22 AERO Second-Quarter 2003 — April

MAINTENANCE

OF HIGH-STRENGTH COMPONENTS

Many landing gear, flap supporting, and flapactuating components on Boeing airplanes aremade of high-strength alloy steels. Operationaladvantages are realized when these high-strength, high-heat-treated materials are used inlimited-space envelopes. To reap the benefits ofhigh-strength alloy steel components and avoidpotential safety issues resulting from damage,airline maintenance and overhaul personnel needto follow proper maintenance procedures andrework practices, checklists, and guidelines during component maintenance and overhaul.

M A I N T E N A N C E RALPH M. GARBER

ASSOCIATE TECHNICAL FELLOW

COMMERCIAL AVIATION SERVICES

BOEING COMMERCIAL AIRPLANES

CRAIG DICKERSON

LEAD METALLURGICAL ENGINEER

BOEING MATERIALS TECHNOLOGY

BOEING COMMERCIAL AIRPLANES

AERO 23Second-Quarter 2003 — April

ALLOY STEEL

24 AERO Second-Quarter 2003 — AprilAERO

operators achieve the benefits asso-ciated with high-strength alloy steelsand avoid potential safety issuesresulting from damage caused bystress concentrations, detrimental sur-face conditions, corrosion, improperprocessing, or other factors.

This article discusses some

factors that cause damage in service or during overhaul. Most can beattributed to a lack of familiarity with high-strength alloy steels.Operators usually recognize the benefits of using these steels; how-

ever, certain characteristics of thesteels are not always given properconsideration during componentmaintenance or overhaul. Thesecharacteristics, including sensi-tivity to corrosion pitting, sus-ceptibility to microstructural damage resulting from embrittle-ment, and notch sensitivity, canlead to rapid crack growth in some load environments.

This article describes

1. Benefits of high-strength alloy steel.

2. Importance of proper inspection and rework.

3. Guidelines for reworking high-strength alloy steel components.

any landing gear, flaptrack, flap carriage, and

other flap actuating components on Boeing airplanes are made ofhigh-strength alloy steels, such as 300M, Hy-Tuf, 4340M, and 4330M.These components provide structural benefits (e.g., reliable, durable design)and strength characteristics that permit an efficient use of availableairframe space. Other steels inuse, including 9Ni-4Co-0.3C,AerMet 100, and precipitation-hardened stainless steels, havesimilar maintenance and over-haul requirements. (Note: High-strength alloy steels referencedin this article generally havebeen heat-treated above 180 ksi[180,000 psi]; most have beenheat-treated above 220 ksi.)

Airline personnel should follow proper maintenance procedures and Boeing-providedrework practices, checklists,and planning guidelines duringmaintenance and overhaul ofthese components. This will help

M

AERO 25Second-Quarter 2003 — April AERO

BENEFITS OF HIGH-STRENGTHALLOY STEEL

Components made of high-strengthalloy steel generally weigh less andrequire less space to house than compo-nents made of lower strength alloys.Using high-strength alloy steel for com-ponent design provides an opportunityto do the same job with less material.When properly maintained and overhauled, high-strength alloy steelcomponents demonstrate high levels ofservice reliability.

The decision to use high-strengthalloy steels is based on weight and eco-nomic factors. Airframe space for gearcomponents may be reduced because of smaller diameter shock strut com-ponents, smaller pins (reduced space for joints), smaller diameter trucks andaxles, and, in some instances, smallerdrag brace, side brace, and attach fit-tings. By reducing the space requiredfor these components, the wheel wellsize can be minimized and aerodynamicsurfaces can be optimized, which allowan increase in fuel tank size (optimalwing spar location) or additional spacefor other uses.

The use of high-strength alloy steelparts is economical because it reducesweight, thereby allowing for more efficient aerodynamic surfaces and providing the potential for increasedpayload and fuel.

For example, the trailing edge of the wing is relatively shallow. Usinghigh-strength alloy steel flap tracks,flap carriages, and flap actuating com-ponents reduces the profile and decreasesspatial envelope requirements whilemeeting or improving aerodynamicrequirements. This also optimizes wingshape and reduces the potential need forbulging aerodynamic surfaces, which in turn reduces drag and increasesairplane performance.

IMPORTANCE OF PROPERINSPECTION AND REWORK

Following proper rework practices and using Boeing-provided documentsduring maintenance and overhaul arenecessary to achieve the benefits associated with high-strength alloy steel components and help ensure safeairplane operation.

Airline personnel who participate incomponent rework, maintenance, andoverhaul tasks should be familiar withthe properties of high-strength steelsand understand the negative effects thatcan result from■ Sensitivity to stress concentrations

(notch sensitivity).

■ Microstructural damage from embrittlement or overheating.

■ Detrimental surface conditions.

■ Corrosion.

■ Improper processing.

Improper rework practices can resultin unscheduled maintenance or surfacedamage that causes crack initiation.Maintenance efforts focus on corrosionprevention and removal in addition tonormal checks for wear and free play.

High-strength alloy steels can ex-perience rapid crack propagation fromstress corrosion under certain loadingconditions. Therefore, surface damagedetection is important during overhauland on components in service.Removing visible surface corrosionbefore pitting begins (such as during a C-check) helps prevent conditions that can lead to crack initiation. Thebest safeguard against corrosion is to ensure that finishes conform to thedesign and that design improvementsare incorporated as minor changeswhenever possible.

Components manufactured from steel alloys heat-treated above 180 ksi(180,000 psi) should be reworked inaccordance with guidelines inComponent Maintenance Manuals(CMM) 32-00-05, 32-00-06, and

32-00-07. Although these guidelinesapply directly to landing gear com-ponents, they can be used to plan theoverhaul rework of all high-strengthsteel components. Standard OverhaulPractices Manual (SOPM) 20-10-01generally is specified in each CMMsection for the rework of wing compo-nents (e.g., flap tracks, flap carriages).For repair of high-strength, 300M steelparts on DC-10 and MD-11 airplanes,use CMM 20-11-02; for DC-9,MD-80, MD-90, and 717 airplanes,use CMMs 20-10-18 and 20-10-06.

In addition, airline personnel need to understand the importance of main-taining component finishes while in service (in situ, or on the airplane).This includes repairing damaged finishes to prevent corrosion and en-suring that solvents and materials thatcome in contact with the finishes donot result in premature degradation and unscheduled component removal.

Boeing documentation describes the methods for detecting base metaldamage while in service and duringoverhaul. Common techniques includedetailed visual inspections and othernondestructive inspection methods,such as magnetic particle inspection(MPI) and fluorescent penetrant inspec-tion (FPI). (See SOPMs 20-20-01 and20-20-02.) Ultrasonic or eddy currentinspections also may be useful for in situ inspections.

Boeing also is developing sup-plemental, specialized techniques,such as the Barkhausen inspection, todetect base metal heat damage underchrome plating or other protective finishes. This technique can be used successfully to screen components withsuspect damage. For example, if an axlefractures as a result of chrome-grindingheat damage during manufacture oroverhaul, the Barkhausen inspectionallows other suspect components to bescreened without first performing achrome strip and temper etch (e.g., nitaletch) inspection on all suspect axles.

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26 AERO Second-Quarter 2003 — April26 AERO

GUIDELINES FOR REWORKINGHIGH-STRENGTH ALLOY STEELCOMPONENTS

This section provides guidelines forreworking high-strength alloy steelcomponents and describes some of the implications of improper reworkprocedures.■ Stress concentrations.

■ Overheating components.

■ Hydrogen embrittlement.

■ Cadmium embrittlement.

■ Improper finishing.

STRESS CONCENTRATIONS

During component design, eliminatingor minimizing areas of stress concentra-tions is a key objective. Special atten-tion is given to protective finish runoutsadjacent to stress concentration details.In addition, all stress concentrationdetails are subject to extensive testingand/or analysis to ensure that no detrimental effects are introduced into a part. Any rework or repair must notincrease stress concentrations thatdegrade component durability.

High-strength alloy steel components(along with those made from othermaterials) are shot-peened to create ashallow layer of compressive residualstress at the surface. This layer helps to■ Minimize the effects of stress

concentrations in transition areas.

■ Impede crack initiation and initialcrack growth caused by fatigue orstress corrosion.

■ Create a surface that will have minimal adverse effects from theresidual stresses of plating.

When a surface is machined orground to remove damage, thereworked area should be shot-peenedwith proper overlap onto the existingshot-peened surface. During overhaul,personnel must observe the platingrunouts specified in the CMM sectionsand SOPMs 20-10-01 and 20-42-03.

For example, when a coating such as chrome or nickel plating is applied

to surfaces to prevent wear or corrosion,the coating must exhibit proper runoutsthat terminate before the tangent of fil-let radii, edges, or other shape changes.Boeing SOPM guidelines should be fol-lowed for the rework of any componentand for all types of plating or coating.Rework or overhaul of componentsshould not introduce stress concentra-tions, or otherwise increase stresses,which can reduce the service life of acomponent below that of the originaldesign configuration.

Stress concentrations can lead to initiation of cracking by fatigue, stresscorrosion, or hydrogen-assisted stress corrosion. These cracks mayresult in a fracture or scrap of a com-ponent when found while in service or during overhaul. The following areexamples of stress concentrations thatcan lead to cracking.

Transitions or radii that are sharperthan original design. When removingdamaged material from part surfacesduring rework, the new transitions orradii should not cause an unacceptableincrease in stress concentration at thelocation or degrade the original designfeatures. When locally machining outcorrosion or damage during overhaul,a gradual transition into the reworkeddepression is necessary.

The intent is to remove the leastamount of material possible whileensuring that all discrepant material isremoved and the original designstrength and durability are maintained.There are few options to restore thesemachined depressions to meet interfacerequirements. One type of rework oroverhaul, sulfamate-nickel plating, iscommon on shock strut cylinder diame-ters and is used to repair lug faces todesign dimensions as follows:■ Local blends on inner cylinder outer

diameter surfaces and outer cylinderinner diameter surfaces often arefilled with sulfamate-nickel platingto restore them to dimensions thatare suitable for subsequent chromeplate application.

■ Spot facing on lugs is controlled tohave a generous radius at the tran-

sition to the adjacent surface andusually is kept at the minimum depthnecessary to clean up the damagedsurface. Spot face depressions typi-cally are not filled with plating torestore the dimension but instead arefinished in the same manner as theoriginal design. Spot face transitionradii need to be such that they can be shot-peened to the requirements of the adjacent surfaces.

■ When the entire face of a lug mustbe machined to remove damage, thenew lug transition radii should beshaped and positioned in accordancewith CMM requirements. Surfacetransitions into the lug hole and atthe lug edges must have design tran-sitions that will allow restoration ofshot-peening on all reworked areasand permit complete seating of bush-ings without contacting hole edges.

Abrupt changes in sections, holes,and sharp-cornered keyways should be avoided. Proper design will reflect generous fillets, gradual changes ofshape, and the use of relief grooves inareas of high stress. Finer surface fin-ishes also may be needed to eliminateunnecessary stress concentrations,especially in areas of machined radii or undercuts. Overhaul should reflectthe same careful, detailed review thatoccurred during the original design.

Plating conditions and runout con-trols that are not in accordance withdesign standards. During overhaul,many landing gear components arecompletely stripped to replace nickeland chrome plating. In most instances,these repairs involve rework of the base metal. The new plating depositsfrequently are thicker than the originaldesign configuration.

In all cases, it is important to adhereto the SOPM recommendations. Thiswill ensure that the restored plating is of high quality and that it does notterminate with an abrupt edge. Through-thickness cracks in chrome plate (gen-erally present where there is evidence of chicken-wire cracking) can lead to corrosion at the base metal interface

3

AERO 27Second-Quarter 2003 — April

and deterioration of the plating ad-hesion. Through-thickness crackingalso can lead to fatigue or stress corrosion cracking of the base metalbeneath the plating.

Visual evidence of chicken-wirecracking after chrome grinding indi-cates poor chrome quality and also mayindicate the possibility of base metal heatdamage. Chicken-wire cracking notedin SOPM 20-10-04 indicates that thechrome should be stripped and replated.

If the plating runouts are blended ormachined to remove the abrupt platingedge, the techniques must be well con-trolled to avoid damaging the adjacentbase metal. Improper blending canremove the required shot-peened layeror create undercuts or grooves at theedge of the plating that can causecracking in service.

Several in-service fractures havebeen attributed to improper platingtechnique, poor-quality plating,improper runout conditions, and basemetal damage caused by poor blendingor machining control.

Proper use of special plating tech-niques, such as conforming anodes androbbers, can control plating thicknessesand runouts. This can reduce the possi-bility of chrome chicken-wire crackingand poor runout details.

Plating into a transition (radius transition or undercut) will create astress concentration that can causecrack initiation. For example, figure 1shows an outer cylinder clevis platedinto the lug transition. In service,fatigue cracking initiated

at the platingrunout led tolug fracture.

Corrosion and pitting.Corrosion pitsare stress con-centrations. Asthe pit forms,it damages theshot-peenedlayer locally at the surface.The pit thengrows throughthe compressivelayer, and thechange in resid-ual stress stateand the pitgeometry initiatestress corrosioncracking. This type of cracking mostoften occurs on surfaces that are bothprone to corrosion and exposed to sus-tained tensile stresses while in service,such as the lower surface of landing geartrucks, axles, and the surfaces of forwardand aft trunnions.

Corrosion pitting also can lead tofatigue crack initiation depending onthe component, the location of pitting,and cyclic loading conditions. In thesecases, the cracks can propagate to thecritical length and result in ductile fracture of the component. The degreeof cracking tolerated before fracturevaries by component, crack location,and component loading conditions.

To prevent excessive corrosion, thor-ough visual inspections should be per-formed on a regular basis to evaluatethe condition of the protective finishes.Damage should be repaired soon afterit is found. Touching up damage to ac-cessible enamel and primer in a timelymanner can prevent the formation ofcorrosion pits and reduce the need forexcessive rework during overhaul.Rework that requires low-hydrogen-embrittlement (LHE) cadmium stylusplating should be performed when thecomponent is not loaded.

When the component is removed

CRACK INITIATION AT CHROMEPLATE RUNOUT INTO RADIUS

FIGURE

1

FINISH DAMAGE ALONG LOWER ID SURFACE AND AREAS OF CORROSION PITTING AT FRACTURE ORIGIN

FIGURE

2

for overhaul, all evidence of corrosionmust be removed and finishes restored to design requirements or better. The sequence of rework operations is provided in CMMs 32-00-05, 32-00-06,and 32-00-07.

Landing gear truck fractures haveoccurred in service because of corro-sion on the inner diameter of the maingear truck beam (figs. 2 and 3). Thesefractures may be caused by a combina-tion of degraded protective finishes onthe truck inner diameter, poor drainage,and contact with the corrosive chemi-cals in washing solutions or deicingcompounds. Truck fractures most oftenoccur at maximum ground loads suchas after fueling or during preflight taxi.

Figures 4 and 5 show a drag bracefrom which corrosion was not removedcompletely during overhaul. The partwas subsequently shot-peened, and new protective finishes were appliedover the residual active corrosion. This resulted in crack initiation andpropagation while in service and theeventual fracture of the component.

Mechanical damage. Stress concentrations can be created bymechanical damage that compromisesthe protective finishes and alters thecompressive shot-peen layer. This

AERO28 Second-Quarter 2003 — April

damage often is caused by impropermaintenance practices such as jackingadjacent to a jack pad or an inadvertentimpact with tools or ground-supportequipment (e.g., tow vehicles).Although high-strength alloy steels arehard and resist dents, scratches, andnicks, stress concentrations caused bymechanical damage can dramaticallyreduce the service life of a component.

High-strength alloy steel componentsalso can be damaged by mishandlingduring shop rework (e.g., dropping,impact), and in some circumstances, byforeign object debris. Possible mechani-cal damage to a high-strength alloysteel component should be evaluated bythe operator and repaired as needed.

If the damage is local and widespreaddeformations are not evident, repair maybe similar to that required for corrosionand pitting. All deformed material mustbe removed before refinishing; deformedhigh-strength steel alloy componentsmust not be straightened. ContactBoeing for assistance, if needed.

LOWER PORTION OF FRACTURED DRAG BRACE

FIGURE

4

CROSS-SECTION VIEW OF RESIDUAL CORROSION UNDER CADMIUM PLATING

FIGURE

5

CLOSE-UP VIEW OF CORROSION PITTING AT FRACTURE ORIGIN

FIGURE

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AERO 29Second-Quarter 2003 — April

OVERHEATING COMPONENTS

Overheating of components can changethe original steel temper and mechani-cal properties of the affected area.Overheating damage can be caused by■ Frictional heating while in service.

■ Abusive machining and grindingoperations during manufacture or overhaul.

■ Exposure to high temperatures during over-haul bake cycles.

■ Unusual conditions such as refused takeoffs and local fires.

The degree to which themechanical properties arechanged depends on the tempera-ture and duration of exposure.

Overheating can result inovertempered martensite (OTM) or untempered martensite (UTM)formations in the base metal.Both conditions can be detectedby a temper etch (i.e., nital etch)inspection of the base metal.UTM indications show whiteduring temper etch inspectionsand often are found withinpatches of OTM, which showdark gray to black during temperetch inspection. SOPM 20-10-02provides details about theinspection process and inter-pretation of the results.

Heat damage generally isremoved by carefully machin-ing the base metal. Afterward,another temper etch inspectionis done to ensure that themachining did not create moreheat damage.

UTM formations may be accompanied by heat-induced crack-ing within these overheated areas that,if left in place, can propagate while in service. Figures 6 and 7 show service-induced heat damage on theinside diameter of a main gear outercylinder. This component developedextensive frictional heat damage in the upper bearing contact area

as a result of improper clamp-up. The heat damage led to crackingthrough the cylinder wall. Salvage wasnot possible.

Less severe friction-induced heatdamage can be found on inner cylindersduring component overhaul. This dam-age, which occurs on a more frequentbasis, is caused by vertical motionagainst the lower bearing surfaces. This

CRACKED MAIN GEAR OUTER CYLINDER WALL

FIGURE

6 damage generally is shallow and can beremoved by machining. After overhauloperations are completed, the compo-nent is returned to service in accor-dance with CMM requirements.

When grinding chrome to finishdimensions, overheating the base metalcan create UTM and OTM formationsunder the chrome. Figures 8 and 9show a severe grinding burn on a mainlanding gear axle that resulted in a

fracture. Similar grinding burnsalso have led to the fracture offlap carriage spindle journals(figs. 10 and 11).

Any visible evidence ofchrome plate distress canindicate the likelihood of basemetal heat damage. Figures 12and 13 show a grinding burnthat led to the fracture of apivot pin. SOPM 20-10-04 andCMMs 32-00-05, 32-00-06,and 32-00-07 provide guide-lines that indicate whenchrome must be removed during overhaul.

Some heat damage is sosevere that the heat-treat con-dition of material is altered inadjacent areas. This widespreadreduction in metal hardness(Rockwell-C hardness readings)may indicate that the compo-nent cannot be salvaged. Axleheat damage caused by a wheelbearing fracture may lead tosuch a condition.

Shop procedures such as magnetic particle inspectionand LHE cadmium stylusplating can cause arc burns ifappropriate precautions are notmaintained during processing.Figures 14 and 15 show a

fracture resulting from an arc burnthat developed during LHE stylus cad-mium plating. (Note: In this article,cadmium plating means cadmium-titanium or LHE cadmium plating.)

Overheating will not alter the heat-treat conditions of the base metal if the temperatures are below the originaltempering temperature. However, thecomponent still may require special

NITAL ETCH INDICATIONS OF HEAT DAMAGEON ID OF MAIN GEAR OUTER CYLINDER

FIGURE

7

30 AERO Second-Quarter 2003 — April

consideration (or rework) because ■ Shot-peening may be compromised

(heated above 400°F).

■ Cadmium embrittlement may occur(heated above 450°F with cadmiumplating present).

■ Chromate conversion coating maybe degraded (heated above 400°F).

■ Organic coatings or sealants maycrack or become brittle or discolored(wide range of temperatures).

These situations often occur whencomponents are ■ Inadvertently overheated in an oven.

■ Exposed to elevated temperatureswith some finishes intact or bushings installed.

■ Exposed to fire.

Residual cadmium often is left on a part during overhaul processing toprotect it from corrosion. The part isthen stripped of all cadmium and re-plated near the end of overhaul. Partswith residual cadmium should not beheated over 400°F during overhaul.

Bushings should not remain installedduring overhaul unless retained byspecific CMM requirements. Bushingsmust be removed to permit a thoroughinspection of the base metal and toavoid bushing-to-bore interface degra-dation during bake cycles. Design fin-ishes are restored and new bushingswith design interferences and dimen-sions are installed because bushing wearlimits do not apply during overhaul.

Wheel bearing fractures or high-energy refused takeoffs often result inhigh local heat on an axle. Discoloration

of the enamel, primer, orchrome or evidence of cad-mium damage on the innerdiameter of the axle mayrequire the heat-damagedcomponent be removedfrom service.

Overheating affectscomponents to variousdegrees; in some instances,only finish durability isdegraded. This may resultin a shorter than plannedtime between componentoverhauls. Contact Boeingfor assistance with ques-tions about repairing or salvaging high-strength alloy steel componentsthat appear to have been damaged byoverheating.

HYDROGEN EMBRITTLEMENT

Hydrogen embrittlement occurs when a high-strength alloy steel component absorbs hydrogen, whichis not removed in a timely manner in accordance with the SOPM (e.g., embrittlement relief baking).

When hydrogen remains in a component for an extended time, themicrostructural damage that developssignificantly degrades the mechanicalproperties of the steel. The infusedhydrogen migrates to areas of highstress (e.g., material internal stresses)and creates local microstructural dam-age. When the component is installedon an airplane, this internal damagecan lead to crack initiation and propa-gation, resulting in component fracture.

The elevated temperatures reachedduring hydrogen embrittlement relief

baking, which is performed directlyafter stripping or plating operations dur-ing overhaul, effectively remove hydro-gen generated during these operations.Processes that must be followed withrelief baking include chrome, sulfamate-nickel, and LHE cadmium plating; strip-ping operations; and many nital etchinspections. After hydrogen-generatingoperations, relief bake delay time limitsmust be observed to ensure completehydrogen removal. In general, the bestpractice is to initiate baking as soon aspossible following a plating operation.

The delay time between platingcompletion and baking start typically isobserved. However, when thick platingdeposits or multiple plating operationsare performed on a single component,the total time between initial platingstart and baking start is a key factorwhen determining the maximum delaytime allowed. For example, embrittle-ment relief baking must begin 10 hrafter sulfamate-nickel plating is com-pleted or within 24 hr after plating

FRACTURED MAIN LANDING GEAR AXLE

FIGURE

8 NITAL ETCH INDICATIONS OF HEAT DAMAGE ON OD OF AXLE AFTER CHROME REMOVED

FIGURE

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AERO 31Second-Quarter 2003 — April

begins, whichever results in the shortestoverall bake delay.

Figure 16 shows a flap track thatcracked because of hydrogen embrittle-ment 149 flight cycles after overhaul.Figure 17 is a scanning electron

FRACTURED PIVOT PIN

FIGURE

12

FRACTURE THROUGH FLAP CARRIAGE SPINDLE

FIGURE

10 GRINDING DAMAGE VISIBLE ON CHROME PLATE AT FRACTURE ORIGIN

FIGURE

11

CLOSE-UP VIEW OF PIVOT PIN OD AT FRACTURE ORIGIN

FIGURE

13

microscope view of a typical hydrogenembrittlement crack where separationoccurs along grain boundaries.Typically, hydrogen embrittlementcracks propagate rapidly once loads are applied to the part. In some cases,internal residual stresses are sufficient-ly high to cause cracking even before the part is installed.

CADMIUM EMBRITTLEMENT

Overheating LHE cadmium orcadmium-titanium plated componentscauses embrittlement of high-strengthalloy steel by cadmium, resulting in cadmium diffusion into the steel grain boundaries. Solid-metal

embrittlement by cadmium can occur at temperatures below the cadmiummelting point. These effects on thebase metal can begin to occur at 450°F,whereas the cadmium melting point is generally 610°F. The microstructuralanomalies resulting from cadmiumembrittlement can lead to componentfractures in service.

Determining whether cadmium hasmigrated into the grain boundaries ofcadmium-plated, high-strength alloysteel components requires destructivetesting of the components. If thesecomponents have been overheated,salvage may not be possible. However,if high-temperature exposure was

short and discoloration of the enamelor primer was minimal, the componentmay be a candidate for salvage. Slightor no discoloration of the enamel orprimer may indicate the cadmium plating was not heated to the extentthat cadmium embrittlement would be suspected. Boeing can assist in this determination.

32 AERO Second-Quarter 2003 — April

IMPROPER FINISHING

Improper applicationof protective finishesduring manufacture or overhaul can lead to finish degradation,corrosion, and cor-rosion pitting, which can result in com-ponent fracture whilein service (figs. 2 and 3, pp. 27–28).Some cleaners andchemicals may accelerate finishdegradation and lead to corrosion.Operators should ensure that cleaners and chemicals are testedbefore use in accordance with Boeingdocument D6-17487, Evaluation of Airplane Maintenance Materials.Testing to these requirements willdetermine whether a cleaner or chemical is detrimental to protectivefinishes or base metal. However,long-term exposure to the solution or material still may adversely affect finishes.

Personnel must ensure that materialsused for activities such as cleaning anddeicing conform to Boeing documentD6-17487 requirements and willaccomplish the intended task (verifiedby the material provider or operator).Refer to the Aircraft MaintenanceManual for materials specified for air-craft cleaning and deicing. The CMMspecifies the materials for use in repair.

High-strength alloy steel compo-nents should be stripped completely

during overhaul (including removal ofbushings and bearings in all structuralcomponents). This allows a thoroughinspection of the base metal (a primarycomponent overhaul requirement) andensures that all finishes, including theLHE cadmium plating and conversioncoating, are restored to the original design requirements. This is addressedin an all-model Boeing service letterdated April 23, 2002, Overhaul of High Strength SteelComponents–Cadmium Strip

Required (e.g.,757-SL-20-036-A,767-SL-20-038-A,747-SL-20-062-A).

Restoration ofthe shot-peenedlayer during over-haul is importantto ensure that theshot-peen com-pressive residualstresses are main-tained or restored.Removing or

damaging the shot-peened layer canreduce the protection that this com-pressive layer provides against fatigueand stress corrosion crack initiation.Discontinuous shot-peening can leadto crack initiation at the tensile surfacestresses adjacent to edges of abruptcompressive layer runouts (no fade-out). All reworked surfaces must beshot-peened after removing materialdamaged by corrosion, heat, anddeformation.

As a rule, if material removal exceeds0.0015 in (or 10 percent of the Almenstrip intensity), the surface should thenbe shot-peened to CMM requirements.Exceeding shot-peen requirements isbetter than leaving areas without shot-peening. All portions of a componentthat are to be shot-peened should first be completely stripped; no cadmiumresidue should remain on the surface.

FRACTURED ACTUATOR BEAM

FIGURE

14

NITAL ETCH INDICATIONS OF ARC BURNHEAT DAMAGE THAT LEAD TO FRACTURE

FIGURE

15

VISIBLE CRACKS IN FLAP TRACK (ARROWS)

FIGURE

16

INTERGRANULAR FRACTURE CAUSEDBY HYDROGEN EMBRITTLEMENT

FIGURE

17

AERO 33Second-Quarter 2003 — April

High-strength alloy steels are used

widely in landing gear, flap track, flap

support carriage, and flap actuating

components on Boeing airplanes. These

high-strength materials provide signifi-

cant structural benefits and can result in

a weight savings. These parts often are

selected for placement in limited-space

envelopes (e.g., wheel wells and wing

trailing-edge support structures)

because of their reduced profile or

smaller diameters.

With these benefits comes a need

for airline personnel to exercise

precise care when reworking high-

strength alloy steel components

during scheduled maintenance and

overhaul. They need to understand the

importance of maintaining component

finishes while in service, follow

proper rework practices, and use

Boeing-provided maintenance

procedures, planning guidelines, and

checklists during scheduled main-

tenance and overhaul processes.

Improper rework and overhaul

practices may result in loss of finish,

corrosion, and damage to or alteration

of the base metal, which may require

unscheduled maintenance between

overhauls. The resulting damage also

could precipitate crack initiation and

removal of the part from service.

Removing corrosion and restoring worn

interfaces on a periodic basis are the

main emphases of high-strength alloy

steel component overhaul rework.

Key benefits of proper rework

and maintenance practices include the

possibility of extending the gear or

component overhaul intervals (time

between overhaul). Operators also will

benefit from the enhanced reliability

and durability of high-strength alloy

steel components on their airplanes.

Operators should ensure that

proper SOPM and CMM documen-

tation is used during overhaul and

rework of high-strength alloy steel

components. The planning flowcharts

in CMMs 32-00-05, 32-00-06, and

32-00-07 are value-added guidelines

for planning the rework of any

high-strength alloy steel component

on a Boeing airplane.

Editor’s note: The SOPMs and CMMs identified in this article can be ordered through the Data and Services Catalog.

S U M M A R Y

34 AERO Second-Quarter 2003 — April

About the Authors

Craig Dickerson has been a metallurgical engineer in the aero-space industry for 18 years. As leadengineer for the Boeing MaterialsTechnology Landing Gear DesignCenter support group and former leadengineer for the Fracture Analysisgroup, he is involved with all aspectsof landing gear structure materials

and processes, including detail part manufacture, in-serviceperformance, component overhaul, analysis of parts returnedfrom service, and accident and incident investigations.

Ralph M. (Mike)Garber has been astructures engineer inthe aircraft industry for 38 years, including28 years as lead engineer and the past 7 years as an AssociateTechnical Fellow. His primary responsibilitieshave involved design and certification analysis of the wing, empennage, nacelle struts, and landing gear with a focus on airline workshops and identifying design improvements toincrease durability and fatigue resistance. He is a licensed professional engineer and hasassisted in NTSB investigations and Boeing- and FAA-recommended structural changes.

identified for both steel and titaniumalloy components in production and as a substitute for chrome plating duringcomponent overhaul. Some repair agencies and airlines are purchasingequipment and preparing facilities forHVOF coating application during over-haul as an alternative to chrome plating.

MAINTAIN FINISH DURABILITY THROUGH PROPER WASHING,CLEANING, AND FREQUENTRELUBRICATIONProperly restoring the finish of high-strength alloy steel components to origi-nal design conditions during overhaulminimizes the effects of washing andcleaning solutions, solvents, and com-pounds on the structure. Design-qualityfinishes are less likely to degrade inservice. In addition, frequent relubri-cation of these components soon afterwashing protects finishes at lubricatedinterfaces. Relubrication intervals arespecified in the Aircraft MaintenanceManual (AMM) but generally are ad-justed based on operator experience.

Boeing continues to receive reportsof premature corrosion from operatorsthat use pressure-washing techniqueson their airplanes. The following guide-lines will help operators maintain thefinishes of their high-strength alloysteel components through heightenedawareness and knowledge about keyaspects of airplane washing processes.

Washing and Cleaning Techniques■ Operators should avoid using

high-pressure washing.

■ When cleaning landing gear and othermechanical, electrical, or hydrauliccomponents, operators must followthe requirements in the AMM procedure Remove Material AroundSensitive Components (e.g., 747-400AMM 12-25-01, p. 301, 2.A.[2]).

■ After washing landing gear and con-trol surfaces, operators should com-plete rinsing within the specified timeperiod. Rinsing must not be delayed.

■ Operators should cover joints without relubrication fittings to

AERO 35Second-Quarter 2003 — April

avoid contamination from water and cleaning solutions.

■ If washing is done at the beginningof a scheduled maintenance period,operators should not wait until theend of the period to perform lubri-cation — the elapsed time may not be acceptable.

■ When flushing or rinsing landinggear assemblies, operators shouldreduce spray pressure, ensure thatthe nozzle is at least 12 in from thejoints, and replace the corrosion preventive compound after washing.(See the multimodel maintenancetip, “Airplane High PressureWashing,” May 18, 1999.)

Impact of Aggressive Washing onFinishes and Lubrication

■ Short-term exposure to materials thatnormally contact properly restoredfinishes, such as solvents, should notcause premature degradation or lossof finish requiring repair or unsched-uled removal between overhaul.However, premature corrosion anddeterioration can occur in servicewhen water or foreign material entersjoints as a result of spraying cleaningsolutions directly into joints. Thisaggressive washing technique dis-places grease and negatively affectslubricated joints even though immedi-ate relubrication will purge most con-taminates from the lubricated cavities.

■ Most corrosion-related cracking andfractures in service are aggravated byaggressive washing techniques andcorrosive solutions. To help ensurethat finishes do not degrade pre-maturely between overhaul, operatorsshould lubricate all greased bearingsand cavities no later than 12 hr afterairplane washing. Relubrication andreplacement of corrosion preventativecompounds within this time periodminimizes finish exposure to cor-rosive cleaning agents following airplane washing.

INCREASED USE OF TITANIUMALLOYS AND TUNGSTEN-CARBIDE COATINGSBoeing and the industry are workingtogether to develop thermal spray coat-ings to replace chrome plating. Thesecoatings, which currently are used insome production and repair applications,can be applied using the high-velocity oxygen fuel (HVOF) process. The twocoatings primarily in use are tungsten-carbide-cobalt and tungsten-carbide-cobalt-chrome.

These tungsten-carbide coatings canbe applied to steel and titanium alloys.Steels can be either chrome plated orcoated with tungsten-carbide. Titaniumalloys cannot be chrome plated but canbe coated with tungsten-carbide throughthe HVOF process. With this coating, thelighter weight titanium can be used inmore landing gear and flight controlapplications where high-strength alloysteels would have been used in the past,resulting in a weight savings.

Titanium alloys are being used moreoften in the design of new landing gearand flight control support structure.These materials, which are becomingmore readily available, exhibit higherstrength-to-weight ratios than do steelalloys. Titanium alloy componentsrequire less finishing, are more easilymaintained, are less prone to corrode inservice, and require less overhaul pro-cessing than most high-strength alloysteel components. The durable, HVOF-applied tungsten-carbide coatings broad-ens possible titanium alloy applications.

HVOF-applied tungsten-carbide coat-ings also provide multiple process bene-fits when compared with chrome plating:■ Embrittlement baking is not needed

after application because hydrogendoes not infuse into the base metal.

■ When grinding the coating to thedesired finish, overheating the basemetal is less likely.

■ Applying coatings using the properHVOF procedures and equipment cansave shop processing and flow time.

These benefits are driving more tungsten-carbide applications to be

High-pressure washing is detrimental and

should not be used under any circumstances.