CASA B2-13b Aircraft Structures PDF

Student Resource Subject B2-13b: Aircraft Structures 2013-06-28 B2-13b Aircraft Structures Page 1 of 6 Training Material Only  Copyright © 2017 Aviation Australia All rights reserved. No part of this document m...

Student Resource Subject B2-13b: Aircraft Structures 2013-06-28 B2-13b Aircraft Structures Page 1 of 6 Training Material Only  Copyright © 2017 Aviation Australia All rights reserved. No part of this document may be reproduced, transferred, sold, or otherwise disposed of, without the written permission of Aviation Australia CONTROLLED DOCUMENT 2013-06-28 B2-13b Aircraft Structures Page 2 of 6 Training Material Only CONTENTS Definitions 2 Study Resources 3 Introduction 5 Topic 13.2.1: Airframe Structures – General Concepts 5 2013-06-28 B2-13b Aircraft Structures Page 3 of 6 Training Material Only DEFINITIONS Define To describe the nature or basic qualities of. To state the precise meaning of (a word or sense of a word). State Specify in words or writing. To set forth in words; declare. Identify To establish the identity of. List Itemise. Describe Represent in words enabling hearer or reader to form an idea of an object or process. To tell the facts, details, or particulars of something verbally or in writing. Explain Make known in detail. Offer reason for cause and effect. 2013-06-28 B2-13b Aircraft Structures Page 4 of 6 Training Material Only STUDY RESOURCES Jeppesen A&P Technicians General Textbook Jeppesen A&P Technicians Airframe Textbook Jeppesen A&P Technicians Powerplant Text book Cessna 100 Series Airplane AMM Boeing 737NG AMM Boeing 747-400 AMM Boeing 727 AMM Fokker F28 AMM Piper PA31 AMM AC-43 FAR 25 (FAA USA B2-13b Student Resource 2013-06-28 B2-13b Aircraft Structures Page 5 of 6 Training Material Only INTRODUCTION The purpose of this subject is to familiarise you with the basic construction, requirements and maintenance of aircraft structures. On completion of the following topics you will be able to: Topic 13.2 Airframe Structures – General Concepts Identify airworthiness requirements for structural strength. Identify inspection programs associated with ageing aircraft. Identify primary, secondary and tertiary structural classifications. Identify fail safe, safe life and damage tolerance concepts. Describe zonal and station identification systems. Identify the following concepts with regards to airframe structures: Stress, strain, bending, compression, shear, torsion, tension, hoop stress and fatigue. Identify drain and ventilation provisions used in airframe structures. Identify provisions used in aircraft for system installations. Describe provisions for lightning protection. Describe aircraft electrical bonding concepts. 2013-06-28 B2-13b Aircraft Structures Page 6 of 6 Training Material Only TOPIC 13.2 AIRFRAME Structures – General CONCEPTS TABLE OF CONTENTS TABLE OF FIGURES............................................................................................................................ 3 Topic 13.2 Airframe Structures – General Concepts Part a................................................................. 5 Airworthiness requirements............................................................................................................ 5 Transport Category Aircraft.............................................................................................................. 6 FAR 25.301 Loads............................................................................................................................. 6 FAR 25.303 Factor of Safety............................................................................................................. 7 FAR 25.305 Strength and Deformation............................................................................................ 7 FAR 25.307 Proof of Structure......................................................................................................... 8 Structural classifications................................................................................................................... 9 Primary Structure.......................................................................................................................... 9 Secondary Structure..................................................................................................................... 9 Tertiary Structure........................................................................................................................ 10 Structural Design Requirement...................................................................................................... 10 Static Strength............................................................................................................................ 11 Fatigue Design Philosophy.......................................................................................................... 11 Safe Life....................................................................................................................................... 13 Safe Life Disadvantages.............................................................................................................. 13 Fail Safe....................................................................................................................................... 13 Fail Safe Disadvantages.............................................................................................................. 14 Damage Tolerance...................................................................................................................... 14 Durability.................................................................................................................................... 15 Ageing Aircraft Inspection Requirements................................................................................... 16 Task 5 – 14 CFR/JAR 25 Ageing Aircraft.......................................................................................... 16 Background................................................................................................................................. 16 Zonal and Station Identification Systems....................................................................................... 19 Aircraft Reference Zones............................................................................................................ 19 Major Zone.................................................................................................................................. 19 Sub Zone..................................................................................................................................... 20 Zone............................................................................................................................................ 20 Reference Datum........................................................................................................................ 20 Body Station Line (BS)................................................................................................................. 22 Buttock Line (BL)......................................................................................................................... 22 2015-08-24 B2-13.2 Structures – General Concepts Page 1 of 42 Training Material Only  Water Line (WL).......................................................................................................................... 23 Wing............................................................................................................................................... 23 Wing Buttock Line (WBL)............................................................................................................ 23 Engine......................................................................................................................................... 24 Clock Position.............................................................................................................................. 25 Stress.............................................................................................................................................. 25 Strain.............................................................................................................................................. 26 Stress-Strain Curves.................................................................................................................... 26 Tension........................................................................................................................................... 27 Compression............................................................................................................................... 28 Shear........................................................................................................................................... 28 Torsion........................................................................................................................................ 29 Bending....................................................................................................................................... 30 Hoop Stress................................................................................................................................. 31 Fatigue............................................................................................................................................ 31 Aircraft Drainage............................................................................................................................ 32 Drainage...................................................................................................................................... 33 Pressurised Fuselage Drain Valve............................................................................................... 33 Drain Hole................................................................................................................................... 34 Ventilation...................................................................................................................................... 34 System Installation Provisions.................................................................................................... 35 Hydraulic Systems....................................................................................................................... 35 Air Conditioning System............................................................................................................. 35 Electrical/Avionics....................................................................................................................... 35 Fuel Storage................................................................................................................................ 36 Main Landing Gear (MLG).............................................................................................................. 36 Waste Disposal........................................................................................................................... 37 Potable Water............................................................................................................................. 38 Auxiliary Power Unit (APU)............................................................................................................. 38 Lightning Strikes.......................................................................................................................... 39 Aircraft Electrical Bonding.............................................................................................................. 41 Bonding....................................................................................................................................... 41 Grounding................................................................................................................................... 41 Bonding Leads............................................................................................................................. 41 2015-08-24 B2-13.2 Structures – General Concepts Page 2 of 42 Training Material Only TABLE OF FIGURES Figure 1: Regulatory Authorities.......................................................................................................... 5 Figure 2: Transport Category Aircraft.................................................................................................. 6 Figure 3: Load Analysis......................................................................................................................... 7 Figure 4: Fuselage Pressure Testing – B787......................................................................................... 8 Figure 5: Proving Structural Analysis - B787 Wing............................................................................... 9 Figure 6: Structural Classification...................................................................................................... 10 Figure 7: Static Strength Testing – B787............................................................................................ 11 Figure 8: Fatigue Testing.................................................................................................................... 12 Figure 9: Fatigue Design History........................................................................................................ 12 Figure 10: Fail Safe Design - Multi Spar Structure............................................................................. 13 Figure 11: Damage Tolerance Inspection Regime............................................................................. 14 Figure 12: Non Destructive Testing.................................................................................................... 15 Figure 13: Aircraft Durability.............................................................................................................. 15 Figure 14: Major Zones...................................................................................................................... 19 Figure 15 Sub Zones........................................................................................................................... 20 Figure 16: Reference Datum.............................................................................................................. 21 Figure 17: B737NG Datum Line.......................................................................................................... 21 Figure 18: Fuselage Station Lines....................................................................................................... 22 Figure 19: Buttock and Water Lines................................................................................................... 23 Figure 20: Wing Buttock Lines........................................................................................................... 24 Figure 21: Engine Station Lines.......................................................................................................... 24 Figure 22: Clock Position.................................................................................................................... 25 Figure 23: Stress-Strain Curves.......................................................................................................... 26 Figure 24: Compression Stresses....................................................................................................... 28 Figure 25: Rivets Under Shear Loads................................................................................................. 29 Figure 26: Torsional Stress on a Fuselage.......................................................................................... 29 Figure 27: Bending Force on a Fuselage............................................................................................ 30 Figure 28: Bending Force on Wings................................................................................................... 30 Figure 29: Hoop Stress to Fuselage.................................................................................................... 31 Figure 30: Fatigue Damage................................................................................................................ 32 Figure 31: Pressurised Fuselage Drain Valve..................................................................................... 33 Figure 32: Drain Hole......................................................................................................................... 34 2015-08-24 B2-13.2 Structures – General Concepts Page 3 of 42 Training Material Only Figure 33: Ventilation......................................................................................................................... 34 Figure 34: Fuel Tank System.............................................................................................................. 36 Figure 35: Main Landing Gear Fittings............................................................................................... 36 Figure 36: A Waste Water Drain System............................................................................................ 37 Figure 37: Potable Water System...................................................................................................... 38 Figure 38: Lightning Strike on B747................................................................................................... 39 Figure 39: Bonding Strap attached to flight control.......................................................................... 39 Figure 40: Radome lightning diverter strips (F28)............................................................................. 40 Figure 41: Static Discharge Wicks...................................................................................................... 40 Figure 42: Radome Bonding Straps (F28)........................................................................................... 41 2015-08-24 B2-13.2 Structures – General Concepts Page 4 of 42 Training Material Only TOPIC 13.2 AIRFRAME STRUCTURES – GENERAL CONCEPTS PART A Airworthiness requirements An aircraft is airworthy only when it conforms to the regulations under which it has been certified. Regulatory authorities work closely on setting uniform requirements. Airworthiness standards are organised into sections, called parts, and each part deals with a specific type of activity. As an example: Part 23 covers airworthiness standards for Normal, Utility, Acrobatic and Commuter Aircraft Part 25 covers airworthiness standards for Transport Category Aircraft Part 27 covers airworthiness standards for Normal Category Rotorcraft Part 29 covers airworthiness standards for Transport Category Rotorcraft Part 33 covers airworthiness standards for Aircraft Engines Part 35 covers airworthiness standards for Propellers Part 39 covers Airworthiness Directives Part 43 covers Maintenance, Preventive Maintenance, Rebuilding and Alteration Part 65 (Part 66 for EASA and CASA) covers Aircraft Maintenance Engineer Licensing Part 145 covers Aircraft Maintenance Organisations Part 147 covers Aviation Maintenance Technicians Schools Figure 1: Regulatory Authorities 2015-08-24 B2-13.2 Structures – General Concepts Page 5 of 42 Training Material Only Transport Category Aircraft An aircraft registered in the transport category requires that its structure meets the airworthiness standards specified in the following regulations as applicable for the country in which the aircraft is registered: CASR 1998 Part 25 Subpart 25, as of the 1st Edition - January 2003 for Australia EASA Certification Specifications CS-25 Subpart C, as of the Amendment 6 - July 2009 for Europe FAR Part 25 Subpart C, as of the e-CFR Data – August 2009 for the USA, covers four sections relating to the aircraft’s structure: – 25.301 Loads – 25.303 Factor of Safety – 25.305 Strength and Deformation – 25.307 Proof of Structure The content of these four sections is repeated below. Figure 2: Transport Category Aircraft FAR 25.301 Loads (a) Strength requirements are specified in terms of limit loads (the maximum loads to be expected in service) and ultimate loads (limit loads multiplied by prescribed factors of safety). Unless otherwise provided, prescribed loads are limit loads. The limit load is the maximum anticipated load or combination of loads, which a structure may be expected to experience. Ultimate load is the maximum load which a structure is designed to withstand without failure. (b) Unless otherwise provided, the specified air, ground, and water loads must be placed in equilibrium with inertia forces, considering each item of mass in the aeroplane. These loads must be distributed to conservatively approximate or closely represent actual conditions. 2015-08-24 B2-13.2 Structures – General Concepts Page 6 of 42 Training Material Only Methods used to determine load intensities and distribution must be validated by flight load measurement unless the methods used for determining those loading conditions are shown to be reliable. (c) If deflections under load would significantly change the distribution of external or internal loads, this redistribution must be taken into account. Figure 3: Load Analysis FAR 25.303 Factor of Safety Unless otherwise specified, a factor of safety of 1.5 must be applied to the prescribed limit load which is considered external loads on the structure. When loading condition is prescribed in terms of ultimate loads, a factor of safety need not be applied unless otherwise specified. FAR 25.305 Strength and Deformation (a) The aircraft structure must be able to support limit loads without detrimental permanent deformation. At any load up to limit loads, the deformation may not interfere with safe operation. (b) The structure must be able to support ultimate loads without failure for at least 3 seconds. However, when proof of strength is shown by dynamic tests simulating actual load conditions, the 3-second limit does not apply. Static tests conducted to ultimate load must include the ultimate deflections and ultimate deformation induced by the loading. When analytical methods are used to show compliance with the ultimate load strength requirements, it must be shown that: The effects of deformation are not significant The deformations involved are fully accounted for in the analysis, or The methods and assumptions used are sufficient to cover the effects of these deformations 2015-08-24 B2-13.2 Structures – General Concepts Page 7 of 42 Training Material Only (c) Where structural flexibility is such that any rate of load application likely to occur in the operating conditions might produce transient stresses appreciably higher than those corresponding to static loads, the effects of this rate of application must be considered. (d) Reserved (e) The aircraft must be designed to withstand any vibration and buffeting that might occur in any likely operating condition up to Vd/Md, (design dive speed), including stall and probable inadvertent excursions beyond the boundaries of the buffet onset envelope. This must be shown by analysis, flight tests, or other tests found necessary by the Administrator. (f) Unless shown to be extremely improbable, the aircraft must be designed to withstand any forced structural vibration resulting from any failure, malfunction or adverse condition in the flight control system. These must be considered limit loads and must be investigated at airspeeds up to Vc/Mc, (design cruise speed). Figure 4: Fuselage Pressure Testing – B787 FAR 25.307 Proof of Structure (a) Compliance with the strength and deformation of this subpart must be shown for each critical loading condition. Structural analysis may be used only if the structure conforms to that for which experience has shown this method to be reliable. The Administrator may require ultimate load tests in cases where limit load tests may be inadequate. (b), (c) Reserved. (d) When static or dynamic tests are used to show compliance with the requirements of 25.305(b) for flight structures, appropriate material correction factors must be applied to the test results, unless the structure, or part thereof, being tested has features such that a number of elements contribute to the total strength of the structure and the failure of one element results in the redistribution of the load through alternate load paths. 2015-08-24 B2-13.2 Structures – General Concepts Page 8 of 42 Training Material Only  Figure 5: Proving Structural Analysis - B787 Wing Structural classifications Aircraft structures and components are classified according to their level of importance for the part they play in maintaining the structural integrity of the aircraft. This classification determines the significantly different levels of intensity and frequency of inspection and maintenance schedules, component replacement, damage assessment and repair schemes. The classifications are: Primary structure Secondary structure Tertiary structure Primary Structure Primary structure is that structure that is critical to the safety of the aircraft. Should a primary structural component fail, during take-off, during flight, or during landing, there is potential for structural collapse, or loss of control, or failure of motive power, or fatality, or serious injury to aircrew. Secondary Structure Secondary structure refers to structure that, if it were to fail, may cause significant damage that would affect the operation of the aircraft, but not lead to its loss. 2015-08-24 B2-13.2 Structures – General Concepts Page 9 of 42 Training Material Only  Figure 6: Structural Classification Tertiary Structure Tertiary structure refers to structure that, if it were to fail, would not significantly affect the operation of the aircraft. Structural Design Requirement The structural design criteria are determined by the type of aircraft and its intended use. From this it is possible to define the structural design analysis to consider the types of manoeuvres, speeds, useful loads and gross weights the structure will be subject to. In addition the criteria must consider such items as inadvertent manoeuvres, effects of turbulent air, and severity of ground contact during landing. The strength provided in the airframe structure to meet these conditions must be adequate for the aircraft to perform its intended mission in a safe and profitable manner as operated by qualified personnel under regulated conditions. Certification of airframe structure generally requires that the structure be subject to testing and/or analysis to demonstrate the following capabilities: Static strength Fatigue strength 2015-08-24 B2-13.2 Structures – General Concepts Page 10 of 42 Training Material Only Static Strength The Design Limit Load (DLL) is the maximum load anticipated on the aircraft during its service life. The aircraft structure shall be capable of supporting the limit loads without suffering detrimental permanent deformation, and not interfere with the safe operation of the aircraft. Design Ultimate Load (DUL) is equal to the design limit load multiplied by a factor of safety. Generally the safety factor is 1.5. Figure 7: Static Strength Testing – B787 Fatigue Design Philosophy Fatigue considerations comprise an important part of aircraft structural design. The common objective is to define aircraft life to ensure flight safety while at the same time minimising maintenance and operating costs. There are three distinct design approaches for protecting an aircraft structure from failure due to fatigue damage: Safe life – Also known as safety by retirement, was introduced in the 1940s after fatigue was recognised as a failure mechanism. It specifies a safe lifespan within which there is no significant risk of structural failure of a component Fail-safe - The fail-safe design principle was introduced in the 1950s as an improvement to safe-life. A fail-safe structure should be able to sustain the limit load even when one of the elements has failed Damage tolerance - Damage tolerance, or safety by inspection, was developed as a design philosophy in the 1970s as an improvement on the fail-safe principle. Damage tolerance is the present method of achieving structural operating safety. 2015-08-24 B2-13.2 Structures – General Concepts Page 11 of 42 Training Material Only  Figure 8: Fatigue Testing Although the objectives of three approaches are the same, they vary with regard to the fundamental definition of service life. Figure 9: Fatigue Design History Note: SSIP means Supplemental Structural Inspection Program. The establishment of a SSIP is to reach the safety standard according to FAR 25.571 Amendment 45 MSG refers to Maintenance Steering Group and is the method that aircraft manufacturers, operators and regulators use to develop the manufacturer’s initial maintenance schedule, as part of the work towards aircraft certification. The process continues throughout the aircraft type’s life. MSG-3 is the methodology that focuses on aircraft systems and the loss of system function or functions 2015-08-24 B2-13.2 Structures – General Concepts Page 12 of 42 Training Material Only Safe Life Safe-life is an approach where a structure is designed to withstand a certain number of events (flight cycles, landings, or flight hours) with a low probability that the strength of the structure will degrade below its designed ultimate strength before the end of its approved life. The structural component must remain crack free during service. This means the component must not develop any cracks as a result of fatigue, corrosion, or accidental damage during its specified service life. Once the service life of a component has been reached it is considered unserviceable and the aircraft’s airworthiness can only be maintained if the component is replaced. Safe Life Disadvantages There are two major drawbacks to the safe life approach: Components are taken out of service even though they may have substantial remaining lives, creating unnecessary costs Despite all precautions, cracks sometimes occur prematurely, creating a safety problem Fail Safe Fail safe is an approach, applied to a structure, designed to retain its required residual strength for a period of unrepaired use after a failure or partial failure of a principal structural element To achieve this requirement, a fail-safe design uses backup structures and secondary load paths. This principle relies on the fact that if the main load path fails, there is a secondary load path to ensure the safety of the aircraft until the failure can be detected. For example; a multi spar structure is a fail-safe structure – if one spar failed, the adjacent spar will support the load. Figure 10: Fail Safe Design - Multi Spar Structure 2015-08-24 B2-13.2 Structures – General Concepts Page 13 of 42 Training Material Only Fail Safe Disadvantages Despite all precautions, cracks sometimes occur prematurely and may not be detected, creating a safety problem. Fail safe methodologies do not consider material or manufacturing flaws. Damage Tolerance This approach can be applied to a structure that is able to sustain a given level of fatigue, corrosion, manufacturing defects, or accidental damage, and still withstand design loads without structural failure or excessive structural deformation for a predetermined period that allows for a set number of opportunities to detect the damage. Figure 11: Damage Tolerance Inspection Regime The damage tolerance approach is based on the principle that while cracks due to fatigue and corrosion will develop in the aircraft structure, the process can be understood and controlled. A key element is the development of a comprehensive programme of inspections to detect cracks before they can affect flight safety. That is, damage tolerant structures are designed to sustain cracks without catastrophic failure until the damage is detected in scheduled inspections and the damaged part is repaired or replaced. In addition, damage tolerance takes into account initial material or manufacturing flaws by assuming an initial crack, which the fail-safe principle does not do. A damage tolerant design should allow cracks to be detected before they reach the critical length that will lead to failure. To ensure that this occurs there should be at least two opportunities to detect the crack prior to it reaching its critical length. The damage tolerance philosophy uses testing and analysis to determine the critical crack length, the residual strength, and the inspection intervals. Tests include flight testing to determine the loads on the structure, and ground testing to determine the fatigue and crack growth characteristics. From the testing and analysis, the critical sites and components susceptible to fatigue can be determined. Fatigue analysis based on flight, ground, and pressurisation loads can then be used to determine crack growth performance and residual strength. 2015-08-24 B2-13.2 Structures – General Concepts Page 14 of 42 Training Material Only  Figure 12: Non Destructive Testing The damage tolerance philosophy is used for transport category aircraft. For other categories of aircraft, the damage tolerance method may be used but is not mandatory; the safe-life or fail-safe methods may be used instead. Durability Durability is the ability of the structure to sustain degradation from such sources as fatigue, accidental damage and environment deterioration, to the extent that they can be controlled by economically acceptable maintenance and inspection programs. Durability of an aircraft structure comes from having a slow crack growth characteristic and the ability to contain or restrict the progress of damage. Figure 13: Aircraft Durability 2015-08-24 B2-13.2 Structures – General Concepts Page 15 of 42 Training Material Only Ageing Aircraft Inspection Requirements The ageing aircraft program requires each transport aircraft to undergo a series of tests and inspections after it has been in service for a specified number of years and/or has accumulated a certain number of flight cycles. These inspections are generally a combination of visual and non- destructive testing methods such as ultrasound, eddy current and x-rays. Corrosion control and structural repairs are made when necessary. Since major disassembly of the aircraft is required to complete the relevant inspections, rigging of the aircraft control system will be necessary. The aged aircraft program must be part of the scheduled maintenance program conducted by an approved Part 145 organisation. Some manufacturers may require major structural disassembly after a certain number of flight hours. For example, some Lear jet models require removal of the wing after approximately 5,000n flight hours for a series of Inspections. The manufacturer of an aircraft may specify this type of disassembly and inspection after a specific period of flight hours, flight cycles, months or years. Task 5 – 14 CFR/JAR 25 Ageing Aircraft Background In 1988, the industry experienced a significant failure of the airworthiness system. This system failure allowed an aeroplane to fly with significant unrepaired multiple site fatigue damage to the point where the aeroplane experienced a rapid fracture and loss of a portion of the fuselage. As a direct result of this accident, the FAA hosted The International Conference on Ageing Aeroplanes on June 1-3, 1988 in Washington D. C. As a result of this conference, an organisation of Operators, Manufacturers and Regulators was formed under the Federal Advisory Committee Act to investigate and propose solutions to the problems evidenced as a result of the accident. This group is now known as the Airworthiness Assurance Working Group (AAWG) (Formally known as the Airworthiness Assurance Task Force). During the 1988 conference, several Airline/Manufacturer recommendations were presented to address the apparent short falls in the airworthiness system including Recommendation 3, which stated: "Continue to pursue the concept of teardown of the oldest airline aircraft to determine structural condition, and conduct fatigue tests of older aeroplanes as per attached proposal" In June 1989, the National Transportation Safety Board (NTSB) made Recommendation 89067 that requested the FAA to pursue necessary tasks to ensure continued safe operations with probable Widespread Fatigue Damage (WFD). WFD was noted by the NTSB to be a contributing cause of the April 1988 Aloha Airlines 737 accident. The NTSB specifically recommended extended fatigue testing for older aeroplanes. In November 1989, the FAA responded by issuing a ‘strawman’ Supplementary Federal Aviation Regulation (SFAR) RE: TWO-LIFE-TIME FATIGUE TEST FOR OLDER AEROPLANES. (The premise behind building a ‘strawman’ – creating a first draft for criticism and testing, and then using the feedback you receive to develop a final outcome that is rock solid.) In June 1990, the AAWG tasked the formal evaluation of the Aerospace Industries Association of America International Air Transport Association (AIAIATA) Recommendation 3. An alternative approach, to the strawman SFAR was developed by the AAWG and presented to the FAA in March 1991. 2015-08-24 B2-13.2 Structures – General Concepts Page 16 of 42 Training Material Only The FAA accepted this alternative approach in June 1991. The AAWG was informally tasked to institutionalize the position in July. It is urged that any future publications on the subject of widespread fatigue damage should include, or at least reference this standard terminology, in order to avoid possible confusion within the industry. Damage Tolerance is the attribute of the structure that permits it to retain its required residual strength without detrimental structural deformation for a period of use after the structure has sustained specific levels of fatigue, corrosion, accidental or discrete source damage. Widespread Fatigue Damage (WFD) in a structure is characterized by the simultaneous presence of cracks at multiple structural details that are of sufficient size and density whereby the structure will no longer meet its damage tolerance requirement (i.e. to maintain its required residual strength after partial structural failure). Multiple Site Damage (MSD) is a source of widespread fatigue damage characterized by the simultaneous presence of fatigue cracks in the same structural element (i.e. fatigue cracks that may coalesce with or without other damage leading to a loss of required residual strength). Multiple Element Damage (MED) is a source of widespread fatigue damage characterized by the simultaneous presence of fatigue cracks in similar adjacent structural elements. In addition, the AAWG proposes the adoption of the following terminology during discussion of programs to ensure continuing structural integrity: Fatigue Crack Initiation is that point in time when a finite fatigue crack is first expected. Point of WFD is a point reduced from the average expected behaviour, i.e. lower bound, so that operation up to that point provides equivalent protection to that of a two-lifetime fatigue test. Monitoring Period is the period of time when special inspections of the fleet are initiated due to an increased risk of MSD/MED, and ending when the point of WFD is established. Design Service Goal (DSG) is the period of time (in flight cycles/hours) established at design and/or certification during which: 1. The principal structure will be reasonably free from significant cracking 2. Widespread fatigue damage is not expected to occur. Extended Service Goal (ESG) is an adjustment to the design service goal established by service experience, analysis, and/or test during which: 1. The principal structure will be reasonably free from significant cracking 2. Widespread fatigue damage is not expected to occur. Furthermore, certain terminology has been considered by past working groups in relation to the problem of WFD, but was not used in the final ARAC definitions. The following terms have been previously identified as being open to misinterpretation, and should be avoided, or defined carefully if their use is essential. 2015-08-24 B2-13.2 Structures – General Concepts Page 17 of 42 Training Material Only Threshold has been used in various contexts, such as: Fatigue Threshold, which may be defined as the first typical fatigue crack in the fleet for that element Inspection Threshold, which may be defined as the start of supplemental inspections for WFD The AAWG believes that the real meaning of WFD in this context is MSD/MED. Onset has been used as an alternative to Threshold, although the simultaneous use of both terms may cause confusion. Sub-Critical has been used in relation to certain fatigue cracks. However, this may require clarification of what are critical fatigue cracks with reference to occurrence of WFD. There are a number of general conditions and details that must be met in order that a monitoring period concept can be used. These conditions are: No aeroplane may be operated beyond the defined Point of WFD without modification or part replacement. The first special inspections, to occur in the monitoring period, should be in line with the estimation of fatigue crack initiation. To use a monitoring period for a detail suspected of developing MSD/MED, it must be determined that inspections will reliably detect a crack before the crack becomes critical. If a crack cannot be reliably detected, a monitoring period cannot be used. By empirical analysis, evaluation of test evidence and/or evaluation of in-service data, the inspection requirements will be defined for application during the monitoring period. The purpose of these inspections is to collect data for reassessment of WFD parameters and to maintain structural integrity (e.g., acceptable level of risk during the monitoring period). Inspections within the monitoring period are mandatory on every aeroplane as well as reporting of inspection results. In the case of MSD or MED findings, the Point of WFD will be re-established in accordance to the inspection results. The area of concern will be repaired following a detailed inspection of adjacent areas using NDI technology that will detect small cracks with a high degree of confidence. The remaining aeroplanes may be operated up to the revised Point of WFD, with application of a revised monitoring program. Prior to the Point of WFD, the aeroplane must be repaired, modified, or retired. If no MSD/MED cracking is detected by the time the high time aeroplane reaches the predicted Point of WFD, the predicted Point of WFD could be re-evaluated and the special inspection program may be continued after revalidation. The monitoring period will terminate at the point in time at which there is sufficient findings to confirm a MSD/MED problem exists and/or the Point of WFD is reached. This will be recommended with the assistance of the STG using an established process. 2015-08-24 B2-13.2 Structures – General Concepts Page 18 of 42 Training Material Only Zonal and Station Identification Systems Aircraft Reference Zones In order to facilitate maintenance and component location, an aircraft is divided into Major Zones. Each zone is further broken down into Sub Zones and Zones. Different manufactures refer to the term Major Sub-Zones instead of Sub Zones; however the identification principle is the same. Each of the zones sequentially produces a three-digit identifier. The identifier is in a standard format as defined by the ATA100 Specification. Major Zone The aircraft is divided into 8 major zones. Each zone is numbered from 100 to 800, with the left hand digit from 1 to 8 and each followed by two zeros to create the three digit identifier. For example, major zone 100 is the lower fuselage; major zone 200 is the upper fuselage; and major zone 300 is the Empennage. Figure 14: Major Zones 2015-08-24 B2-13.2 Structures – General Concepts Page 19 of 42 Training Material Only Sub Zone It is possible to have 9 sub zones in each major zone. A sub zone is identified by the middle digit of the three digit identifier, using the numbers 1 to 9, and the right hand digit a zero. For example, sub zone 820 refers to cargo compartment doors, sub zone 830 refers to the left side passenger compartment doors, and sub zone 840 refers to the right side passenger doors. Figure 15 Sub Zones Zone The third digit, comprising the numbers 1 to 9 identifies a component or group of components that are in the subzone. For example, zone 821 identifies the forward cargo door and zone 822 identifies the aft cargo door. Access doors and panels in a zone are identified by the zone number and a two or three letter suffix. This alpha-numeric label is different for each access door or panel. For example: 821 is the forward cargo door (RH side) – 821AR is an access panel on the forward cargo door – 821AZ is an access panel on the forward cargo door liner Reference Datum A reference datum or datum is established by the manufacturer for each aircraft model. The datum is an imaginary vertical plane or line that sits at a right angle to the aircraft’s longitudinal axis. The datum is the point from which all horizontal measurements are taken, with the aircraft in a level flight attitude. The location of all items, including equipment, tanks, baggage compartments, seats, engines, and propellers are listed as being so many inches from the datum. 2015-08-24 B2-13.2 Structures – General Concepts Page 20 of 42 Training Material Only Information on the location of the datum is found in the Aircraft Specifications or Type Certificate Data Sheets. Figure 16: Reference Datum There is no fixed rule for the location of a datum. It may be located on the nose of the aircraft, the fire-wall, the leading edge of the wing, or even at a point in space ahead of the aircraft. The manufacturer chooses a location for the datum where it is most convenient for measurement, the location of equipment, and for weight and balance computation. For example, the datum plane in B737NG is perpendicular to the fuselage centreline and found 130.0 inches forward of the aeroplane nose. Figure 17: B737NG Datum Line Position Identification and Location Using the datum line as the prime reference point, it is possible to identify an exact point, for example, a specific rivet, anywhere on the aircraft. 2015-08-24 B2-13.2 Structures – General Concepts Page 21 of 42 Training Material Only Fuselage For the fuselage, the three planes that give an exact location are termed: The Fuselage Station line (FS) measured in millimetres from the datum line (Airbus), or Body Station (BS), measured in inches from the datum line (Boeing) The Left Buttock Line (LBL), or Right Buttock Line (RBL), measured from the longitudinal centreline of the fuselage The Water Line (WL), measured vertically from an imaginary water line zero, placed some distance below the fully extended undercarriage of the aircraft Body Station Line (BS) A fuselage station is a vertical line perpendicular to the body centreline. In the case of the 747 aircraft, the datum line is 90 inches forward of the radome. So B STA 90 refers to the body or fuselage station which is 90 inches back from the datum line. Likewise, it is possible to determine that body station 2360 is where the stabiliser leading edge contacts the fuselage. Figure 18: Fuselage Station Lines Buttock Line (BL) A body buttock line is a vertical line which establishes lateral distance to the left and right of the fuselage’s vertical centreline. Buttock Line zero is the centre line of the aircraft. Distances left (LBL) or right (RBL) identify position on the left or right side of the aircraft. Left and right designations are always referenced facing forward in the aeroplane. 2015-08-24 B2-13.2 Structures – General Concepts Page 22 of 42 Training Material Only Water Line (WL) A waterline is a horizontal line which establishes vertical distances from the top to the bottom of the aeroplane. Water line 0, in the case of a 747 is 91 inches below the lowest point of the fuselage. Figure 19: Buttock and Water Lines Wing Wing Buttock Line (WBL) The use of a wing buttock line is one method of providing a location reference along the wing. These stations are measured from the centre line of the aircraft, or buttock line zero. They indicate the distance in inches along the wing towards the wing tip. More accurate references are obtained using the term wing station (WS) where the reference is against the wing’s rear spar. 2015-08-24 B2-13.2 Structures – General Concepts Page 23 of 42 Training Material Only  Figure 20: Wing Buttock Lines Engine The engine station numbers show the locations of the structural components and features on the engine. Figure 21: Engine Station Lines 2015-08-24 B2-13.2 Structures – General Concepts Page 24 of 42 Training Material Only Clock Position Another position referencing system often used by engineers or flight crew is clock positions. They are used to report faults on the aircraft or engines using a clock face as a reference, e.g. damage at 2 O’clock position. The reference for this is from within the aircraft looking forward. Figure 22: Clock Position Being able to identify the zone and exact point anywhere within that zone provides important information, especially when writing up defects found during inspections. Correctly identifying the location of components prevents the possibility of replacement of an incorrect component. If the defect found requires a structural repair, it ensures that the correct repair scheme appropriate for the structural classification is selected. Finally, the correct identification of the location of a defect ensures that the certification of the work performed, does directly refer to that work. Stress Aircraft structures must be able to withstand all flight conditions and be able to operate under all payload conditions. In aircraft structures stress is defined as the force or load (F) applied to an element (beam, bulkhead or skin) divided by its cross sectional area (A), being: l = F/A, where the l indicates the longitudinal direction. The SI unit for stress is the Pascal (symbol Pa), which is equivalent to one Newton (force) per square meter (unit area). The unit for stress is the same as that of pressure, which is also a measure of force per unit area. 2015-08-24 B2-13.2 Structures – General Concepts Page 25 of 42 Training Material Only There are five main types of stress loads that an aircraft is subjected to: Tension Compression Shear Bending Torsion Strain The stresses within a structure must be kept below a defined permitted level. Where stress is, so is strain. It is impossible to be subjected to stress without experiencing strain. Strain is the linear deformation of the element divided by its original length/size. When considering length deformation only the applicable formula is: Strain = change in length / original length Stress-Strain Curves Stress-strain curve is used by engineers to interpret material strength. To obtain this type of curve it is necessary to carry out a tensile test on a ductile material such as a piece of low carbon steel. Figure 23: Stress-Strain Curves A test piece of known dimensions is placed in the testing machine. To determine its strength, a load (or force) is applied to “stretch” the test piece. The size of the load is measured and the stress induced in the piece is calculated by dividing the applied load by the cross sectional area of the test piece. The calculated value, stress, is represented on the vertical axis of the graph. 2015-08-24 B2-13.2 Structures – General Concepts Page 26 of 42 Training Material Only As the load increases the material “stretches”, it elongates. The elongation is measured and converted to strain by dividing the change in length by the original length of the test piece. The calculated value, strain, is represented on the horizontal axis of the graph. As the stress increases, the material stretches. Initially it is “elastic”. That is, if the load is removed the test piece will return to its original length. This occurs until the material reaches its “proportional limit” shown on the curve. This part is a straight line. This deformation is called “elastic deformation” and the ratio of stress to strain is a constant known as “Young’s Modulus”, denoted E, thus, σ = Eε. As the stress increases beyond the proportional limit many materials show a definite ‘yield point’ and this is often used as a basis (yield stress) for design calculations. For materials without a definite yield point (many brittle materials) an offset method is used to calculate a ‘theoretical’ yield which is used for design calculations. After reaching the proportional limit, if the stress increases, the material continues to stretch, but the elongation is now permanent. This is called “plastic deformation”, so if the load is removed the material will NOT return to its original length. The point of “ultimate stress” on the curve represents the material’s greatest ability to support a load. In engineering it is often called the “Ultimate tensile stress” (UTS) and is the figure quoted when we compare the strengths of different materials. Once the UTS has been reached the cross sectional area of the test piece reduces (a process known as necking) and the applied load drops accordingly. Fracture occurs at the end of the curve. When aircraft are designed the size of the parts is determined by the loads those parts have to withstand. Other factors, such as the cyclic nature of the loads, are also considered. As mentioned earlier, depending on the design process being applied a ‘factor of safety’ can be applied to either the ‘UTS’ or the ‘yield strength’ to provide a ‘maximum allowable stress’ from which the minimum safe size of components can be calculated. Tension Tension describes forces that tend to pull an object apart. It is a stress produced in a body by forces acting along the same line but in opposite directions, for example: Flexible steel cable used in aircraft control systems are designed to withstand tension loads Wing struts of a high wing aeroplane are under tension during normal flight The bolts used to fasten a windscreen to the fuselage are under tension stress during normal flight In the BAC1-111 accident in June 1990, the error was fitting the wrong bolts to the windscreen. The accident happened when the aircraft was climbing through 17,300 feet on departure from Birmingham. The left windscreen, which had been replaced prior to the flight, was blown out under effects of the cabin pressure when it overcame the retention of the securing bolts, 84 of which, out of a total of 90, were of smaller than specified diameter. The commander was sucked halfway out of the windscreen aperture and was restrained by cabin crew whilst the co-pilot flew the aircraft to a safe landing at Southampton Airport. This illustrates well the category of ‘wrong parts’. 2015-08-24 B2-13.2 Structures – General Concepts Page 27 of 42 Training Material Only Compression Compression is the opposite of tension. It is the resultant stress of two forces which act along the same line pushing against each other. Wing struts of a high wing aeroplane are under compression while on the ground stationary or taxiing – wings not producing lift. Aircraft rivets are driven with a compressive force. When compression stresses are applied to a rivet, the rivet shank expands until it fills the hole and forms a butt to hold materials together. Figure 24: Compression Stresses A landing gear oleo strut is under compression when an aircraft is on the ground. Shear Shear in an aircraft structure is a stress exerted when two pieces of fastened material tend to separate. Shear stress is the outcome of sliding one part over the other in opposite directions. The rivets are designed to withstand shear stresses when riveted the aluminium skin panels to the stringers. Shear forces try to rip the rivet in two therefore, selection of rivets with adequate shear resistance is critical. 2015-08-24 B2-13.2 Structures – General Concepts Page 28 of 42 Training Material Only Generally, in aircraft structure rivets are subjected to shear only, but bolts may be stressed by shear and tension, and in particular material, shear strength is less than tensile or compressive strength. Figure 25: Rivets Under Shear Loads Torsion Torsion is the stress applied to a material when it is twisted. It is a combination of tension and compression loads. For example, torsional stress on the fuselage is created by the action of the ailerons when the aircraft is manoeuvred. Torque (also a twisting force) works against torsion. Figure 26: Torsional Stress on a Fuselage 2015-08-24 B2-13.2 Structures – General Concepts Page 29 of 42 Training Material Only Bending Bending is the stress in an object caused by load being applied to one end while the other is restrained. Like torsion, bending stress is also a combination of tension and compression stresses. When an aircraft is on the ground, there is a bending force on the fuselage. This force occurs because of the weight of the aircraft. The bending action creates a tension stress on the lower skin of the fuselage and a compression stress on the top skin. These stresses are transmitted to the fuselage when the aircraft is in flight. Figure 27: Bending Force on a Fuselage When the aircraft is in flight, lift forces act upward against the wings, tending to bend them upward. The wings are prevented from folding over the fuselage by the resisting strength of the wing structure. The bending action creates a tension stress on the bottom of the wings and a compression stress on the top of the wings. Figure 28: Bending Force on Wings 2015-08-24 B2-13.2 Structures – General Concepts Page 30 of 42 Training Material Only Hoop Stress Considering a thin cylindrical shell subjected to an internal pressure as shown in Figure, tensile stress acting in a direction tangential to the circumference is called hoop stress or circumferential stress. Pressurised aircraft are subjected to hoop stress when the pressure inside the fuselage increases. The aircraft skin tries to expand and split along the longitudinal axis. Hoses are also susceptible to hoop stress. Figure 29: Hoop Stress to Fuselage Fatigue Fatigue is progressive localized structural damage. It occurs in a material subjected to repeated or fluctuating strains at stresses having a maximum value less than the ultimate tensile strength of the material. There are three requirements for a fatigue crack to form and spread in metals: There must be a local plastic stress There must be a tension stress There must be a cyclic (repeated or fluctuating) stress If we can eliminate any one of these three requirements, we can stop the fatigue process. It should be noted that composite materials can fatigue under compression loads. The insidious feature of fatigue failure is that there is no obvious warning, a crack forms without appreciable deformation of structure making it difficult to detect the presence of growing cracks. Fractures usually start from small nicks or scratches or fillets which cause a localised concentration of stress. Failure can be influenced by a number of factors including size, shape and design of the component, condition of the surface or operating environment. 2015-08-24 B2-13.2 Structures – General Concepts Page 31 of 42 Training Material Only On April 28, 1988, Aloha Airlines Flight 243 suffered a massive structural failure as part of the fuselage tore away from the Boeing 737 while in the air en route to Honolulu. It was later determined that the failure was caused by widespread fatigue damage in the aluminium skin of the fuselage. It was calculated that this particular aircraft had experienced 89,090 flight cycles over its 19 year life span. The investigation into this failure concluded that the widespread fatigue damage that led to the ultimate failure was caused by the aircraft’s exceptionally large number of flight cycles and accelerated by the exposure of the aircraft to corrosive salt water vapour during its inter-island flights. The combination of these two factors led to the creation of small cracks throughout the fuselage lap joints that eventually linked to form larger cracks and eventually complete failure. Figure 30: Fatigue Damage Aircraft Drainage Corrosion is one of the biggest threats to the integrity of aircraft structure. Three conditions must exist simultaneously for corrosion to take place: An anode and a cathode A metallic connector between the anode and cathode An electrolyte such as water Water cannot be avoided, but it can be controlled with drain paths, drain holes, sealants, and corrosion-inhibiting compounds. Controlling the presence of water is usually the most effective means of preventing corrosion Small quantities of water collect in the cabin because of condensation. Drains are installed in the fuselage structure to let the water out and to give more protection to the structure from corrosion. The centre fuselage drains are located at the lowest point of the cabin, below the cabin floor. 2015-08-24 B2-13.2 Structures – General Concepts Page 32 of 42 Training Material Only Drainage Effective drainage of all structure is vital to prevent fluids from becoming trapped in crevices. The entire lower fuselage of a pressurized aircraft is drained by a system of valved drain holes. Fluids are directed to these drain holes by a system of longitudinal and cross-drain paths through the stringers and frame shear clips. Pressurised Fuselage Drain Valve As the cabin pressurizes, the pressure sensitive rubber seal will move down to close the drain hole. When the hole is closed, loss of cabin pressure through the drain hole is prevented. When the cabin pressure decreases the rubber seal will move up to its usual position and open the drain hole. Any water that has collected will drain away through the main gear bay. The drain valves may be electrically heated to prevent freezing in a pressurised aircraft. Figure 31: Pressurised Fuselage Drain Valve 2015-08-24 B2-13.2 Structures – General Concepts Page 33 of 42 Training Material Only Drain Hole In non-pressurized aircraft, and non-pressurized components on pressurized aircraft, (flaps, ailerons, rudder, elevators), small drilled holes in the aircraft skin act as moisture drainage holes. They also allow fresh air to ventilate through the structure to help dry out any residual moisture. Figure 32: Drain Hole Ventilation Although partial fuselage and wing-to-body fairings are not pressurised structures, differential pressures can build up between the internal sections. Equal pressure in the sections is maintained by the pressure relief valves, located in the structure. The valves open, against spring tension, to release any overpressure build-up in the fairing sections. When the valves open, they remain open until the pressure equalises between the sections of the structure. Figure 33: Ventilation 2015-08-24 B2-13.2 Structures – General Concepts Page 34 of 42 Training Material Only System Installation Provisions A number of provisions are made on aircraft which help make a maintenance engineer’s job safer and easier. System installations are designed to operate under all aircraft attitudes and flight loads. Systems need to pass through holes in frames and bulkheads. Fluid hoses and tubes are supported where they pass through the structure. Grommets made from nylon or rubber compounds protect the components, moving control systems and structure from damages and wear caused by fretting. Hydraulic Systems Most transport category aircraft have hydraulic systems located in dedicated panels. These panels are commonly located in an unpressurised zone of the aircraft. Fluid leakage will drain overboard and not contaminate the cabin. Having the components grouped together simplifies servicing and maintenance. Some aircraft do have hydraulic systems components located within the pressure cell. E.g. F28/F100 and BAE146. In these aircraft, fluid leak tends to pool under the cabin floor. Hydraulic fluid misting from a small leak under very high pressure (3000 psi) has been known to enter the air conditioning system to the discomfort of the passengers and crew. For this reason most hydraulic components are eliminated from the pressurised area. Air Conditioning System Air conditioning systems are also located in an unpressurised zone The cabin air distribution system includes air ducts, filters, heat exchangers, silencers, non-return (check) valves, humidifiers, and mass flow control sensors, and mass flow meters. Components have lugs and stays which attach to brackets mounted on the aircraft structure. Electrical/Avionics Avionics and electrical equipment centres are commonly located in the pressurised cell of the aircraft. Some of these systems have their own cooling and temperature monitoring systems. Equipment centres are usually located below the cockpit or cabin floor and some equipment centres can be accessed from the cabin. However, entry via the cabin is not intended for flight crew access in flight, normal access is through an external door. The battery is stowed in a compartment designed to support its weight in all aircraft attitudes and under high “g” loads. The inside of the compartment should be adequately protected with a tar- based paint or with polyurethane enamel to protect against corrosion. Battery cables should be adequately supported by clips to protect them from chafing and flexing. All hardware in the battery compartment should be corrosion resistant. 2015-08-24 B2-13.2 Structures – General Concepts Page 35 of 42 Training Material Only Fuel Storage In small aircraft the fuel tank or tanks are located near the centre of gravity so the balance changes very little as the fuel is used. In large aircraft, fuel tanks are installed in every available location and fuel valves allow the flight crew to keep the aircraft balanced by scheduling the use of the fuel from the various tanks. The weight of the fuel is a large percentage of an aircraft’s total weight, and the balance of the aircraft in flight changes as the fuel is used. These conditions add to the complexity of the design of an aircraft fuel system. Figure 34: Fuel Tank System Main Landing Gear (MLG) Fittings for the wing mounted MLG are installed on the aft face of the rear spar and an auxiliary or false spar. These fittings include: The gear support rib The trunnion fitting The side stay attachment bracket The actuator attachment bracket Figure 35: Main Landing Gear Fittings 2015-08-24 B2-13.2 Structures – General Concepts Page 36 of 42 Training Material Only Waste Disposal The sub-systems of the waste disposal system which discard waste products and fluids from the galleys and the lavatories are: The Toilet System The Waste Water Drain System Figure 36: A Waste Water Drain System The toilet system removes waste from the toilet bowl through a vacuum drain to an underfloor waste holding tank. The system uses potable water from the aircraft pressurised water system to flush the toilet. During ground service, the waste holding tank is emptied, cleaned and filled with a prescribed quantity of sanitary fluid. The waste water drain system discards waste water from the lavatory washbasins and the sinks of the galley through the aircrafts heated drain masts. Waste storage tanks are located in rear of the aircraft in the pressurised zone. The support structure must support the heavy full tanks at all aircraft attitudes and during high ‘g’ manoeuvres. 2015-08-24 B2-13.2 Structures – General Concepts Page 37 of 42 Training Material Only Potable Water The aircraft potable water system is another heavy storage system which supplies drinking water to the galleys and the lavatory wash basins. The potable water storage tanks are located in various locations within an aircraft’s pressurised zone. The support structure must support the weight of the full tanks at all aircraft attitudes and during high ‘g’ manoeuvres. Figure 37: Potable Water System Heating is required to prevent the water freezing where plumbing passes through unpressurised areas. Galley and wash basin waste water is drained overboard via heated drain masts. Auxiliary Power Unit (APU) Each engine and auxiliary power unit mount and its supporting structure must be designed for a limit load factor in lateral direction. For the side load on the engine and auxiliary power unit mount must be at least equal to the maximum load factor obtained in the yawing conditions but not less than 1.33. 2015-08-24 B2-13.2 Structures – General Concepts Page 38 of 42 Training Material Only Lightning Strikes Aircraft cannot always avoid being struck by lightning. To minimise the potential danger of a lightning strike, principally arcing and electrical shock, the electrical current created must be able to find a path of least resistance from the point of impact to a suitable discharge point. Figure 38: Lightning Strike on B747 This path is normally created by the metal skin and frame of the aircraft, provided that the individual structural components that make up the aircraft are adequately connected electrically. Composite materials, which are not electrically conductive, must have provision to conduct electrical current incorporated into their design and manufacture. To provide a suitable path to dissipate a lightning strike safely, the following procedures are employed: High capacity electrical conductors, called bonding straps, link together all parts of the airframe to provide a low resistance path, thereby reducing the possibility of arcing. Figure 39: Bonding Strap attached to flight control 2015-08-24 B2-13.2 Structures – General Concepts Page 39 of 42 Training Material Only The radome, which must be manufactured of a non-conductive composite material, has conductive lightning diverter strips incorporated to collect the charge built up on the radome surface and safely transfer it to the airframe structure. Figure 40: Radome lightning diverter strips (F28) Static electricity or a lightning strike will generally discharge from the sharp trailing edges of the airframe structure. To prevent arcing damage to these edges, static wicks are placed at these locations to harmlessly dissipate the electrical charge to the atmosphere. Figure 41: Static Discharge Wicks 2015-08-24 B2-13.2 Structures – General Concepts Page 40 of 42 Training Material Only Aircraft Electrical Bonding Bonding Bonding is the electrical connecting of components of an aircraft structure together which are not otherwise adequately connected. This is to prevent static electricity (including lightning) from building up on one part of the structure to a point where it is high enough to allow it to jump to another part. Bonding failure has the potential to cause Electromagnetic interference (EMI), physical damage to mechanical and electronic components, and risk of electrical shock. Grounding The term grounding, when used in reference to aircraft electrical circuits, is where conductive parts of the structure are used to provide a return path, instead of using an insulated wire, for completing an electrical circuit normally. Grounding also referred to as earthing, provides an alternative return path should a two wire electrical circuit becomes faulty, reducing the potential for electric shock. Bonding ensures there is a continuous path between different parts of a conductive structure. On an aircraft, the primary structure is commonly referred to as ground. Bonding Leads To accomplish the purpose of grounding, it is necessary to provide a conductive path where direct electrical contact does not exist. Jumpers or bonding straps used for this purpose in such applications as between moving parts, between shock-mounted equipment and structure, and between electrically conducting objects and structure. Bonding straps (or leads) are pre-fabricated from braided copper or aluminium terminated with crimps. The maximum permissible resistance of a bonding strap is 0.003 ohms. Figure 42: Radome Bonding Straps (F28) 2015-08-24 B2-13.2 Structures – General Concepts Page 41 of 42 Training Material Only This page is intentionally blank 2015-08-24 B2-13.2 Structures – General Concepts Page 42 of 42 Training Material Only