and means to show compliance with the airworthiness
criteria;
• Develop and maintain a maintenance program and a structural integrity plan for each type; and
• Maintain a system for data collection, assessment, analysis and reporting related to failures and
defects that pose risks for maintaining airworthiness of the aircraft.
The FDF’s TCH does not have an organic design capability and is typically supported by OEMs and the
domestic partners. For aircraft like the F/A-18, a national design capability has been necessary.
STO-TR-AVT-275 21 - 1
AIRCRAFT MECHANICAL SYSTEMS
INTEGRITY MANAGEMENT ACTIVITIES IN FINLAND
As Finland has a unique operating environment, tasks and usage spectrum, the FDF must have its own
capability to:
• Assess effects of the usage spectrum on the service life and structural integrity of individual
airframe and systems;
• Implement changes and repairs needed to the structure due to differences in actual usage vs. design
usage spectrum; and
• Adjust maintenance requirements and intervals to maintain optimal availability with minimized
costs.
Furthermore, the national capability should enable re-certification of the fleet (component) when extension
of service life is deemed necessary.
MECSIP is described as “a task-based approach with the goal of achieving the desired level of safety and
aircraft availability at the most economic cost across the life cycle of the weapon system”. Regarding
suitable phase to initiate a MECSIP program, it is stated that “it will be easier to comply with MECSIP Force
Management tasks if the program enters the sustainment phase with a robust execution of MECSIP Tasks I
through IV, but completing Tasks I through IV is not essential for a successful MECSIP sustainment
program. Many aircraft programs will initiate their MECSIP Programs in the sustainment phase” [1].
MECSIP Tasks I – IV are by definition related to design, development and test phases; however,
they comprise significant information, requirement base and practices to be applied during the sustainment
phase. As an example, airworthiness certification is based on certification analyses which may need to be
reviewed/updated during sustainment in order to solve integrity issues, particularly when differences are
identified between certification analyses and actual behavior of the component (e.g., measured operational
loads or performance). Considering implementation of MECSIP during sustainment phase, CAMO should
collect all the available data related to the Tasks I – IV.
MECSIP Task V (Sustainment – Force Management) forms the basis for the implementation of MECSIP
for the aircraft already in service, including “those actions necessary to ensure that the performance, safety,
reliability, and durability requirements established in Tasks I through IV are met and maintained throughout
the entire life of the weapon system” [1].
In general, considering possible implementation of the MECSIP standard or similar approach, the following
are considered as the first steps:
• Identify the critical mechanical equipment and subsystems with respect to reliability/availability and
remaining service life;
• Identify the kind of data and knowledge base required to manage integrity of the selected MECSIP
parts; and
• Define processes and methods, as well as roles and responsibilities for MECSIP planning and
execution.
21 - 2 STO-TR-AVT-275
AIRCRAFT MECHANICAL SYSTEMS
INTEGRITY MANAGEMENT ACTIVITIES IN FINLAND
The following sub-sections briefly review Subtasks 1 – 5 of Task V with regard to the existing and foreseen
FDF activities.
Subtask 1: Data gathering and task planning can be considered primarily as the OEM and the main user
responsibility. However, many of the subsequent tasks link back to the information, practices and
requirements set at the initial stage. For example, “developing new MECSIP parts” (under the Subtask 3)
will reach back to updating FMECAs and safety-critical component lists.
Subtask 2: Develop and utilize a functional systems integrated database can be associated with the FDF
activity of developing and utilizing the in-country Logistic Information Management System (LTJ-system)
for monitoring and tracking of components usage, condition and maintenance data. This subtask also refers
to the fundamental task to determine component inspection and/or replacement intervals and criteria.
Actually, the FDF logistic support system facilitate most of the “MECSIP functionalities” by the LTJ-system
as it serves as a primary maintenance database and is utilized extensively for aircraft maintenance and fleet
management actions. However, implementation of MECSIP within the FDF would obviously require the
LTJ-system to be improved and/or accompanied with more active use of both in-country developed and
OEM delivered (analysis) tools.
Subtask 3: Force management execution is described in the standard as “a roadmap on how the fleet’s
components will be managed” and comprises performing tasks through monitoring, risk assessment,
analysis, and actual management of components. Furthermore, the subtask addresses development of new
MECSIP parts and assessment of overall program.
From the FDF perspective, this subtask is essentially about performing assessment, analysis and tests to support
decisions on airworthiness, reparability, service life, maintenance program, and overall fleet management
actions. Considering implementation of MECSIP, a fundamental action at this stage will be to compile the
MECSIP master plan, to be updated regularly, including essential planning and documentation of activities.
In addition to the monitoring systems and programs developed under the previous subtasks, an organization
with sufficient personnel resources and knowhow is required for execution of this subtask. FDFLOGCOM
Joint Systems Centre, Air Systems Division has been organized into responsibility areas covering:
• Life cycle planning;
• Maintenance planning;
• Sleet management; and
• Supply chain management.
The FDF is in the process of implementing the European Military Airworthiness Regulations (EMAR),
particularly EMAR 21 MTCH and Part M CAMO.
Subtask 4: Preventative maintenance actions covers essentially performing MSG-3 analysis in order to
review the maintenance requirements, actions and intervals, resulting in the updated aircraft maintenance
program.
The FINAF F/A-18 Hornet’s maintenance program has been maintained from its early years of service by
setting up a maintenance system development working group, led by the FDFLOGCOM and represented
by maintenance experts from the FINAF squadrons and the strategic industrial partner Patria.
The necessity of the most recent review/update of the maintenance program for the FINAF F/A-18
originated from the increasing maintenance workload due to high number of new structural inspection tasks
STO-TR-AVT-275 21 - 3
AIRCRAFT MECHANICAL SYSTEMS
INTEGRITY MANAGEMENT ACTIVITIES IN FINLAND
based on the main user’s maintenance bulletins. The project aimed at an optimized maintenance program,
determining inspection thresholds and intervals with a special objective to extend the depot level
maintenance intervals that were considered as conservative.
The analyses have been performed following the guidance given in ATA-MSG3 document that defines
the data as the foundation for the analysis for a certain system:
• Fault/failure history and other experience/data related to the system;
• OEMs test data;
• Information from other users/operators; and
• Engineering judgement made by an experienced working group, based on operational experience.
Regarding the basis for the analyses, the LTJ-system holds reported fault indications/findings recorded
during flight or maintenance operations, as well as any performed measurements related to condition of a
mechanical system (e.g., clearances). The failure data can then be assessed as trends over specified time, as
well as distributions of data in terms of in-flight, daily maintenance, and depot level maintenance. There is
also an in-country developed tool for reliability analysis that assists in processing the failure data and
calculates reliability values such as MTBF for equipment/subsystems to be used as reliability parameters for
analysis purposes.
Subtask 5: Manage service life extension and final five years prior to retirement relates closely to the
activity that has been executed for the F/A-18 within the international user group collaboration, e.g., F/A-18
Subsystem Life Extension Program (SLEP) actions together with US Navy and other Hornet users.
Besides research towards larger scale scientific goals, some efforts have also been focused into problem
oriented smaller scale research subjects. The main R&D topics on FINAF F/A-18 Hornet have been [2]:
• Aircraft component failure prognosis related Early Warning System (EWS);
• Modelling and simulation of the hydraulic system of the FINAF F-18 Hornet;
• Modelling and simulation of the main landing gear system of the FINAF F-18 Hornet;
• Detecting and eliminating the measurement error caused by free air in aircraft hydraulic system
on-line particle counting;
• Aircraft hydraulic system fault finding and troubleshooting using on-line particle counters; and
• Process for evaluation and validation of non-original components for aircraft hydraulic systems.
21 - 4 STO-TR-AVT-275
AIRCRAFT MECHANICAL SYSTEMS
INTEGRITY MANAGEMENT ACTIVITIES IN FINLAND
Aircraft component failure prognosis related Early Warning System (EWS) is a project aimed at providing
an extra tool for the maintenance personnel to support the Condition Based Maintenance (CBM) decision
making process. The EWS project is briefly described as follows:
• In practice, EWS tries to forecast the subsystem failures of the FINAF aircraft systems. The starting
point is the process data of the aircraft systems. Thus, what is done at a broad level is data analysis
on the aircraft systems process data. The research of EWS is especially focused on the subset of data
analysis methods called Machine Learning (ML) methods or vernacularly Artificial Intelligence
(AI) methods [2].
• In the EWS project the software for independent operation on the FINAF servers is under
development. Currently the tool is capable of mining relevant parameters out of the data, like
landing gear times from take-offs and landings of the aircraft. Also, some conventional statistical
reliability data analyses and spare part stock estimation with obsolescence aspect research are
ongoing in the project [2].
Modelling and simulation of the main landing gear system of the FINAF F/A-18 is part of the efforts to find
solutions for integrity issues with the landing gear. In general, the objective of modelling and simulation of
certain mechanical systems/components is to better understand the functioning of the systems and their
failures. The F/A-18 main landing gear is a good example of a critical mechanical system that has been a
subject for intensive modelling and simulation efforts. The studies and results are briefly described as
follows:
• The studies were initiated by the FINAF in order to investigate reasons for landing gear failures
(so-called Planing Link failures) that had caused several aircraft landing mishaps in Finland and
elsewhere. The research focused on creating a realistic simulation model to be able to assess the
main landing gear function under different landing conditions and the effect of different landing
gear components. The landing gear model consists of the nose landing gear and both main landing
gears allowing asymmetric touchdown simulation capability. To achieve realistic approach flight
path (including ground effect) and desired aircraft orientation, in-country developed flight
simulation software with the F/A-18C Hornet aircraft model had been used in conjunction with the
landing gear model [3].
• A conclusion from the studies was that there was not a single cause for Planing Link failures, but it
was a combination of causes related to rigging, landing parameters and the condition of the MLG.
The solution focused on reducing clearances of the MLG, to be achieved by incorporating
in-country changes to wear limits and manufacturing tolerances of bushings and bearings, including
in-country modified axle-lever bushings. The capability for machining the bushings was set up in
the FDF’s industry partner Patria to meet the tightened tolerances. Furthermore, periodic overhaul
for the landing gears was implemented in order to reduce the risk of failures.
• More recently, the necessity of re-evaluating the landing gear simulation results has become
apparent, as several upgrades/modifications have been completed on the FINAF F/A-18 Hornet
fleet. The goals of the ongoing study are to find out the effect of mass distribution change on landing
behavior, to improve the landing gear model realism by coupling it to a complex shock absorber
model, and to find out stress concentration locations of three main parts: Trunnion, Lever and Lower
Sidebrace [4].
21.5 CONCLUSION
The FDF is investigating implementation of a systematic, comprehensive approach for aircraft mechanical
systems integrity management, with the intention of avoiding duplication of effort by examining the
generally known standards such as the MIL-STD-1798 (MECSIP). The establishment of integrity
STO-TR-AVT-275 21 - 5
AIRCRAFT MECHANICAL SYSTEMS
INTEGRITY MANAGEMENT ACTIVITIES IN FINLAND
management processes along with associated critical capabilities is considered crucial to supporting the
FDF’s Type Certificate Holder (TCH) and Continuing Airworthiness Management Organization (CAMO)
functions.
The initial review of MECSIP tasks suggests at least the following activities that could support the
implementation and execution of the mechanical systems integrity management for the FDF fleets:
• Develop the integrated logistic information management system accompanied with analysis and
reporting tools, in order to facilitate the MECSIP functions;
• Conduct the process to revise and update the aircraft maintenance program based on collected data
and information from in-service operation, for example by using MSG-3 analysis and an
experienced working group assisted with data analytics tools;
• Develop data analytics methodology and tools in order to exploit the increasing amount of collected
data with a potential to provide automatized, intelligent and efficient support for decision making;
• Make use of modelling and simulation to improve the understanding of critical mechanical systems
function/failures; and
• Promote active participation to international user groups and R&D forums in order to share best
practices and technology advancements along with actual collaborative research efforts.
Regarding best practices, the experience gained from implementing Aircraft Structural Integrity Program
(ASIP) may provide some applicable practices when considering implementing a Mechanical Equipment and
Subsystems Integrity Program (MECSIP). For example:
• Identify all the information related to design and development phases, and incorporate them in the
MECSIP to be assessed when the need arises during the sustainment;
• Develop and maintain an up-to-date MECSIP plan/database for the fleet, documenting all the
relevant information and actions; and
• Evaluate regularly the overall effectiveness of the MECSIP and prepare an action plan to improve
the identified focus areas.
21.6 REFERENCES
[1] Department of Defence, MIL-STD-1798C, Department of Defence Standard Practice: Mechanical
Equipment and Subsystems Integrity Program (MECSIP), 8 August 2013. Available at:
http://everyspec.com/MIL-STD/MIL-STD-1700-1799/MIL-STD-1798C_47320/.
[2] Viitanen, T., and Siljander, A. (Eds.), A Review of Aeronautical Fatigue Investigations in Finland
March 2015 - March 2017, Presented at the 35th Conference of the International Committee on
Aeronautical Fatigue and Structural Integrity (ICAF), Nagoya, Japan, 5-6 June 2017. Available at:
http://www.vtt.fi/inf/julkaisut/muut/2017/ICAF_Doc2433_Finland_Review_2017.pdf.
[3] Siljander, A. (Ed.), A Review of Aeronautical Fatigue Investigations in Finland May 2007 - April
2009, Presented at the 31st Conference of the International Committee on Aeronautical Fatigue
(ICAF), Rotterdam, the Netherlands, 25 ‒ 26 May 2009. Available at: http://www.vtt.fi/inf/julkaisut/
muut/2009/ICAF_Doc2418.pdf.
[4] Viitanen, T., and Siljander, A. (Eds.), A Review of Aeronautical Fatigue Investigations in Finland
April 2017 - March 2019, Presented at the 36th Conference of the International Committee on
Aeronautical Fatigue and Structural Integrity (ICAF), Krakow, Poland, 3-4 June 2019. Available at:
https://cris.vtt.fi/files/24743703/ICAF_Finland_Review_2019.pdf.
21 - 6 STO-TR-AVT-275
Chapter 22 – EWIS MONITORING PROGRAM IN CANADA
Research has also demonstrated that wiring can be harmed by collateral damage when maintenance is being
performed on other aircraft systems. For example, a person performing an inspection of an electrical power
centre or avionics compartment may inadvertently cause damage to wiring in an adjacent area.
In recent years, the regulatory authorities and industry groups have come to the realization that current
maintenance practices may not be adequate to address aging non-structural systems. While age is not the sole
cause of wire degradation, the probability that inadequate maintenance, contamination, improper repair, or
mechanical damage causing degradation to a particular Electrical Wiring Interconnection System (EWIS)
increases over time. Studies by industry and government agency working groups have found that, although
EWIS management is an important safety issue, there has been a tendency to be complacent about EWIS
(see paragraph 5 – Related Reading Material – in Ref. [1]). These working groups have concluded that there
is a need to manage EWIS, so that they continue to function safely.
To address the effects of aging on EWIS, a recommendation on best maintenance practices was created to
mitigate the effects of aging on electrical and wiring systems. As indicated by the acronym, it was based on a
new maintenance philosophy that wire and its associated systems must be viewed as a complete system and
not simply a means of connecting other independent systems. All aircraft are filled with miles of wiring and
hundreds of wiring devices that connect and transfer power and signals to and from electrical components.
Virtually all aircraft systems rely heavily on some type of wiring for safe operation. The health and integrity
of the EWIS can be significantly compromised due to deterioration from aging, compounded maintenance,
damage and the failure of wiring installation. It is integral to the overall maintenance and sustainment of all
aircraft that the EWIS be treated as a system and afforded the same level of importance as the aircraft
structure and other critical flight control systems.
STO-TR-AVT-275 22 - 1
EWIS MONITORING PROGRAM IN CANADA
22 - 2 STO-TR-AVT-275
EWIS MONITORING PROGRAM IN CANADA
All RCAF aircraft fleets have a zonal inspection program within their approved maintenance program, so the
EZAP process essentially was a gap analysis to update the inspection program to reflect current EZAP guidance.
To address design changes, guidance was also provided, the design change categorization process was
amended to require further increased TAA oversight for any change that affected EWIS risk levels or
conversely added mechanical systems that increased risk e.g., new flammable fluid lines. Ref. [1] provides
guidance with respect to assessing any design change adding or modifying EWIS, that should include an
assessment, to determine whether the modification has affected the zone EWIS attributes to warrant
reapplication of the EZAP to the entire zone. Further guidance is available in Appendix B of Ref. [4].
22.1.5 Reporting
Type certificate holders are required to report on their Electrical Wiring Interconnection System Monitoring
Program at a minimum every two (2) years, with the end of the review schedule such that the review data
and results are available for presentation to the Airworthiness Review Board (ARB) for that particular year.
22.2 REFERENCES
[1] Transport Canada Civil Aviation, Transport Canada Civil Aviation (TCCA) TP 14331E – Enhanced
Zonal Analysis Procedures, 30 September 2005, https://www.tc.gc.ca/eng/civilaviation/publications/
tp14331-menu-608.htm.
[2] Department of National Defence, DND Flight Safety Investigation Report – CC130342, 21 February
2012 http://www.rcaf-arc.forces.gc.ca/en/flight-safety/article-template-flight-safety.page?doc=cc1303
42-hercules-epilogue-flight-safety-investigation-report/hky1dmyy.
STO-TR-AVT-275 22 - 3
EWIS MONITORING PROGRAM IN CANADA
[4] Federal Aviation Administration. FAA AC25-27A – Development of Transport Category Airplane
Electrical Wiring Interconnection Systems Instructions for Continued Airworthiness Using and
Enhanced Zonal Analysis Procedure. https://www.faa.gov/regulations_policies/advisory_circulars
/index.cfm/go/document.information/documentID/319197.
22 - 4 STO-TR-AVT-275
Part 5: SUMMARY ON BEST PRACTICE AND LESSONS LEARNED
STO-TR-AVT-275 Part 5 - i
Part 5 - ii STO-TR-AVT-275
Chapter 24 – BEST PRACTICES SUMMARY
K. Jones M. Liao
United States Air Force National Research Council Canada
UNITED STATES CANADA
C. Babish
United States Air Force
UNITED STATES
This report collected 23 chapters on high-level overview of Continuing Airworthiness (CAW) of various
aging systems, they are documented in four Sections on CAW Policies and Approaches (Section I),
Structural Systems (Section II), Propulsion Systems (Section III), Aircraft Mechanical Systems/Subsystems
(Section IV). A number of chapters on case studies are also collected in Section IV, which provided more
detailed technical procedures to maintain structural integrity, specifically through aircraft Operational Loads
Monitoring (Chapter 25), Full-Scale Fatigue Testing (Chapter 26), NDT (Chapter 27 to Chapter 28),
environmental sensing for corrosion (Chapter 29, and failure models and NDI of composite parts
(Chapter 30 to Chapter 31). Overall many chapters provide some excellent references/sources for the readers
to learn more in-depth information.
The United States Air Force (USAF) has numerous written policies and instructions that establish
requirements for aircraft development, production, and sustainment phases. The overall policy is titled
Integrated Life Cycle Management and requires the acquisition chain of authority to apply standard systems
engineering processes and practices to ensure the integrity, mission assurance, operational safety, suitability,
and effectiveness of each system throughout the life cycle from concept development through disposal. The
USAF has issued written policy and instruction that establishes the requirements for formal airworthiness
assessments to ensure that USAF operated aircraft are airworthy over their entire life cycle and maintain
accepted levels of safety. An important aspect of the USAF policy is that it establishes the requirement for an
independent technical airworthiness authority (outside the acquisition chain authorities) to make
airworthiness determinations and issue flight certifications and authorizations [1]. The original USAF
airworthiness certification requirements were primarily aimed at initial airworthiness certification only.
A significant change was made in 2009 when the inclusion of service life limits in aircraft airworthiness
certificates was mandated for all existing and future aircraft. This requirement led to the inclusion of
additional airworthiness certification requirements, most notably in the aircraft structures section of the
Airworthiness Certification Criteria Handbook. The USAF airworthiness process requires a continual
evaluation of fielded aircraft throughout their life cycle to ensure that all aircraft are in a condition for safe
operation, i.e., that they have not exceeded the approved service life limit, that environmental
(e.g., corrosion) or other factors have not degraded airworthiness, that aircraft are properly maintained in
accordance with approved maintenance documentation, and the aircraft are operated in accordance with the
approved flight manual within the approved mission usage. The USAF recognizes that the initial service life
limit is not a fixed value and the airworthiness process includes procedures for determining service life
extension potential and for executing a service life extension program when desired and practical.
STO-TR-AVT-275 24 - 1
BEST PRACTICES SUMMARY
For the RCAF fleets, the technical airworthiness program, which is outlined in the Technical Airworthiness
Manual (TAM), consists of three primary elements: Initial Airworthiness, Continuing Airworthiness, and
Disposal. Initial Airworthiness includes the processes and activities necessary to ensure that aeronautical
products are airworthy (including Type Certification and Equipment Certification), prior to their introduction
into service. The framework of the RCAF CAW includes:
• Conduct and control of maintenance;
• Design change certification;
• In-service configuration management; and
• In-service monitoring program.
The key element of the CAW requirement is to establish an In-service Monitoring Program (ISMP). The
RCAF experience with the Aging Aircraft Assessment (AAA) has shown that the assessment activity needs
to be tailored for each fleet based on how the fleet is operated and maintained. There is a heavy reliance on
the OEM for fleets that are operated in a predominantly civilian role, while a more active involvement is
required for military role fleets. Thus, both military and civil airworthiness standards are used in the TAM.
The level of effort to perform AAA is dependent on the strength of the integrity monitoring program, with a
strong program requiring minimal effort to carry out the AAA. This highlights the importance of the
integrity monitoring program.
The United Kingdom Ministry of Defence (UK MOD) has been undertaking Aging Aircraft Structural
Audits (AASA) since the early 1990s. These audits were initiated following the well-known Aloha Flight
243, Boeing 737 pressure cabin failure in 1988. However, several civil accidents in the mid-1990s, including
TWA Flight 800 and Swiss Air Flight 111, made the aviation community aware of the need to consider
aging effects of electrical and mechanical systems alongside structural implications. For the UK MOD, the
need to understand these aging systems issues was reinforced starkly by the loss of Nimrod XV230 over
Afghanistan in 2006. Thereafter, the inclusion of aircraft systems aspects into the UK MOD’s Aging Aircraft
Audit (AAA) process was introduced. Lessons from these initial systems elements (Aging Aircraft Systems
Audit (AASysA)) and Condition Surveys (CS), including programs conducted on the Nimrod, VC10,
C-130, Sentry, Tucano and Historic Aircraft were collated, with gratefully acknowledged inputs from across
the UK MOD and industrial support base. Common recommendations were identified under the categories
of regulatory or policy issues, program conduct issues and wider airworthiness issues. Another interesting
part is that UK developed a combined AAA approach for different systems (structures, propulsion, and other
subsystems) with common and specific sections. It was proven to be a far more efficient approach, with an
estimated 40% cost saving for the AAA process. Key recommendations from this study were incorporated
into the UK MOD Regulatory Framework and are summarized below:
• To include a requirement to validate the effectiveness of integrity assurance methods by undertaking
a detailed physical condition survey of a sample of aircraft from within each fleet;
• To reinforce the airworthiness requirements of the AAA, rather than the cost of ownership and
availability issues;
• To include a requirement to address the effectiveness of emergency systems within the AAA; and
• To streamline the three independent audits (structures, systems and propulsion) into a single AAA.
From the Polish perspective, the basic Airworthiness (AW) principles are common for all military aviation
stakeholders. However, the detailed approaches and ways of implementation in the particular systems differ
between Nations and users. The Polish Air Force also faces the challenge to operate both post-Soviet and
“western design” aircraft which are certified with different AW regulations. The European Defence Agency
has initiated an activity aiming for development of a common (commonly understood by all EU Nations)
system of AW requirements, and further on regulations, for European military aviation. The EDA Military
24 - 2 STO-TR-AVT-275
BEST PRACTICES SUMMARY
Airworthiness Authority Forum (EDA MAWA Forum) was established to achieve this goal, and a set of
European Military Airworthiness Requirements (EMAR) was developed under the Forum umbrella. The
importance of having a common approach to the AW issues in military domain is very well understood by
all national military AW authorities (or their equivalents) and a process of harmonization, bilateral and
multilateral recognition is continuously in progress. An interesting example is provided on the airworthiness
system of the NE-3A fleet, including 18 early warning aircraft operating within multi-NATO nations. From
the Polish standpoint, the challenge does not necessarily lie in the general understanding of EMAR
requirements, but in the detailed implementation of distinct requirements in polish military airworthiness
regulations in particular EMAR 145 and EMAR M (on CAW) planned for the near future.
The Royal Canadian Air Force (RCAF) was an early adopter of ASIP MIL-STD-1530. An ASIP Master Plan
was mandated for all RCAF fleets. The ASIP Tasks IV and V of the MIL-STD-1530 covered the Continuing
Airworthiness (CAW) requirements for the aircraft structure. Around the same time, the Department of
National Defence (DND) started to formalize their technical airworthiness program and the assembly of the
Technical Airworthiness Manual (TAM). As it was understood that the DND would have very limited
influence on and access to the design information (i.e., design analysis and full-scale testing), it was decided
to NOT mandate the MIL-STD-1530 in its entirety. Rather, it was decided to focus on those tasks for which
the DND had control and would affect the CAW of the aircraft. This was brought under the In-service
Monitoring Program (ISMP) chapter of Part 3 – Continuing Airworthiness of the TAM. Several annexes
were included to cover the various systems of the aircraft. One of these annexes is the Aircraft Structural
Integrity Monitoring Requirements. Since the RCAF aircraft fleets have been certified using different
airworthiness standards (ex. MIL-STD-1530 vs. DEF-STAN-0970) (Table 8-1), the RCAF also gained some
experience to manage structural integrity using two major lifing methods, i.e., Safe Life and
Damage-Tolerant, either in a separate or combined manner. Future challenges include assessing the health of
the execution of the ASIP requirements due to the lack of aircraft engineering design data acquired with the
new fleet acquisitions. Current mitigation is to build into the acquisition and in-service support contact the
ability of the TAA staff to access the engineering data.
In Australia, a Joint Directive signed by Secretary, Department of Defence and Chief of the Defence Force
established the Defence Aviation Safety Framework as of 30 Sep 2016. The framework required the
implementation of a credible and defensible aviation safety framework that recognizes and supports
compliance with statutory safety obligations. Where appropriate the framework was aligned with International
Civil Aviation Organization (ICAO) principles and European Military Airworthiness Requirements (EMAR).
As at 01 Jan 19, the Australian Defence Force (ADF) fully transitioned to the Defence Aviation Safety
Regulation (DASR). Under the previous Technical Regulations, the ADF managed Aircraft Structural Integrity
(ASI) via ASIP MIL-STD-1530, documented within an Authority approved Aircraft Structural Integrity
Management Plan (ASIMP). Under the DASR, the ASIP construct was preserved for existing weapon systems
and established as Acceptable Means of Compliance (AMC) for new acquisitions. Overall the safety record of
the RAAF fleets is improving (Figure 9-1). One of best practices is presented using an Australia developed
enhanced teardown technique to extend structural life while maintaining CAW.
STO-TR-AVT-275 24 - 3
BEST PRACTICES SUMMARY
The enhanced teardown technique was further demonstrated by NLR for extending the life Royal
Netherlands Air Force (RNLAF) F-16 fleet in a cost-efficient and effective manner, as documented in
Chapter 10. Some lessons learned on the application this technique are also documented in this Chapter,
along with a future outlook on a proposal to jointly collect EIDS data for next generation fighter fleet F-35,
which is one of the key inputs for fleet management and CAW. It is also notable that the AFIT developed
some SHM and NDI techniques to efficiently support the ASIP program of their fleets (examples presented
in Section VI (Annexes)).
The Polish AFIT also adopted the USAF ASIP MIL-STD-1530 principles, and applied the process for some
Aircraft Structural Life extension Program (SLAP) for some fleets. The two chapters presented two
examples using the typical OLM, FSFT, and IAT program to extend the structural life of PZL-130 “Orlik”
trainer and Su-22 fighter-bomber.
Finnish Defence Forces (FDF) uses MIL-STD-1530 as a guideline for Aircraft Structural Integrity Program
(ASIP) for Finnish Air Force’s (FINAF) F/A-18 Hornet, BAE Hawk fleets, and the Finnish Army NH90
rotorcraft fleets. The FDF ASIP has evolved as a primary means to manage structural integrity in order to
maintain safety and availability of aging systems and ensure economic service life of the fleets.
Although initiated during the sustainment phase, the FDF ASIP is intended to cover all the tasks defined in
the MIL-STD-1530. Design and development (tasks I-III) information is to be identified as far as possible
and incorporated into the program, to be assessed when the need arises during sustainment.
Increasing structural issues with the aging aircraft pose risks and challenges to meet requirements for
operational availability and economic service life. Major concerns include crack findings on fracture critical
structures for which there is no “full life” modification available and findings on structures that allow very
small critical crack sizes. In addition, exceedance of certified safe life for some components may necessitate
either significant re-certification activities or introduction of new structural inspections to further add to the
increasing workload. These are examples of structural issues that justify the necessity to maintain a
responsive and effective ASIP through the remaining service life.
The United States Army “Basis for Airworthiness” includes analysis and testing during product
development. Many propulsion components are deemed ‘on condition’, whereas the component may remain
in service an indefinite period of time as long as it meets inspection requirements. Some components have a
prescribed retirement life (aka, life limit) due to a limitation in capability, such as Low Cycle Fatigue (LCF).
These life limits are established to maintain safe operation of the component based on data and analyses
available at that time. As the fleet matures into production and fielding, the basis for those life
determinations must be reviewed periodically. Several potential changes that impact the life limit may have
24 - 4 STO-TR-AVT-275
BEST PRACTICES SUMMARY
occurred, including material processing, changes in design, and analysis tool revisions, as well as field
experience and mission requirements. The US Army approach for continued airworthiness is established in
Army Regulation 70-62, Airworthiness of Aircraft Systems.
The Royal Canadian Air Force (RCAF) established The Engine Structural Integrity Program (ESIP) and
requires each Weapon System Manager to maintain defined program to, at regular intervals:
1) Reassess and document actual usage conditions;
2) Compare actual usage conditions against design usage conditions; and
3) Determine life consumption based on actual usage (of critical components, as a minimum)
accordingly.
The ESIP also mandates issue tracking: safety-related issues are managed through the “Record of
Airworthiness Risk Management” (RARM) process documented elsewhere in this report (Chapter 2,
Sections 2.1 to 2.6), which is not a propulsion-specific process. At present, it is not yet possible to ascertain
the effectiveness of the ESIP for the RCAF fleets, as compliance with the regulation is just started this year.
One lesson learned: reassessment of the cyclic exchange rate (low cycle fatigue major cycles consumed per
flying hour) of one RCAF fleet, performed shortly before the release of ESIP requirements, found that
evolutionary changes in role and mission mix had resulted in significant increases in life consumption rate,
on the order of 20 – 30 % per flying hour. This case is representative of the rationale for the ESIP. To
support the RCAF airworthiness authority, NRC-Aerospace has been collecting failure mode data and
developing damage tolerance methods, some of the recommendations have been implemented for the
maintenance and operation of the military gas turbine engines.
Experimental research support from VTT Technical Research Centre of Finland (VTT) has enabled the FDF
to increase the operational life (flight hours) of the high pressure turbine blades of the FINAF F/A-18
engines. The 10% increase in life yielded approx. 3 million USD savings to the taxpayers, with a nearly
hundredfold return of the research investment (ROI).
The RCAF adopted Aircraft Mechanical Systems Integrity Program (MSIP) in 2015, with significant updates
in 2019 in the form of a Technical Airworthiness Authority Advisory which provides further guidance on
MSIP implementation. RCAF fully expects to revise preventive maintenance schedules and tasks based on
operational experience in order to maintain the necessary level of safety. Canada’s submission on Electrical
Wiring Interconnect System (EWIS) monitoring focused on aircraft wiring issues. The regulatory authorities
and industry groups have come to the realization that current maintenance practices may not be adequate to
address aging non-structural systems. While age is not the sole cause of wire degradation, the probability that
inadequate maintenance, contamination, improper repair, or mechanical damage causing degradation to a
particular Electrical Wiring Interconnection System (EWIS) increases over time. Studies have shown a
STO-TR-AVT-275 24 - 5
BEST PRACTICES SUMMARY
tendency to be complacent about EWIS. Any efforts to combat aging wiring problems should be considered to
be a best practice since this is an area that is often ignored but has the potential for catastrophic impacts to an
aircraft.
The Finnish Defence Forces (FDF) is currently investigating implementation of a systematic, comprehensive
approach for mechanical systems integrity management, with intention to avoid duplication of effort
by examining MECSIP or other generally known standard practices. Many activities have been put
into practice to manage integrity of mechanical systems of the F/A-18C/D, the BAE Hawk trainers, and
NH90 helicopters. Some of interesting lessons learned are reported including, Invest on R&D related to data
analysis methodologies as applied to analysis of increasing data on usage, health and environment of the
aircraft and its systems; Make use of modelling and simulation to better understand of functioning and
failures of the critical systems, aiming at the sufficient level of realism.
The German input is the only one in this report on life extension of fuel system and secondary power system
of TORNADO aircraft. Overall the life extension program of TORNADO includes structures, propulsion,
and subsystems, conducted by a tri-national consortium Panavia aircraft GmbH, for a number of NATO
operators including German Air Force, British Royal Air Force, and Italian Air Force). The key safety parts
of the fuel and power systems showed some similarity with structural integrity including mechanical fatigue
and corrosion problems. A harmonized process was developed to implement the life extension program, with
a good cooperation of government and industrial organizations. A lesson learned: for future developments,
potential life extension requirements should be considered already in the design and qualification phase. Life
demonstration for instance should be continued until equipment failure, but not stopped after having
demonstrated the specified life.
24.5 SUMMARY
The best practices implemented by various nations include establishing and implementing integrity
programs for the major aircraft systems: structures, propulsion, mechanical equipment, and subsystems.
The United States developed military standards that provide general and detailed requirements for execution
of each of these integrity programs (and others not addressed in this report) throughout the aircraft’s life
cycle. Many Nations have utilized ASIP MIL-STD-1530 as the basis for establishing structural integrity
requirements for their specific aircraft applications. Due to lack of OEM support and access to design
information, some nations may not be able to implement the Tasks I-III of MIL-STD-1530. Alternatively
these nations often collaborate together to share data and experience, in order to maintain CAW. However,
there appears to be much more diversity in approaches for other safety-critical aircraft systems; but this
could simply be a reflection of the contributions made by each Nation to this report. Other systems integrity
programs have been established using the similar Task structure as in MIL-STD-1530. Compared to
structures, the failure modes of other systems may not be well understood.
In general, the common best practices described throughout this report (and the foundation for all integrity
programs) follows:
1) Collect and evaluate operations and maintenance data.
2) Compare actual usage and maintenance actions with expectations during initial airworthiness
certification through analysis, testing, or assessments.
3) Revise aircraft maintenance requirements based on results of above and continually reassess.
4) Determine when the maintenance program is no longer sufficient to maintain safety to approved levels.
24.6 REFERENCES
[1] Air Force Instruction 62-601, “USAF Airworthiness”, June 2010.
24 - 6 STO-TR-AVT-275
Part 6: ANNEX OF COLLECTION OF CASE STUDIES
ON BEST PRACTICE AND LESSONS LEARNED
STO-TR-AVT-275 Part 6 - i
Part 6 - ii STO-TR-AVT-275
Chapter 25 – FLOW CONTROL SYSTEM INTEGRATION INTO THE
SUBSTITUTE MODELS FOR STRUCTURAL COMPONENTS
LOADS ESTIMATION BASED ON FLIGHT PARAMETERS
AND STATISTICAL INFERENCE METHODS –
MIG-29 VERTICAL FIN USE CASE
25.1 INTRODUCTION
Over 40 years have passed since damage tolerance philosophy was introduced into the design process of
aircrafts [1]. One of the pillars of maintaining structural integrity of aircrafts is to control the loads spectrum
of the structure. In particular, structural load determines the intervals between subsequent Non-Destructive
Inspections (NDI) as well as the safety margins for further aircraft operation. The way that a particular
aircraft is operated after its introduction into service does not necessarily fit its pre-assumed profile. Thus, it
is necessary to monitor the loads of every aircraft in use. The modern approach to meeting this requirement
is to implement OLM systems [2] as a necessary component of aircraft avionics. In the OLM case, aircraft
loads related information could be available from a sensor network e.g., strain gauges or Fiber Bragg
Gratings (FBG) permanently mounted in the aircraft structure and measuring its local strains (direct load
monitoring). However this is not always possible, e.g., in the case of older aircraft, or economically efficient.
In that case, some selected flight parameters, e.g., load factor, are used to determine aircraft load. Due to
complexity and costs of OLM programs with direct load monitoring, indirect load monitoring, based on
flight parameters analysis has lost nothing of its attractiveness. What is available instead of local strain
distribution is a set of recorded flight parameters, which by the laws of inertia and aerodynamics should
determine the dominant part of loads, acting on a given element.
For both approaches to load monitoring, different modes of element deflection need to be established
(Figure 25-1). Depending on the mode, stress distribution at particular critical location can be different,
therefore if different modes can occur, they need to be identified and corresponding fatigue cycles need to be
counted accordingly. In the case of load monitoring based on flight parameters, an additional difficulty
emerges. Beside identification of different stress distribution, functional relations between stress field and
flight parameters also need to be established. One of the methods which can be applied for the task is
Canonical Correlation Analysis (CCA) [3].
STO-TR-AVT-275 25 - 3
FLOW CONTROL SYSTEM INTEGRATION INTO THE SUBSTITUTE
MODELS FOR STRUCTURAL COMPONENTS LOADS ESTIMATION
BASED ON FLIGHT PARAMETERS AND STATISTICAL INFERENCE METHODS
The CCA method is defined as follows. Given two sets of variables, i.e., response variables, e.g., measured
stresses or strains E = {σ 1 , … , σ M } at critical locations and predictor variables, e.g., flight parameters
P = { p1 , …, pN } , one is looking for uncorrelated linear combinations of elements of E and P such that the
correlation between them is subsequently maximized, i.e., the first two canonically conjugated combinations
σ 1c , p1c , satisfy:
(25-1)
whereas the next two combinations are uncorrelated with the first pair:
(25-2)
The CCA method can be illustrated by the example of two modes of simple beam bending. For the two
considered modes, forces are applied at the tip of the beam in its centre along the y-axis and z-axis
respectively, resulting in a pure bending condition (Figure 25-2).
Note: Strain gauges are mounted symmetrically, A and B on top, C and D on bottom side (D is not
visible). Stress is visualized: red is tensile, blue- compressive and green- neutral.
25 - 4 STO-TR-AVT-275
FLOW CONTROL SYSTEM INTEGRATION INTO THE SUBSTITUTE
MODELS FOR STRUCTURAL COMPONENTS LOADS ESTIMATION
BASED ON FLIGHT PARAMETERS AND STATISTICAL INFERENCE METHODS
Theoretical strains along the original x-direction of the beam at selected locations (Figure 25-2) are presented
in Figure 25-3, assuming sinusoidal force-time relations:
• Loading in y-direction results in positive (tensile) strain for strain gauges A and B, and negative
(compressive) for C and D (Figure 25-3(a)); and
• Loading in z-direction results in positive strain for strain gauges A and D, and negative for B and C
(Figure 25-3(b)).
Denoting as F1 , F2 load values of the first and the second modes respectively, the two sets of parameters
upon which CCA method depends can be the following:
(25-3)
For the first case (Figure 25-3(a)), the following condition is satisfied:
(25-4)
(25-5)
Also, for the first type of loads ε 2c = 0 whereas for the second type ε1c = 0 . Therefore the pairs describing
both modes of loads:
(25-6)
can be considered as canonical. The response of any strain gauge ε I at the predefined locations (Figure 25-2)
on any linear combination of the two considered forces F1 , F2 can be represented as a linear combination of
ε1c , ε 2c . When two modes of bending are present in the beam consecutively, readings of strain gauges A and
B are presented in Figure 25-4. The sequence of forces is presented in Figure 25-5 and corresponding
canonical strains are shown in Figure 25-6. Canonical parameters show the mode of bending currently being
present in the beam. Indeed:
STO-TR-AVT-275 25 - 5
FLOW CONTROL SYSTEM INTEGRATION INTO THE SUBSTITUTE
MODELS FOR STRUCTURAL COMPONENTS LOADS ESTIMATION
BASED ON FLIGHT PARAMETERS AND STATISTICAL INFERENCE METHODS
(25-7)
therefore canonically conjugated pairs Equation (25-6) would be obtained as a result of CCA method in that
case. If there is any combination of bending modes present in the load spectrum, it can be captured by CCA,
but in that case both of the canonical strains ε1c , ε 2c would be nonzero.
25 - 6 STO-TR-AVT-275
FLOW CONTROL SYSTEM INTEGRATION INTO THE SUBSTITUTE
MODELS FOR STRUCTURAL COMPONENTS LOADS ESTIMATION
BASED ON FLIGHT PARAMETERS AND STATISTICAL INFERENCE METHODS
The stress distribution can be split into two terms: σˆ inert coming from inertial forces and σˆ aero resulted from
aerodynamic forces:
(25-8)
For vertical stabilizer, the second term was considered as the main contribution to stress distribution. The
following flight parameters were initially considered for modelling of σˆ aero :
The function relating the dependence between flight parameters and stress was assumed to be of the
following form:
(25-9)
STO-TR-AVT-275 25 - 7
FLOW CONTROL SYSTEM INTEGRATION INTO THE SUBSTITUTE
MODELS FOR STRUCTURAL COMPONENTS LOADS ESTIMATION
BASED ON FLIGHT PARAMETERS AND STATISTICAL INFERENCE METHODS
where ρ 0 is the air density at the ground level according to International Standard Atmosphere (ISA) model,
σ nozzle , σ slip , σ roll , σ rudder are dimensionless functions describing contribution to stress distribution due to:
• The Venturi effect, i.e., the aircraft is twin tailed (Figure 25-7) which can be considered as a nozzle;
• The sideslip angle;
• The rate of rotation; and
• The position of the rudder.
Thus the first factor in the Equation (25-9) describes the aerodynamic pressure at the altitude H , and
σ σ σ σ
functions nozzle , slip , roll , rudder should return the stress distribution for different manoeuvres of the
σ σ σ σ
aircraft. The exact form of nozzle , slip , roll , rudder functions was not known, therefore for CCA application
Taylor expansion was used, e.g.,:
(25-10)
with unknown constants cV , cVV , cα , cαα , cV α , etc. All the variables resulting from that expansion,
e.g., paeroV , paeroV 2 , paeroα , paeroα 2 , were considered as independent input variables to CCA model,
i.e., a new set of flight parameters was considered in comparison to the set of initially selected variables.
The CCA model was applied on the data obtained during flight test campaign of two MiG-29 aircrafts. Thus,
in total there were four independent installations of OLM systems (Figure 25-8, Figure 25-9):
• OLM1 – Mounted on the left stabilizer of the first aircraft;
• OLM2 – Mounted on the right stabilizer of the first aircraft;
• OLM3 – Mounted on the left stabilizer of the second aircraft; and
• OLM4 – Mounted on the right stabilizer of the second aircraft.
25 - 8 STO-TR-AVT-275
FLOW CONTROL SYSTEM INTEGRATION INTO THE SUBSTITUTE
MODELS FOR STRUCTURAL COMPONENTS LOADS ESTIMATION
BASED ON FLIGHT PARAMETERS AND STATISTICAL INFERENCE METHODS
The CCA model was trained using only the data obtained by the OLM1 installation and tested on the other
data from strain gauges corresponding to OLM2, OLM3 and OLM4.
A single mode of stress distribution was noticed as a result of CCA. The first canonical flight parameter was
in good correspondence with about 70% of the data obtained from strain gauges in OLM2, OLM3, OLM4
and they had the highest amplitude of the records. The remaining data from the strain gauges had small
variations during flights and were too noisy to be predicted. Examples of the correspondence between
normalized canonical parameter and normalized records of a strain gauge obtained during steady and high
manoeuvre flights are presented in Figure 25-10.
Figure 25-10: Examples of Normalized Strain Records from Strain Gauge (Red)
and Canonical Flight Parameter (Black) for: (a) Steady
and (b) High Manoeuvre Flights.
STO-TR-AVT-275 25 - 9
FLOW CONTROL SYSTEM INTEGRATION INTO THE SUBSTITUTE
MODELS FOR STRUCTURAL COMPONENTS LOADS ESTIMATION
BASED ON FLIGHT PARAMETERS AND STATISTICAL INFERENCE METHODS
The leading contribution to stress distribution conjugated to the canonical parameter ε1c was due to the
σ nozzle term. The nozzle characteristics are dependent on the parameters Equation (25-10):
• V – The velocity of aircraft relative to the atmosphere; and
• α – The attack angle.
Based on that data, several scenarios of flight configurations for Computational Fluid Dynamics (CFD)
computation were prepared. In table below, distribution of occurrences of peaks and valleys of σ nozzle term
with respect to parameters V , α are presented for a certain number of flights. Parameter V ranged between
V1 to V9 and parameter α took values in the range α1 το α31 (true values of V and α are hidden intentionally).
Colours of cells separates between different geometric configurations of the aircraft, e.g., for high angles of
attack, the geometry of the wing changes). Flight configurations Values marked in red in Table 25-1 were
selected for CFD computations. The selected cases were supposed to span subspace of the V – α space
sufficiently broad to capture different distributions of aerodynamic loads and whose occurrence is not rare at
the same time.
Figure 25-11 gives an example of CFD computation for a particular flight configuration and Figure 25-12
gives an example of aerodynamic force distribution on the surface of the stabilizer based on CFD
computations.
Table 25-1: Frequency of Occurrences of Peaks and Valleys of the Nozzle Term.
α \V V1 V2 V3 V4 V5 V6 V7 V8 V9
α1 0 2 1 13 6 4 5 0 0
α2 214 20 66 152 113 48 26 37 5
α3 318 82 232 732 734 1202 3059 2240 205
α4 793 181 1255 4432 13752 24227 27488 15057 911
α5 448 1295 9198 62916 146908 259183 250014 70120 2661
α6 220 9088 147983 576181 1145989 1266522 636115 91436 2986
α7 1295 59789 956659 2956306 3010393 1624642 469213 36944 4092
α8 4511 314354 3527864 4534276 2514661 951120 234361 18216 2712
α9 12784 1076472 3779997 2613920 1150829 429741 110004 7703 1514
α10 30178 1302743 1842749 1050538 463809 202810 52845 3965 727
α11 65712 807516 765251 495699 253180 120762 24739 1771 400
α12 74316 386490 392761 286264 157597 73329 12160 865 157
α13 72827 223026 269199 208723 116088 42178 6274 541 73
α14 50370 128424 190580 154871 79449 22045 3127 280 64
α15 31037 89725 144010 118497 52089 12288 1888 145 13
α16 18421 66638 113588 89235 30763 6718 921 73 2
25 - 10 STO-TR-AVT-275
FLOW CONTROL SYSTEM INTEGRATION INTO THE SUBSTITUTE
MODELS FOR STRUCTURAL COMPONENTS LOADS ESTIMATION
BASED ON FLIGHT PARAMETERS AND STATISTICAL INFERENCE METHODS
α \V V1 V2 V3 V4 V5 V6 V7 V8 V9
α17 11683 53869 98927 63059 17741 3735 442 24 0
α18 8761 48944 86673 43162 11430 2400 240 8 0
α19 6326 36340 56983 23422 6382 1320 117 2 0
α20 4189 25274 32298 12281 3435 751 47 0 0
α21 2852 16523 19248 7376 1964 384 15 0 0
α22 2062 12336 13370 5214 1485 242 1 0 0
α23 1729 11366 11324 4557 1173 161 6 0 0
α24 1636 10605 8959 3591 969 109 6 0 0
α25 1595 9372 7234 2737 786 91 0 0 0
α26 1566 8543 5866 2163 608 58 4 0 0
α27 1226 6777 4737 1454 353 33 0 0 0
α28 966 5211 3455 1010 194 15 1 0 0
α29 629 3190 1908 549 66 7 0 0 0
α30 229 1144 651 112 26 1 0 0 0
α31 59 256 138 45 4 1 0 0 0
STO-TR-AVT-275 25 - 11
FLOW CONTROL SYSTEM INTEGRATION INTO THE SUBSTITUTE
MODELS FOR STRUCTURAL COMPONENTS LOADS ESTIMATION
BASED ON FLIGHT PARAMETERS AND STATISTICAL INFERENCE METHODS
25 - 12 STO-TR-AVT-275
FLOW CONTROL SYSTEM INTEGRATION INTO THE SUBSTITUTE
MODELS FOR STRUCTURAL COMPONENTS LOADS ESTIMATION
BASED ON FLIGHT PARAMETERS AND STATISTICAL INFERENCE METHODS
For comparison using CCA analysis, a single value needs to be estimated, which characterizes the pressure
distribution. Since the mode selected by means of CCA analysis is dominated by the nozzle contribution
σ nozzle , a proper parameter to be compared with the obtained canonical parameter ε1c (Figure 25-10) is the
bending momentum of the stabilizer, i.e., the momentum of aerodynamic forces distribution computed with
respect to the axis of the stabilizer attachment to the frame (Figure 25-13).
In the coordinates frame presented in Figure 25-13, the bending momentum of the stabilizer M g is given by
the formula:
(25-11)
where:
• U denotes the surface of the stabilizer (including rudder surface);
r0
• is a point on the bending axis of the stabilizer (Figure 25-13);
•
e x denotes unit vector along the x-axis;
• × denotes cross product of vectors;
• • denotes the scalar product of vectors;
STO-TR-AVT-275 25 - 13
FLOW CONTROL SYSTEM INTEGRATION INTO THE SUBSTITUTE
MODELS FOR STRUCTURAL COMPONENTS LOADS ESTIMATION
BASED ON FLIGHT PARAMETERS AND STATISTICAL INFERENCE METHODS
Given the momentum M g calculated for every CFD case (Table 25-1), a model of dependence of M g in
the form similar as in Equation (25-9) was considered:
(25-12)
The unknown constants were estimated by fitting it to values of M g calculated by the formula
Equation (25-11). This allowed the interpolation of the obtained CFD results to the entire V – α space under
consideration (Table 25-1) and also to predict values of bending momentum during flight. Examples of the
correspondence between normalized canonical parameter ε1c , normalized records of a strain gauge and
normalized values of bending momentum M g obtained during steady and high manoeuvre flights are
presented in Figure 25-14 There exists a very good correspondence between calculated bending momentum
and the canonical parameter, which confirms the applicability of the CCA method for the determination and
prediction of loads on aircraft structures.
25.3 SUMMARY
This chapter presented Canonical Correlation Analysis (CCA) as a proper method for selection of flight
parameters well suited to predict loads of the aircraft structure. CCA allows both for identification of
different modes of stress distribution as well as identification of flight parameters which are the best suited
for their prediction. This chapter also presented the results of the application of this method to identifying
loads acting on the vertical stabilizer of twin tailed aircraft. The determined canonical predictors were in very
good correspondence with loads values acting on vertical stabilizer which were calculated using
Computational Fluid Dynamics simulations (CFD).
25 - 14 STO-TR-AVT-275
FLOW CONTROL SYSTEM INTEGRATION INTO THE SUBSTITUTE
MODELS FOR STRUCTURAL COMPONENTS LOADS ESTIMATION
BASED ON FLIGHT PARAMETERS AND STATISTICAL INFERENCE METHODS
25.4 REFERENCES
[1] Osgood, C., Fatigue Design – 2nd edition. Pergamon Press, Oxford, 1982.
[2] Boller, C., and Staszewski, W.J., “Aircraft structural health and usage monitoring”, In W.J. Staszewski,
C. Boller, Tomlinson G.R. (Eds.), Health Monitoring of Aerospace Structures, John Wiley and Sons,
Ltd, 2004, pp. 29-73.
[3] Hastie, T., Tibshirani, R., and Friedman, J., The Elements of Statistical Learning: Data Mining,
Inference, and Prediction, Second ed., Springer Science+Business Media, New York, 2009.
STO-TR-AVT-275 25 - 15
FLOW CONTROL SYSTEM INTEGRATION INTO THE SUBSTITUTE
MODELS FOR STRUCTURAL COMPONENTS LOADS ESTIMATION
BASED ON FLIGHT PARAMETERS AND STATISTICAL INFERENCE METHODS
25 - 16 STO-TR-AVT-275
Chapter 26 – SU-22 FSFT STUDY CASE AND
MI-8 HARD LANDING TEST STUDY CASE
26.1.1 Introduction
Poland is one of the last countries in the world that still operates Sukhoi Su-22 Fitter aircraft (Figure 26-1).
Nowadays the Su-22 is a rather obsolete construction and has limited combat value. The Polish Air Force has
decided to extend their service live and keep them operating. The reason for such a decision lies in the lack
of an available fleet that could fulfil auxiliary tasks e.g., anti-aircraft defence training or cooperation with the
Army at the test range.
In 2014, the decision was made to extend the service life of Su-22 aircraft for another 10 years. The Original
Equipment Manufacturer (OEM) had also set the limitation of the total number of landings allowed to
accumulate by each structure during its service life. Operated aircraft, especially the two-seaters version,
have approached the limitation. Taking into account the actual status and planned needs, there was a need to
increase the limit in number of landings as well as a total number of allowed flight hours. It was decided to
implement a research program, whose task was to confirm the possibility of further operation of
Su-22 aircraft and to establish a new limitation.
STO-TR-AVT-275 26 - 1
SU-22 FSFT STUDY CASE AND MI-8 HARD LANDING TEST STUDY CASE
The essential element of the program was the implementation of a Full-Scale Fatigue Test (FSFT) [1]. In
order to determine the adequate loads acting on the aircraft structure during the fatigue test, the spectrum of
operating loads of aircraft used by the Polish Air Force was determined and load measurements were
performed at selected phases of the flight during dedicated research flights.
Based on collected data from on-board recorders, different statistics were calculated. As a result of those
analyses, an average operational profile was obtained. An appropriate load spectrum for FSFT was
developed based on vertical acceleration statistics (Nz values) [2], [3].
A lack of design data had to be overcome by means of extensive reverse engineering activity. The geometry
of the structure had been previously digitalized. Based on this data, computer models were developed, one
series for Finite Element (FE) calculations and second for Computational Fluid Dynamics (CFD) simulation.
The flight tests were carried out to collect data for validation of computer calculations. One aircraft was
equipped with a dedicated flight recorder (ACRA KAM-500) and instrumented with strain gauge sensors.
Before flight tests the measuring system was calibrated. During the calibration, known forces were applied to
the structure and the outputs of the sensors were recorded. Different methods for force application were
adopted, e.g., dedicated sequences of refuelling, adding masses (bombs), and acting forces (wing clamps).
Forces and moments were measured in different sections of the structure during several test flights. Special
attention was paid to landings. Available historical data from recorders could not be used for determination
of loads during landings because of low recording frequency / sampling rates. The numbers of landings with
different gross mass and vertical speed were carried out.
The test article for Full-Scale Fatigue Testing (FSFT) was taken from a monument (Figure 26-2). One of the
withdrawn two-seaters had served as a monument at an Air Base. The operational history of that aircraft was
known. After some preparations performed by Military Aviation Works No. 2 JSC (WZL-2), the aircraft was
delivered to VZLU Prague (Czech Aerospace Research Centre) for FSFT.
Figure 26-2: The Test Rig of the Su-22 UM3K Full-Scale Fatigue Test.
26 - 2 STO-TR-AVT-275
SU-22 FSFT STUDY CASE AND MI-8 HARD LANDING TEST STUDY CASE
The aim of Stage 1 was to prove that the structure can withstand an increased number of landings. The
landing gear itself was not the aim of the test because the Polish Air Forces (PLAF) has a satisfactory
number of spare landing gears at their disposal to ensure that the desired number of landings for the whole
fleet can be achieved.
A mix of different landing scenarios were simulated, followed by NDI checks. The structure passed the
exam.
The aim of Stage 2 was to prove that the structure can withstand more flight hours. Flight hours
are a subsidiary measurement unit. The procedure was based on applying cyclic loading. The
Su-22 aircraft structure has swept wings. This could significantly increase the level of complexity of
FSFT. Based on developed statistics, it turned out that most of missions were performed with a constant
wing swept angle, 45 degrees. A 30 degree angle was used during the landing phase and a 63 degree angle
for supersonic flights. The number of supersonic flights was very low and these flight categories could
be neglected in total fatigue life considerations. Thus the FSFT test rig had to be adjusted for wing
angle configuration changes between Stage 1 and 2. A dedicated NDI program was developed and
executed while Stage 2 of the FSFT was being performed. There were no serious damages or cracks
discovered in Stage 2.
The Stage 3 dealt with the flaps. Forces acting upon the flaps during take-off and landing were calculated
analytically and using a CFD model. Flaps were released and blocked. After the test, no damages were found
by NDT inspectors.
After repeating numerous cycles, the main wing spar broke. This automatically terminated the Su-22 FSFT.
The test article was delivered back to Poland (WZL-2) where the Teardown Inspections (TI) were started.
Teardown Inspections provided the opportunity to inspect areas that were not accessible during FSFT.
Though some new inspection findings appeared, they were not significant enough to change the general
opinion about the durability of the Su-22 structure under the FSFT.
STO-TR-AVT-275 26 - 3
SU-22 FSFT STUDY CASE AND MI-8 HARD LANDING TEST STUDY CASE
The main aim of the project was to further develop the methodologies researched in the preceding
EDA project HECTOR which dealt with devising of a predictive expert system used for detection of cracks
and defects in the airframe [4], [5], [6], [7]. The expert system parameters were devised based on a series of
finite element simulations. In the ASTYANAX project this approach has been extended from a helicopter
panel to the full structure of the Mil Mi-8 Hip helicopter. One of the damage modes for which the predictive
network had to be calibrated was the harsh landing event. An experiment was devised to provide data for this
calibration [8].
Figure 26-3: Mi-8 Hip Helicopter – Gantry Rail Test Set Up with Release Mechanism.
26 - 4 STO-TR-AVT-275
SU-22 FSFT STUDY CASE AND MI-8 HARD LANDING TEST STUDY CASE
26.3.4 Instrumentation
The helicopter was instrumented with an elaborate sensor suite to capture the movement, rotation as well as
structural response of the airframe. The ACRA KAM-500 data recorder has been used for recording of the
main test signals. The measurement was performed with a 15 kHz sampling rate. The sensors used include:
• 54 strain gauge channels in various bridge configurations;
• 4 accelerometers;
• 4 laser distance meters;
• 2 shock absorber length sensors;
• 25 wireless polymer strain gauges;
• 6 MicroStrain® nodes (strain gauges and Micro-Electro-Mechanical-System (MEMS)
accelerometers);
• 4 lines of 190 Fibre Bragg Grating (FBG) sensors; and
• 12 Piezoelectric Transducer (PZT) sensors in three network groups.
Location of the sensors is presented in Figure 26-4. PZTs, MicroStrain® no Nodes, and FBGs used a separate
acquisition system [8].
The FBG sensors were placed in the main areas of concern for the ASTYANAX project, and are crucial in
calibrating the designed sensor system that will be the result of the project. The strain gauge locations were
established based on the operator experience as well as on a preliminary finite element simulation. The
findings from the test will be used to validate a finite element model of the Mi-8 helicopter.
In addition to the on-specimen measurement systems described above, additional external systems were used
to measure geometry before and after each of the drop attempts. Also a fast camera system (1000 frames per
second) were used in the experiment.
STO-TR-AVT-275 26 - 5
SU-22 FSFT STUDY CASE AND MI-8 HARD LANDING TEST STUDY CASE
Two acceleration sensor types were used in the experiment – the piezoelectric, triaxial ENDEVCO 65R-10
accelerometer as well as the classic, soviet produced, uniaxial MP-95 spring-mass g-meter. The MP-95 was
placed in the centre of mass and oriented along the vertical axis of the helicopter, the 65R-10s were installed
in the centre of mass as well as in the front and back sections of the fuselage. The soviet MP-95 is the same
type of sensor that is used in routine service of the Mi-8 – this enables a comparison of flight test data and
the drop experiment signals.
Because of the relatively low structural impact of the 48 cm drop, an additional drop height of 75 cm was
employed. The 0 cm level was defined as the height at which the main gear wheels barely touch the ground.
The fast camera markers used for the velocity measurement are presented in Figure 26-5.
26 - 6 STO-TR-AVT-275
SU-22 FSFT STUDY CASE AND MI-8 HARD LANDING TEST STUDY CASE
Some difficulties arose during the experiment. Firstly, the fluid pressure in the cylinder was low for the 0 cm
level, because the absorber fluid chamber was not fully refilled. Because of that, the vertical displacement
was excessive for the first (0 cm level) drop attempt. For the subsequent drops the fluid was replenished
before each attempt. This resulted in disproportionately higher loads and strains for the 0 cm drop. The
influence of this low cylinder load is clearly seen in the presented signals. Also a problem with the release
mechanism was encountered – for some of the drop attempts the helicopter got stuck in the grips of an
opened release mechanism (due to friction), and the specimen didn’t drop. This was remedied by applying
lubrication to the grip area of the release mechanism, so the helicopter could slide off the grips.
STO-TR-AVT-275 26 - 7
SU-22 FSFT STUDY CASE AND MI-8 HARD LANDING TEST STUDY CASE
The experiment provided an extensive amount of data, enabling a detailed insight into the dynamics of
a harsh landing event [10]. Analysis of the recorded data (strain gauge readings in particular) suggests that
the Mi-8 airframe is resistant to ground impacts for impact velocities below 3.05 m/s – i.e., the assumed
harsh landing threshold. For the highest drop level, only an onset of plastic deformation is observable. The
geometry measurement methods used routinely in Mi-8 helicopter geometry assessment did not record any
significant deformation. This finding is confirmed by the strain gauge readings – the highest recorded strains
were below the yield limit of the material.
More results and deeper analysis of the harsh landing test is accessible in conference proceedings and
Refs. [6], [10], [11].
26 - 8 STO-TR-AVT-275
SU-22 FSFT STUDY CASE AND MI-8 HARD LANDING TEST STUDY CASE
26.4 REFERENCES
[1] Leski, A., Reymer, P., Kurdelski, M., Zieliński, W., and Jankowski, K., (2016). “Full Scale Fatigue
Test of the Su-22 Aircraft – Assumptions, Process and Preliminary Conclusions”, in D. Skibicki (Ed.),
Fatigue Failure and Fracture Mechanics XXVI, pp. 1-7, Melville, New York.
[2] Reymer, P., Kurdelski, M., Leski, A., and Jankowski, K., (2016). The Definition of the Load
Spectrum for Su-22 Fighter-Bomber Full Scale Fatigue Test”, Fatigue of Aircraft Structures, 1(2015),
pp. 28-33. doi:10.1515/fas-2015-0005.
[3] Reymer, P., Kurdelski, M., Leski, A., Leśniczak, A., and Dziendzikowski, M., (2018). “Introduction
of an Individual Aircraft Tracking Program for the Polish Su-22”, Fatigue of Aircraft Structures,
2017(9), pp. 101-108. doi:10.1515/fas-2017-0008.
[4] Giglio, M., Manes, A., Mariani, U., Molinaro, R., and Matta, W., (2009). “Helicopter Fuselage Crack
Monitoring and Prognosis through On-Board Sensor Network”, Proc. of CM 2009 and MFPT 2009
(The Sixth International Conference on Condition Monitoring and Machinery Failure Prevention
Technologies), Stillorgan Park Hotel, Dublin, Ireland, 23–25 June 2009.
[5] Vallone, G., Sbarufatti, C., Manes, A., and Giglio, M., (2013). “Artificial Neural Networks for
Structural Health Monitoring of Helicopter Harsh Landings”, Applied Mechanics and Materials, 390,
pp. 192-197.
[6] Leski, A., Kurdelski, M., and Stefaniuk, M., (2015). “Investigation of a Helicopter Harsh Landing
Based on Signals from installed Sensors”, in Proceedings of the Ninth Australian Defence Science and
Technology Organisation (DSTO) International Conference on Health and Usage Monitoring
Systems, A. Sinha, Ed., Defence Science and Technology Organisation, Melbourne, Australia, 2015.
[7] Sbarufatti, C., Manes, A., Giglio, M., Mariani, U., Molinaro, R., Matta, W., Di Luzio, I., Toscani, D.,
Archetti, F., Bjerkan, L., Hudec, R., Wieser, V., Makys, P., Gozdecki, J.and Katsikeros, C. (2010)
Application of Structural Health Monitoring over a Critical Helicopter Fuselage Component.
http://hdl.handle.net/11311/574961.
[8] Giglio M. et al., AHS Forum Proceedings CD 71 (2015). Available at: http://vtol.org/store/product/
modelbased-structural-integrity-assessment-of-helicopter-fuselage-during-harsh-landing-10147.cfm.
[9] Zimmerman, R.E., Warrick, J.C., Lane, A.D., Merritt, N.A., and Bolukbasi, A.O., (1989). Aircraft
Crash Survival Design Guide. Volume 3. Aircraft Structural Crash Resistance, Simula Inc. (1989),
ADA218 436, USA AVSCOM TR 89-D-22.
[10] Dziendzikowski, M., Dragan, K., Kurnyta, A., Klimaszewski, S., Leski, A., and Vallone, G., (2015).
“Fatigue Cracks Detection and their Growth Monitoring During Fatigue Test of a Helicopter Tail
Boom”, in F.-K. Chang and F. Kopsaftopoulos (Eds.), Structural Health Monitoring, DEStech
Publications, Inc.
[11] Sbarufatti, C., Vallone, G., Giglio, M., Leski, A., Stefaniuk, M., and Zieliński, W., (2016).
“Experimental Validation of a Computational Hybrid Methodology to Estimate Fuselage Damage Due
to Harsh Landing”, Journal of the American Helicopter Society, 61(4), pp. 1-11.
STO-TR-AVT-275 26 - 9
SU-22 FSFT STUDY CASE AND MI-8 HARD LANDING TEST STUDY CASE
26 - 10 STO-TR-AVT-275
Chapter 27 – MIG-29 AND F-16 STUDY CASES
Krzysztof Dragan
Air Force Institute of Technology
POLAND
27.1 INTRODUCTION
The MiG-29 is the basic fighter jet used in Polish Air Force. Two main criteria in the maintenance
approach of MiG-29 are based on Hour Service Life consumption (called HSL) and service life based on
aircraft age (so-called Calendar Service Life – CSL). AFIT played a role in the introduction of new
maintenance approach named On Condition Maintenance (OCM) in connection with the possibility of
extending service life of some airplanes. One of the primary tasks was NDT (Non-Destructive Testing)
work for critical components, as well as a corrosion inspection program for MiG-29 aircraft. Activities
connected with the NDT and corrosion inspection program are also delivered for other aging aircraft of
Polish Armed Forces. In this review, only part of the NDT work will be presented and related to the
composite skin of vertical tail of MiG-29. The MiG-29 is an aircraft with double vertical tail construction.
The vertical tail consists of aluminum substructure and composite skin made of CFRP (Carbon Fiber
Reinforced Plastic). The skin is monolithic, with a thickness range of 1.8 mm – 3.8 mm and it is joined to
the substructure using rivets and adhesive [1].
Techniques which are used for the inspection of MiG-29 vertical stabilizers are also used for F-16 horizontal
and vertical stabilizers are presented in Table 27-1. In Figure 27-1, an automated C-Scan system used for the
inspection is shown.
Table 27-1: Information about NDT Techniques Used and Their Capabilities.
STO-TR-AVT-275 27 - 1
MIG-29 AND F-16 STUDY CASES
As can be observed from Table 27-1, only multimode NDT damage detection is efficient enough to cover
possible damage detection [2], [3]. Due to the different failure mode occurrence in the inspected structure,
such an approach covers the possibility of damage detection and proper description. For implemented OCM
program, detailed NDT with the use of ultrasonic and low frequency acoustics is being under used at present.
To perform structural integrity testing of the composite vertical stabilizer of a MiG-29 jet, ultrasonic testing is
applied. The total surface area for inspection is approximately 11 sq. meters per aircraft. The aircraft
population, which is currently being tested, is more than 30. Thus the total area for inspection is more than
600 sq. m for the whole monitored aircraft population. Classical NDT tests using manual hand scanning are
very difficult or impossible, because the work has to be done on a very large area. An automated scanning
procedure was therefore applied. The MAUS®V system (Mobile AUtomated System) that was used is a hybrid
construction and makes possible inspections with such techniques as:
• Ultrasound, eddy current;
• Eddy current;
• Mechanical Impedance Analysis (MIA);
• Pitch-catch;
• Resonance; and
• Phased Array.
The system is fully portable and enables inspection on horizontal and vertical inverted surfaces due to flexible
track system equipped with special fixturing. Figure 27-2 shows a MAUS®V system installed for an inspection
of the vertical stabilizer composite skin of a MiG-29. Low frequency acoustics and ultrasonic inspection with
MAUS system are quantitative methods. The results are presented in so-called C-scan mode, which makes
damage location and sizing possible. The system’s embedded interface enables amplitude and TOF (Time Of
Flight) C-Scan data presentation. In Figure 27-2, the Time of Flight C-scan of a MiG-29 vertical stabilizer
is presented.
27 - 2 STO-TR-AVT-275
MIG-29 AND F-16 STUDY CASES
The colour range in the TOF data provided gives information about different composite skin thicknesses as
well as the depth of the damage location. This type of data presentation is very helpful for damage
description, as well as for fast determination of the damage location.
In the amplitude data presented in Figure 27-3(a), a different grey scale intensity is indicative of different
signal attenuation. The amplitude mode is useful for determining damage location. The distribution of the
damage locations relative to the side of the vertical tail (internal and external) as well as left or right tail has
been found to have an approximately normal distribution. Some interesting inspection results are presented
in the Figure 27-4 and Figure 27-5.
STO-TR-AVT-275 27 - 3
MIG-29 AND F-16 STUDY CASES
If more information about damage characterization is required, additional full waveform capturing is
performed, based on B-scan visualization (presented on the Figure 27-4). These results are used for the
detection of delaminations and foreign object inclusions. A typical distribution of damages is in the form of
disbonds, located at the bottom of the stabilizer, shown in Figure 27-5. As previously mentioned, monitoring
of all operating aircraft contained in AFIT designed service bulletins has been implemented.
27 - 4 STO-TR-AVT-275
MIG-29 AND F-16 STUDY CASES
Figure 27-7 presents data from the ultrasonic inspection of the horizontal stabilizer (top side). Based on
detailed information of structure design and thickness changes, failure mode occurrence can be analyzed.
This approach to the first inspection of the composite elements provides holistic information about the initial
structure condition. From the manufacturing and maintenance point-of-view, various failure modes may
occur, thus profound knowledge about structural integrity starting from aircraft manufacture and aircraft
delivery is important. Figure 27-8 presents an example of a delamination around a rivet, which occurred
during original structure manufacturing. From the whole aircraft fleet population, damage distribution
information was collected and stored in a database.
STO-TR-AVT-275 27 - 5
MIG-29 AND F-16 STUDY CASES
Figure 27-8: Amplitude and TOF C-Scan F-16 Horizontal Stabilizer with Damage Presence.
27.4 CONCLUSIONS
All the information presented here is related to the automated and advanced NDI of composite aerospace
structures. All these activities are connected with NDI procedures and service bulletins accepted by the
appropriate Technical Airworthiness Authority in the Polish Air Force. MiG-29s are inspected as part of
scheduled maintenance for the purpose of structural integrity monitoring, which is part of on condition
maintenance. Structures on F-16s were inspected for collecting data related the initial structure integrity of
composite structures which will be used for further perspective support of aircraft maintenance.
27.5 REFERENCES
[1] Dragan, K., and Klimaszewski, S., (2006), “In-Service Flaw Detection and Quantification on the MiG-
29 Composite Vertical Tail Skin”, 9th ECNDT Berlin, (Germany) 25 – 29 September 2006.
[2] Dragan, K., Klimaszewski, S., Sałaciński, M., Synaszko, P., and Latoszek, A., (2010), “Damage
Detection and Failure Mode Distribution For The Mig-29 Vertical Stabilizer”, 2010 Aircraft
Airworthiness and Sustainment Conference, 11 – 13 May, Austin Convention Center.
[3] Sałaciński, M., Synaszko, P., Stefaniuk, M., and Dragan, K., (2011), “Monitoring of Crack Growth in a
Structure Under a Composite Patch”, Journal of Fatigue of Aircraft Structures, 1(3) August 2011,
pp. 103-111, ISSN 2081-7738.
27 - 6 STO-TR-AVT-275
Chapter 28 – NDT ACTIVITIES DURING COMPOSITE BONDED
REPAIR OF PL C-130E WING BOTTOM SKIN
Piotr Synaszko
Air Force Institute of Technology
POLAND
28.1 BACKGROUND
28.1.1 General
Continued airworthiness is required to accomplish aviation safety regulations. Non-destructive inspections,
according to regulations (CS 23.573.a) are mandatory to ensure structural integrity and safety for composite
structures. Such inspections also concern composite bonded repair. During the realization of Programmed
Depot Maintenance (PDM) on C-130 owned by Polish Air Force, Military Aviation Works No. 2 identified
composite bonded repair on the bottom side of wing. Based on the documented requirements, a composite
patch should be inspected using infrared thermography. The available documentation consists of a procedure
for thermography inspection and technical documentation for manufacturing of specimens. Documentation
of the repair itself was not available.
STO-TR-AVT-275 28 - 1
NDT ACTIVITIES DURING COMPOSITE
BONDED REPAIR OF PL C-130E WING BOTTOM SKIN
28.2 WORKFLOW
Unacceptable
Damage was NDT Report
detected
NDT Report
28 - 2 STO-TR-AVT-275
NDT ACTIVITIES DURING COMPOSITE
BONDED REPAIR OF PL C-130E WING BOTTOM SKIN
STO-TR-AVT-275 28 - 3
NDT ACTIVITIES DURING COMPOSITE
BONDED REPAIR OF PL C-130E WING BOTTOM SKIN
Figure 28-6: Flaw Indication (Left) and Temperature Profile (Right) Plotted for Blue Line.
The damage found corresponded to the thermography indications. There was evidence of lack of bonding on
two fragments of the wing skin, and voids between layers of the adhesive film.
28 - 4 STO-TR-AVT-275
NDT ACTIVITIES DURING COMPOSITE
BONDED REPAIR OF PL C-130E WING BOTTOM SKIN
STO-TR-AVT-275 28 - 5
NDT ACTIVITIES DURING COMPOSITE
BONDED REPAIR OF PL C-130E WING BOTTOM SKIN
28.3 SUMMARY
This study identified significant manufacturing defects in repair packages, indicative of the poor quality of
previously developed or implemented repair technology. Based upon these inspections results, AFIT
developed their own procedure for repair as well as for NDT.
In this work, the authors used the FLIR SC7000 camera, which was demonstrated to have sufficient
sensitivity to determine the integrity of the repair patch. This procedure was used primarily to determine the
quality of the repair patch after prefabrication, prior to bonding of the repair patch to the structure.
In the AFIT developed repair procedure, attention was paid to the bonding of the repair patch to the wing
skin structure. The NDT inspection procedure has been extended to include the MOI procedure for wing skin
cracks detection.
28.4 REFERENCES
[1] Verhoeven, S., “Bonded Repair of Corrosion Grind-Outs”, (2005), USAFA TR 2005-6.
[2] Günther, G., and Maier, A., (2010), “Composite Repair for Metallic Aircraft Structures Development
and Qualification Aspects”, ICAS 2010.
[3] Davis, M.J., “The Development of an Engineering Standard for Composite Repairs”, (1994), Seville,
AGARD.
[4] Baker, A., and Chester, R., (2003), “Bonded Repair Technology for Aging Aircraft”, RTO Meeting
Proceedings 79(II).
28 - 6 STO-TR-AVT-275
Chapter 30 – HELICOPTER BLADES FAILURES MODES
Krzysztof Dragan
Air Force Institute of Technology
POLAND
The above mentioned criteria are also used to determine the MRB service life. The contribution to MRB life
from actual operations (HSL) is in the range of 30 – 40 % of the service life. This means that theoretically
MRB even in a good condition must be replaced because of aging (CSL). The introduction of Cost Benefit
Analysis was necessary for the AFIT/ITWL (Air Force Institute of Technology) to realize an MRB service
life extension program.
There are types of MRB blade constructions used in the Polish Air Force, Navy, Army: one with single spar
and multi-section honeycomb, the other with single spar and non-section honeycomb structure. Table 30-1
presents information about the main rotor blades used in Polish Armed Forces helicopters. Most of the
blades used in the Polish Armed Forces are metal with multi-section structure. Their differences consist of
size, spar manufacturer and number of blades (Qty), as well as HSL and CSL [1], [2].
Note: 1Spar manufactured in Poland; 2After detailed servicing; 3Operating in sea environment.
STO-TR-AVT-275 30 - 1
HELICOPTER BLADES FAILURES MODES
Detailed analysis of the operating intensity of the helicopters proved that the percentage use of HSL of
helicopters is negligible comparing to assumed level. Figure 30-1 presents the percentage share of the use of
HSL of selected helicopters used in the Polish Armed Forces [2].
The percentage share of HSL use is taken into consideration at the end of the CSL. It means that it is not
economical to replace MRB with new sets when the HSL percentage use is only 17% – 57%. However, the
existing maintenance procedures are not able to take into account the technical condition of aged main rotor
blades. For this, it was necessary to perform exact statistical data analysis which should include such points
(amongst others) as:
• Percentage share of failure mode;
• Percentage share of failure effect;
• Percentage share of MRB elements;
• Number of failures caused by selected mode;
• Percentage share of failures caused by aging process;
• Safety and reliability analysis; and
• Possibility of using cluster analysis.
These analyses give necessary information for further work. In the AFIT there is a database system (SAN)
which stores the data about failures, providing the possibility of various analyses. Figure 30-2 presents
information about percentage share of various failure modes expected during the service life of MRB.
30 - 2 STO-TR-AVT-275
HELICOPTER BLADES FAILURES MODES
The most frequent failure was skin separation, as well as disbond. A much lower contribution was due to
other failures. Failure was determined by several factors. The most important factors were aging and
manufacturer fault. The next one was in-service damage with only 11% share of failure effect (caused by low
level of HSL). During further analysis, the most critical elements were determined, as well as the number of
events, and cluster analyses were performed. These analyses provided a good base for determining what
method should be used and which MRB should be inspected.
Maintenance procedures used so far, not only in the Polish Armed Forces, were based on visual inspection,
with the additional support of hammer (coin) tap testing. Those techniques are fast and fairly simple to use
but are not reliable and are operator dependent. The next important issue is the fact that these procedures are
also very dependent on the so-called “human factor”. Taking that into consideration, as well as analyzed data
from failure analysis, appropriate NDE work was performed. Selected enhanced NDI techniques, such as:
Shearography; MIA; Pitch-Catch; Ultrasonic; Ultrasonic Phased Array; Eddy Current; and X-Ray, were
applied [2], [3]. These techniques were supported with a PC interface, which made post process data analysis
on site. The other advantages of having a PC based interface were: data storing and data comparison;
automation of arm scanning procedures; and enhanced visualization. The first step in the NDE work was to
prepare special specimens using out of service MRB, as well as those still used in service. Using these, it was
possible to create damages similar to the those found in service for comparison and to determine detectability
limits. The next issue was to compare these with serviceable MRB in a different condition, and different
CSL or HSL. One of the tasks was also to perform low energy impact tests on selected MRB from W-3
helicopters and establish damage detection. Some of the results of these inspections are presented below.
Based on extensive R&D programs, special test plans were stablished. A specialized specimen was designed,
and a test scenario for damage detection was implemented.
The first step of that work was to determine damages which could be detected and which happen during
the service life of MRB of helicopters used in the Polish Armed Forces. Those damages could be
described as follows:
• Disbonds (skin to honeycomb, skin to spar);
• Cracks (in the spar);
• Corrosion; and
• Water ingress (honeycomb cells).
One of the important issues connected with the maintenance of the metal MRB is crack growth in the spar
structure of the blade. Detailed analysis of the failure modes made it possible to determine so-called
“hot spots”, i.e., spar areas where the cracks occur the most often.
It is possible to divide the spar structure into three areas – Figure 30-3 A, B, C. Areas A (leading edge) and
B consist of Top Wall (TW) and Bottom Wall (BW). Area C is a location where all the sections are bonded to
the spar, and where the Rear Wall (RW) is located. Detection of the cracks in Areas A and B is easier because
of the geometry. Inspection is difficult on the blade leading edge, where an anti-erosion layer is installed.
In the several specimens made from MRB blades, cracks were created by fatigue cycling in a load frame,
which provided an opportunity to determine which inspection techniques were most efficient at crack
detection. As a result, eddy current and ultrasonic inspections were chosen for the study.
Eddy currents are a useful tool for the inspection of aluminium alloys. The major limitation of this technique
is the so-called depth of penetration, which depends on conductivity and the frequency of excitation of eddy
currents in the material being inspected.
STO-TR-AVT-275 30 - 3
HELICOPTER BLADES FAILURES MODES
The thickness of the spar in most locations is in the range of 3.3 mm. At the airscrew, the thickness increases
to 5.6 mm. Based on conductivity measurements, the estimated depth of penetration was determined to be
equal to 3 mm. This means that most of the TW and BW is inspectable. Figure 30-5 presents a C-scan result
with a visible crack, which was grown by fatigue from the notch shown in Figure 30-4.
30 - 4 STO-TR-AVT-275
HELICOPTER BLADES FAILURES MODES
Another issue, which is important for eddy current inspection, is density. The density of eddy currents
decreases with increasing depth of penetration. Such phenomena should be considered as the inspection is
performed on deeper crack locations.
Because of the large inspection area, some applications should be used to increase inspection speed, while
maintaining reliability. For that reason, automated data collection with the use of eddy currents as well as
Magneto Optic Eddy Current Imager (MOI) were selected. Such inspections enable fast and reliable data
collection and data storing. The MOI inspection is easy and fast because the use of visualization is based on
the “Faraday effect” to polarized light beam. Figure 30-6 shows an example an MOI inspection.
It is necessary to mention that a large area of the MRB leading edge is protected with a non-conductive layer,
which makes crack inspection using conventional eddy current difficult. Moreover, the use of the ultrasonics
in that area is not possible because of large signal attenuation caused by the rubber layer. For that
application, some tests will be conducted using Remote Field Eddy Current, which enables deeper
penetration, especially through the non-conductive layer.
For the further evaluation of the spar structure, ultrasonic and X-Ray tests were carried out. Ultrasonics is a
relevant inspection technique, which enables detailed defect characterization. However, because of extreme
sensitivity for some selected failure modes, strict conditions should be fulfilled, such as: signal to noise ratio,
frequency selection, and coupling assurance. For the blade inspection, shear waves for crack detection were
considered to be the appropriate technique. Other tests were made using X-Ray. The X-Ray technique
enables inspection of difficult to access areas, and data can be stored and compared as images. Tests were
made to assess crack detectability in the spar structure, especially under the leading edge protection layer.
Figure 30-7 presents data from X-Ray inspection made with the energy range 60 – 80 kV according to
EN 584-1. On that figure, there is a visible crack going from the hole in the BW. Results were not as expected
(especially on the leading edge). Cracks were not clearly visible and data interpretation strongly depended on
the experience of the data analyzer. For that particular application, eddy currents and ultrasonic are more
efficient. Another application is pitting corrosion detection. For that application, ultrasonic is efficient. Tests
were made with appropriate specimens which had drilled holes with different sizes and depths.
Another important issue is the detection of disbonds leading to skin separation (and section separation)
between the section to spar, and skin to honeycomb disbonds. Taking into the consideration that some of the
damages are not allowed (or size is specifically limited), determination of detection possibility is required.
STO-TR-AVT-275 30 - 5
HELICOPTER BLADES FAILURES MODES
For that reason, special specimens were designed and made from the blade elements. Figure 30-8 presents
disbonds which were made for the different types of blades. For that specimen, shearography, Mechanical
Impedance Analysis (MIA) and resonance testing were applied.
Figure 30-9 presents results of blade inspection with the use of MIA. For that damage, it is possible to detect
skin to honeycomb disbonds and section separation both with the use of MIA and shearography. Low
resolution results were obtained using resonance testing.
30 - 6 STO-TR-AVT-275
HELICOPTER BLADES FAILURES MODES
For most of the highlighted scenarios, several tests successfully detected and described the damage. Based
on the results of this three year program, several service inspection bulletins have been elaborated and
introduced to service life extension programs.
Figure 30-10 presents the approach used for in-service NDI inspection of the Mi-8 helicopter main
rotor blades.
30.2 CONCLUSIONS
Based on developed NDI bulletins, service life extension programs have been delivered on the selected main
and tail blades for different types of helicopters, which have enabled further helicopter operation. Moreover
all the NDI technologies elaborated on here are currently used as part of on condition based maintenance
procedures. Such procedures are also part of airworthiness policy which is in accordance with planned
implementation of European Military Airworthiness Regulation to Polish MOD.
30.3 REFERENCES
[1] Dragan, K., and Klimaszewski, S., (2008), “Holistic Approach for Structural Integrity Evaluation of
Composite Main Rotor Blades”, WCNDT 2008, Shanghai, China.
[2] Dragan K., (2009), “NDE Activities Connected With Service Life Extension of Main Rotor Blades of
Helicopters used in Polish Armed Forces”, 7th Australian Pacific Vertiflite Conference on Helicopter
Technology, 9 – 12 March 2009, Melbourne.
[3] Dragan, K., Klimaszewski, S., Kudela P., Malinowski P, and Wandowski T., (2010), “Health
Monitoring of the Helicopter Main Rotor Blades With the Structure Integrated Sensors”, Proceedings
of the Fifth European Workshop on Structural Health Monitoring, 2010, p. 66-69.
STO-TR-AVT-275 30 - 7
HELICOPTER BLADES FAILURES MODES
30 - 8 STO-TR-AVT-275
Chapter 31 – METALLIC HELICOPTER BLADES
COMPOSITE REPAIRS STUDY CASE
Michal Sałaciński
Air Force Institute of Technology
POLAND
31.1 INTRODUCTION
The Polish Armed Forces are currently operating hundreds of the Mi family of helicopters. The helicopter
metal fuselage is usually resistant to battle and the human factor conditions. Unfortunately, the metal rotor
blades of Mi helicopters are sensitive to operating conditions. Each blade is made with a monolithic
aluminium spar and mutually separated trailing sections, which are bonded to the spar. The sections are
constructed of metal sandwich panels.
During aggressive military operating conditions, blades sections are often damaged by: debonding from the
spar, fatigue cracks forming in sections of the skin, dents and perforations, as well as erosion. The
manufacturer assumed that structurally damaged sections should be replaced, provided that repair
technologies are applicable only to cosmetic damages. Unfortunately, there is a limit to the number of
repairs, which prevents replacement of two neighboring sections due to the high temperature curing cycle
experienced during the section replacement. Additionally, the old technology repairs are expensive and
time-consuming. Therefore, it has been necessary to develop new repair technologies to enable the repair of
rotor blade structural damages. This chapter focuses on discussing important issues in the field of
repair technology.
STO-TR-AVT-275 31 - 1
METALLIC HELICOPTER BLADES COMPOSITE REPAIRS STUDY CASE
Rotor blades are elements that belong to Category 1 – primary structures because their integrity determines
helicopter integrity and therefore crew safety [1]. The rotor blades are the most loaded aircraft elements,
because they are operated in complex load conditions (Figure 31-2).
Figure 31-2: Scheme of Helicopter Rotor Blade Loads: (a) Tensile Force –
Centrifugal P1, Bending Momentum of the Blade in the Direction of
Flow Mg1; (b) Carrier Force Pz; Aerodynamic Momentum Mg3;
(c) Loads Related to the Change of Blade Position
During the Flight; (d) Vibrations.
The blades are loaded during operation and during maintenance as well. It should be underlined that the
sections are made in a sandwich structure, so they have a very low resistance to point loads, such as those
that occur from impact. This situation can occur during take-off and landing, where elements are picked up
from the ground or during maintenance, i.e., when tools are mistakenly dropped on the blade.
Rotating elements are exposed to erosion. It can be assumed that a helicopter is on the ground with running
engines for about 20% of its service life. This time is related the start-up and periodic operations. The
average contamination of solid particles in air at standstill is 40 mg/m3 – this is the value of air density
during the strongest sandstorms (for example in Afghanistan) experienced during desert take-off. The rest of
the operating time with running engines takes place in the air, where the average contamination in air with
solid particles is 0.3mg / m3 [2]. The rotor blade safe life of the Mi-2 helicopter is 2000 hours. During service
operations, the components on the leading edge and on the trailing sections can be thinned or even
completely eroded through.
Each helicopter flight can experience variable atmospheric conditions. The variable atmospheric conditions
may be result of changes in flight altitude or type of mission. In addition, helicopters in the Mi family are
intended for inland and sea operations. The adhesive joints and other polymeric materials, rubber, sealants
and others present on the blade are exposed to variable temperature and high humidity.
The aviation regulations for the certification of heavy helicopters [3] do not explicitly define the maximum
operating temperature of rotor blade. Subsection CS29.1043 indicates the minimum temperature of lubricant
tests, which is -55.5 °C. The data contained in [4] regarding the Mi-38 heavy helicopter, indicates that the
helicopter meets the European requirements of JAR-29 in terms of operating conditions. According to the
31 - 2 STO-TR-AVT-275
METALLIC HELICOPTER BLADES COMPOSITE REPAIRS STUDY CASE
authors of Ref. [4], the regulations require that the operating temperature should be within the range of
60 ÷ 50 °C. The aviation regulations regarding the certification of heavy helicopters [3] in subsection CS29.45
indicate that the maximum operating humidity is 80%. Unfortunately, in sea conditions, humidity reaches
100%, and in addition, the atmosphere contains a mist of saltwater, which is an electrolyte conducive to
galvanic corrosion.
All the above factors cause the rotor blades to be subject to emergency failures before the end of service life
resulting from the safe life. The repair technologies provided by the manufacturer, unfortunately, do not
allow the restoration of the properties to pre-damage levels. Therefore, AFIT as an institution supporting the
Polish Air Force took up the development of new modern technology in the repair rotor blades.
Figure 31-3 shows typical emergency damage to blades in the Mi family of helicopters.
Figure 31-3: Damage Types in Mi Metal Rotor Blades: (a) Disbonding a Piece of Stainless
Shield on the Leading Edge; (b) Disbonding Between Skin Trailing Section and
Spar; (c) Cracks a Skin’s Section; (d) Dent; (e) Puncture in a Skin’s Sections.
In accordance with the applicable regulations in the Polish Air Force [5], it is unacceptable to operate rotor
blade with punctured a skin’s sections, cracks a skin’s section and disbonding a piece of stainless shield on
the leading edge. However, depending on the damage location and damage size it can be to use blades with
the dent damage and disbonding between skin trailing section and spar.
STO-TR-AVT-275 31 - 3
METALLIC HELICOPTER BLADES COMPOSITE REPAIRS STUDY CASE
The repair consists in removing the original stainless shield and rubber in the area of the leading edge fittings
and replacing the removed components with new ones with better properties (Figure 31-5).
Figure 31-4: Strength Testing of the Bonding Join on the Leading Edge: a) Test Results Before
and After Repair; b) Character of Loading; c) the Failure Type of Samples after Tests.
Figure 31-5: Repair of the Leading Edge: (a) Scheme of the Repair, (b) Scheme
of Assembly Station, (c) the Leading Edge After Repair.
For repair, a rubber with lower stiffness was selected and additionally subjected to etching with sulfuric acid.
Etching with sulfuric acid increases the adhesion properties of the rubber and also makes it more elastic.
The metal hardware was replaced with a thinner stainless-steel sheet than the original one. A thinner metal
sheet is used for a shield because it is more flexible, which allows it to be adapted to the variable geometry
of the blade during operation. Before being bonded, the shield surface was prepared by chemical treatment
and the use of primer hardened at 400 °C. High temperature causes adequate primer hardening
and allows diffusion across the metal. The use of a suitable adhesive paste mix allows for a durable and
flexible connection.
31 - 4 STO-TR-AVT-275
METALLIC HELICOPTER BLADES COMPOSITE REPAIRS STUDY CASE
Although the repair seems relatively simple, the technological process is composed of many complicated
stages. These include etching with sulfuric acid, preparation of the metal surface for bonding and bonding
rubber for the heating package and shield for rubber.
The repair technology of disbonding between skin trailing section and spar provided by the manufacturer
consists in re-bonding the skin to spar by injecting a liquid epoxy resin into the gap between skin and spar.
This method is in fact not effective, because it is not possible to properly clean the surfaces between
bonding. To properly prepare the surface would be to bend out the skin. Unfortunately it is unacceptable,
because it would cause permanent plastic deformation of the skin at the bending point.
During the research at AFIT a method of repair was developed consisting in removing the debonding skin
and reconstructing the structure from the equivalent material (Figure 31-6). This method enabled proper
cleaning of the bonding surfaces, preparation of the surfaces for bonding and protection against corrosion.
Figure 31-6: Repair Technology of Disbonding Between Skin Trailing Section and Spar.
The calculations and tests have shown that the patch carried the load no worse than the original skin [6].
The technology repair is the same for cracks in a skin’s section as well as for punctures in a skin’s sections.
The repair consists in removing the paint coat in the area of bonding, drying the structure (Figure 31-10) and
then bonding the patch (Figure 31-7).
The most difficult process from among those presented is the dent repair. The repair consists in removing
damaged components of section and reconstruction it using new ones with equivalent properties (Figure 31-8).
Restoring a structure is a two-step process. First, the replacement core with the technological excess
protruding above the surface of the skin is bonded. After hardening of the adhesive films, the core is mashed
even with the skin surface. Then an aluminium alloy patch is bonded.
STO-TR-AVT-275 31 - 5
METALLIC HELICOPTER BLADES COMPOSITE REPAIRS STUDY CASE
Figure 31-7: Repair Technology of Cracks in a Skin’s Section and Puncture of a Skin’s Sections.
31 - 6 STO-TR-AVT-275
METALLIC HELICOPTER BLADES COMPOSITE REPAIRS STUDY CASE
The vacuum bag film should have a few small holes so that air circulates in the vacuum bag to allow the
suction of moisture. Therefore, it is important that the process was carried out in a room where the air is
clean and dry.
The boiling point of the liquid depends on its pressure. The diagram in Figure 31-11 can be useful to set the
temperature of the heating blanket depending on the value of the vacuum existing in the vacuum bag.
STO-TR-AVT-275 31 - 7
METALLIC HELICOPTER BLADES COMPOSITE REPAIRS STUDY CASE
Figure 31-12: Surface Preparation for Bonding and Type of Failure During Wedge Test
(ASTM D3762) of the Joint Subjected to 100% Humidity Exposure: a) Mechanical
Treatment – Adhesive Damage, b) Mechanical Treatment and the Use
of Sol-Gel Treatment – Cohesive Damage.
The chemical treatment of the surface before bonding with the use of Sol-Gel allows a chemical joint at the
metal-resin interface. The chemical joint limits the destructive effect of water (Figure 31-12b).
31 - 8 STO-TR-AVT-275
METALLIC HELICOPTER BLADES COMPOSITE REPAIRS STUDY CASE
During composites production, stationary devices such as an oven or autoclave are used to realize the
thermal cycle, but for repairs, a mobile device is used (Figure 31-14).
Figure 31-14: Schema of Device and Vacuum Bag During Cure Cycle of Patch.
31.6 SUMMARY
This paper presents a method for carrying out repairs on metal rotor blades in the Mi family of helicopters.
Some of the repairs presented have already been put into operation in the Polish Air Force. Some repairs will
be implemented in the near future. For all repair methods, necessary tests on the ground in accordance with
the developed test program have been carried out. Preparations are underway for the last stage of the
research – experimental flights.
Details of the research program, calculations and the manner of carrying out ground-based tests together with
the results are described in Ref. [6].
The most important benefit of the new technologies developed is the financial saving. No less important is
the way that the developed technology allows us to become less dependent on materials from the Russian
market. Materials that can be used for the new technology are available primarily in the EU and the USA. In
addition, repairs may be carried out by a trained staff of the Polish Air Force. Some repairs are possible both
in the Military Works and in field bases during the implementation of military missions.
In summary, the developed new technologies significantly increase the possibility of maintaining a
continuous airworthiness of the blades.
31.7 REFERENCES
[1] Greenwell, T.A., “Design of Repair of Battle-Damaged Fixed-Wing Aircraft” (2010), USAF Academy,
Colorado, OMB No. 0704-0188.
[2] Younhong, H., Peiyuan, H. and Anli, W., (2012), “Research on Erosion Mechanical Parameters of
Wind-Sand Environment in the Central and Western Region of inner Mongolia”, Advanced Material
Research, Vol. 10, ISSN:1662-8985.
STO-TR-AVT-275 31 - 9
METALLIC HELICOPTER BLADES COMPOSITE REPAIRS STUDY CASE
[3] European Aviation Safety Agency, (2008), EASA, CS-29 Certification Specifications for Large
Rotorcraft CS-29, Annex to ED Decision 2008/010/R.
[5] Operating Bulletin No. P/O/5073/E/08, AFIT Library, Warsaw, Poland, 2008.
[6] Salacinski, M., Broda, P., and Samoraj, P., (2017), “The Repair Design and Technology of Metal Rotor
Blades for Mi Family Helicopter – The Approach with the Usage of Reverse Engineering”, SAE
Technical Paper.
31 - 10 STO-TR-AVT-275
REPORT DOCUMENTATION PAGE
1. Recipient’s Reference 2. Originator’s References 3. Further Reference 4. Security Classification
of Document
STO-TR-AVT-275 ISBN
AC/323(AVT-275)TP/930 978-92-837-2255-7 PUBLIC RELEASE
5. Originator
Science and Technology Organization
North Atlantic Treaty Organization
BP 25, F-92201 Neuilly-sur-Seine Cedex, France
6. Title
Continuing Airworthiness of Aging Systems
7. Presented at/Sponsored by
Final report on best practices and lessons learned for aircraft structural,
propulsion, and mechanical systems safety and reliability. The report Annex,
STO-TR-AVT-275-A, contains Chapters 10, 23, and 29.
8. Author(s)/Editor(s) 9. Date
Multiple 260
12. Distribution Statement
There are no restrictions on the distribution of this document.
Information about the availability of this and other STO
unclassified publications is given on the back cover.
13. Keywords/Descriptors
Aging systems; Aircraft structures; Aircraft Structural Integrity Process (ASIP);
Continuing airworthiness; Electrical wiring interconnect system; Mechanical system;
Propulsion system
14. Abstract
Due to high procurement costs associated with replacing aging aircraft fleets, NATO Nations are
frequently required to operate their aircraft for longer than the original design life. The objective of
this Task Group was to develop a technical report containing the best practices and lessons learned
that exist in NATO nations for aircraft structural, propulsion, and mechanical systems. The
documented best practices on continuing airworthiness of aging aircraft systems aim at capturing the
unique aptitudes that have developed in each participating nation. With these documented practices,
the different NATO nations will be able to adopt them as they see appropriate.
Aircraft structural systems have the most developed integrity program, with Propulsion systems
well-defined and becoming matured. Mechanical systems are progressing, but many lessons from
structural and propulsion integrity programs could be used to further mechanical systems integrity
program efforts. Employing aircraft structures-like methods may improve the overall aircraft
continued airworthiness, especially at or beyond the original service life limit.
STO-TR-AVT-275
STO-TR-AVT-275
NORTH ATLANTIC TREATY ORGANIZATION SCIENCE AND TECHNOLOGY ORGANIZATION
BP 25
DIFFUSION DES PUBLICATIONS
F-92201 NEUILLY-SUR-SEINE CEDEX • FRANCE
Télécopie 0(1)55.61.22.99 • E-mail mailbox@cso.nato.int STO NON CLASSIFIEES
Les publications de l’AGARD, de la RTO et de la STO peuvent parfois être obtenues auprès des centres nationaux de distribution indiqués ci-
dessous. Si vous souhaitez recevoir toutes les publications de la STO, ou simplement celles qui concernent certains Panels, vous pouvez demander
d’être inclus soit à titre personnel, soit au nom de votre organisation, sur la liste d’envoi.
Les publications de la STO, de la RTO et de l’AGARD sont également en vente auprès des agences de vente indiquées ci-dessous.
Les demandes de documents STO, RTO ou AGARD doivent comporter la dénomination « STO », « RTO » ou « AGARD » selon le cas, suivi du
numéro de série. Des informations analogues, telles que le titre est la date de publication sont souhaitables.
Si vous souhaitez recevoir une notification électronique de la disponibilité des rapports de la STO au fur et à mesure de leur publication, vous pouvez
consulter notre site Web (http://www.sto.nato.int/) et vous abonner à ce service.
Les demandes de documents STO, RTO ou AGARD doivent comporter la dénomination « STO », « RTO » ou « AGARD » selon le cas, suivie du numéro
de série (par exemple AGARD-AG-315). Des informations analogues, telles que le titre et la date de publication sont souhaitables. Des références
bibliographiques complètes ainsi que des résumés des publications STO, RTO et AGARD figurent dans le « NTIS Publications Database »
(http://www.ntis.gov).
NORTH ATLANTIC TREATY ORGANIZATION SCIENCE AND TECHNOLOGY ORGANIZATION
BP 25
DISTRIBUTION OF UNCLASSIFIED
F-92201 NEUILLY-SUR-SEINE CEDEX • FRANCE
Télécopie 0(1)55.61.22.99 • E-mail mailbox@cso.nato.int STO PUBLICATIONS
AGARD, RTO & STO publications are sometimes available from the National Distribution Centres listed below. If you wish to receive all STO
reports, or just those relating to one or more specific STO Panels, they may be willing to include you (or your Organisation) in their distribution.
STO, RTO and AGARD reports may also be purchased from the Sales Agencies listed below.
Requests for STO, RTO or AGARD documents should include the word ‘STO’, ‘RTO’ or ‘AGARD’, as appropriate, followed by the serial number.
Collateral information such as title and publication date is desirable.
If you wish to receive electronic notification of STO reports as they are published, please visit our website (http://www.sto.nato.int/) from where you
can register for this service.
Requests for STO, RTO or AGARD documents should include the word ‘STO’, ‘RTO’ or ‘AGARD’, as appropriate, followed by the serial number
(for example AGARD-AG-315). Collateral information such as title and publication date is desirable. Full bibliographical references and abstracts of
STO, RTO and AGARD publications are given in “NTIS Publications Database” (http://www.ntis.gov).
ISBN 978-92-837-2255-7