MAINTENANCE HIGH-STRENGTH - Boeing

MAINTENANCE

OF HIGH-STRENGTH
COMPONENTS

Many landing gear, flap supporting, and flap
actuating components on Boeing airplanes are
made of high-strength alloy steels. Operational
advantages are realized when these high-
strength, high-heat-treated materials are used in
limited-space envelopes. To reap the benefits of
high-strength alloy steel components and avoid
potential safety issues resulting from damage,
airline maintenance and overhaul personnel need
to follow proper maintenance procedures and
rework practices, checklists, and guidelines
during component maintenance and overhaul.

MAINTENANCE RALPH M. GARBER CRAIG DICKERSON
ASSOCIATE TECHNICAL FELLOW LEAD METALLURGICAL ENGINEER
COMMERCIAL AVIATION SERVICES BOEING MATERIALS TECHNOLOGY
BOEING COMMERCIAL AIRPLANES BOEING COMMERCIAL AIRPLANES

22 AERO Second-Quarter 2003 — April

ALLOY STEEL

Second-Quarter 2003 — April AERO 23

M any landing gear, flap operators achieve the benefits asso- factors that cause damage in service
track, flap carriage, and ciated with high-strength alloy steels or during overhaul. Most can be
other flap actuating components and avoid potential safety issues attributed to a lack of familiarity
on Boeing airplanes are made of resulting from damage caused by with high-strength alloy steels.
high-strength alloy steels, such as stress concentrations, detrimental sur- Operators usually recognize the
300M, Hy-Tuf, 4340M, and 4330M. face conditions, corrosion, improper benefits of using these steels; how-
These components provide structural processing, or other factors.
benefits (e.g., reliable, durable design) ever, certain characteristics of the
and strength characteristics that This article discusses some steels are not always given proper
permit an efficient use of available consideration during component
airframe space. Other steels in maintenance or overhaul. These
use, including 9Ni-4Co-0.3C, characteristics, including sensi-
AerMet 100, and precipitation- tivity to corrosion pitting, sus-
hardened stainless steels, have ceptibility to microstructural
similar maintenance and over- damage resulting from embrittle-
haul requirements. (Note: High- ment, and notch sensitivity, can
strength alloy steels referenced lead to rapid crack growth in
in this article generally have some load environments.
been heat-treated above 180 ksi
[180,000 psi]; most have been This article describes
heat-treated above 220 ksi.)
1. Benefits of high-strength
Airline personnel should alloy steel.
follow proper maintenance
procedures and Boeing-provided 2. Importance of proper
rework practices, checklists, inspection and rework.
and planning guidelines during
maintenance and overhaul of 3. Guidelines for reworking
these components. This will help high-strength alloy steel
components.

24 AERO Second-Quarter 2003 — April

1 BENEFITS OF HIGH-STRENGTH 2 IMPORTANCE OF PROPER 32-00-07. Although these guidelines
ALLOY STEEL INSPECTION AND REWORK apply directly to landing gear com-
ponents, they can be used to plan the
Components made of high-strength Following proper rework practices overhaul rework of all high-strength
alloy steel generally weigh less and and using Boeing-provided documents steel components. Standard Overhaul
require less space to house than compo- during maintenance and overhaul are Practices Manual (SOPM) 20-10-01
nents made of lower strength alloys. necessary to achieve the benefits generally is specified in each CMM
Using high-strength alloy steel for com- associated with high-strength alloy section for the rework of wing compo-
ponent design provides an opportunity steel components and help ensure safe nents (e.g., flap tracks, flap carriages).
to do the same job with less material. airplane operation. For repair of high-strength, 300M steel
When properly maintained and parts on DC-10 and MD-11 airplanes,
overhauled, high-strength alloy steel Airline personnel who participate in use CMM 20-11-02; for DC-9,
components demonstrate high levels of component rework, maintenance, and MD-80, MD-90, and 717 airplanes,
service reliability. overhaul tasks should be familiar with use CMMs 20-10-18 and 20-10-06.
the properties of high-strength steels
The decision to use high-strength and understand the negative effects that In addition, airline personnel need
alloy steels is based on weight and eco- can result from to understand the importance of main-
nomic factors. Airframe space for gear taining component finishes while
components may be reduced because ■ Sensitivity to stress concentrations in service (in situ, or on the airplane).
of smaller diameter shock strut com- (notch sensitivity). This includes repairing damaged
ponents, smaller pins (reduced space finishes to prevent corrosion and en-
for joints), smaller diameter trucks and ■ Microstructural damage from suring that solvents and materials that
axles, and, in some instances, smaller embrittlement or overheating. come in contact with the finishes do
drag brace, side brace, and attach fit- not result in premature degradation
tings. By reducing the space required ■ Detrimental surface conditions. and unscheduled component removal.
for these components, the wheel well
size can be minimized and aerodynamic ■ Corrosion. Boeing documentation describes
surfaces can be optimized, which allow the methods for detecting base metal
an increase in fuel tank size (optimal ■ Improper processing. damage while in service and during
wing spar location) or additional space overhaul. Common techniques include
for other uses. Improper rework practices can result detailed visual inspections and other
in unscheduled maintenance or surface nondestructive inspection methods,
The use of high-strength alloy steel damage that causes crack initiation. such as magnetic particle inspection
parts is economical because it reduces Maintenance efforts focus on corrosion (MPI) and fluorescent penetrant inspec-
weight, thereby allowing for more prevention and removal in addition to tion (FPI). (See SOPMs 20-20-01 and
efficient aerodynamic surfaces and normal checks for wear and free play. 20-20-02.) Ultrasonic or eddy current
providing the potential for increased inspections also may be useful for
payload and fuel. High-strength alloy steels can ex- in situ inspections.
perience rapid crack propagation from
For example, the trailing edge of stress corrosion under certain loading Boeing also is developing sup-
the wing is relatively shallow. Using conditions. Therefore, surface damage plemental, specialized techniques,
high-strength alloy steel flap tracks, detection is important during overhaul such as the Barkhausen inspection, to
flap carriages, and flap actuating com- and on components in service. detect base metal heat damage under
ponents reduces the profile and decreases Removing visible surface corrosion chrome plating or other protective
spatial envelope requirements while before pitting begins (such as during finishes. This technique can be used
meeting or improving aerodynamic a C-check) helps prevent conditions successfully to screen components with
requirements. This also optimizes wing that can lead to crack initiation. The suspect damage. For example, if an axle
shape and reduces the potential need for best safeguard against corrosion is fractures as a result of chrome-grinding
bulging aerodynamic surfaces, which to ensure that finishes conform to the heat damage during manufacture or
in turn reduces drag and increases design and that design improvements overhaul, the Barkhausen inspection
airplane performance. are incorporated as minor changes allows other suspect components to be
whenever possible. screened without first performing a
chrome strip and temper etch (e.g., nital
Components manufactured from etch) inspection on all suspect axles.
steel alloys heat-treated above 180 ksi
(180,000 psi) should be reworked in
accordance with guidelines in
Component Maintenance Manuals
(CMM) 32-00-05, 32-00-06, and

Second-Quarter 2003 — April AERO 25

3 GUIDELINES FOR REWORKING to surfaces to prevent wear or corrosion, sition to the adjacent surface and
HIGH-STRENGTH ALLOY STEEL the coating must exhibit proper runouts usually is kept at the minimum depth
COMPONENTS that terminate before the tangent of fil- necessary to clean up the damaged
This section provides guidelines for let radii, edges, or other shape changes. surface. Spot face depressions typi-
reworking high-strength alloy steel Boeing SOPM guidelines should be fol- cally are not filled with plating to
components and describes some of lowed for the rework of any component restore the dimension but instead are
the implications of improper rework and for all types of plating or coating. finished in the same manner as the
procedures. Rework or overhaul of components original design. Spot face transition
should not introduce stress concentra- radii need to be such that they can
■ Stress concentrations. tions, or otherwise increase stresses, be shot-peened to the requirements
which can reduce the service life of a of the adjacent surfaces.
■ Overheating components. component below that of the original
design configuration. ■ When the entire face of a lug must
■ Hydrogen embrittlement. be machined to remove damage, the
Stress concentrations can lead to new lug transition radii should be
■ Cadmium embrittlement. initiation of cracking by fatigue, stress shaped and positioned in accordance
corrosion, or hydrogen-assisted with CMM requirements. Surface
■ Improper finishing. stress corrosion. These cracks may transitions into the lug hole and at
result in a fracture or scrap of a com- the lug edges must have design tran-
STRESS CONCENTRATIONS ponent when found while in service sitions that will allow restoration of
or during overhaul. The following are shot-peening on all reworked areas
During component design, eliminating examples of stress concentrations that and permit complete seating of bush-
or minimizing areas of stress concentra- can lead to cracking. ings without contacting hole edges.
tions is a key objective. Special atten-
tion is given to protective finish runouts Transitions or radii that are sharper Abrupt changes in sections, holes,
adjacent to stress concentration details. than original design. When removing and sharp-cornered keyways should
In addition, all stress concentration damaged material from part surfaces be avoided. Proper design will reflect
details are subject to extensive testing during rework, the new transitions or generous fillets, gradual changes of
and/or analysis to ensure that no radii should not cause an unacceptable shape, and the use of relief grooves in
detrimental effects are introduced into increase in stress concentration at the areas of high stress. Finer surface fin-
a part. Any rework or repair must not location or degrade the original design ishes also may be needed to eliminate
increase stress concentrations that features. When locally machining out unnecessary stress concentrations,
degrade component durability. corrosion or damage during overhaul, especially in areas of machined radii
a gradual transition into the reworked or undercuts. Overhaul should reflect
High-strength alloy steel components depression is necessary. the same careful, detailed review that
(along with those made from other occurred during the original design.
materials) are shot-peened to create a The intent is to remove the least
shallow layer of compressive residual amount of material possible while Plating conditions and runout con-
stress at the surface. This layer helps to ensuring that all discrepant material is trols that are not in accordance with
removed and the original design design standards. During overhaul,
■ Minimize the effects of stress strength and durability are maintained. many landing gear components are
concentrations in transition areas. There are few options to restore these completely stripped to replace nickel
machined depressions to meet interface and chrome plating. In most instances,
■ Impede crack initiation and initial requirements. One type of rework or these repairs involve rework of the
crack growth caused by fatigue or overhaul, sulfamate-nickel plating, is base metal. The new plating deposits
stress corrosion. common on shock strut cylinder diame- frequently are thicker than the original
ters and is used to repair lug faces to design configuration.
■ Create a surface that will have design dimensions as follows:
minimal adverse effects from the In all cases, it is important to adhere
residual stresses of plating. ■ Local blends on inner cylinder outer to the SOPM recommendations. This
diameter surfaces and outer cylinder will ensure that the restored plating
When a surface is machined or inner diameter surfaces often are is of high quality and that it does not
ground to remove damage, the filled with sulfamate-nickel plating terminate with an abrupt edge. Through-
reworked area should be shot-peened to restore them to dimensions that thickness cracks in chrome plate (gen-
with proper overlap onto the existing are suitable for subsequent chrome erally present where there is evidence
shot-peened surface. During overhaul, plate application. of chicken-wire cracking) can lead
personnel must observe the plating to corrosion at the base metal interface
runouts specified in the CMM sections ■ Spot facing on lugs is controlled to
and SOPMs 20-10-01 and 20-42-03. have a generous radius at the tran-

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

26 AERO Second-Quarter 2003 — April

and deterioration of the plating ad- at the plating 2 FINISH DAMAGE ALONG LOWER ID SURFACE AND
hesion. Through-thickness cracking runout led to AREAS OF CORROSION PITTING AT FRACTURE ORIGIN
also can lead to fatigue or stress lug fracture.
corrosion cracking of the base metal FIGURE
beneath the plating.
Corrosion for overhaul, all evidence of corrosion
Visual evidence of chicken-wire and pitting. must be removed and finishes restored
cracking after chrome grinding indi- Corrosion pits to design requirements or better.
cates poor chrome quality and also may are stress con- The sequence of rework operations is
indicate the possibility of base metal heat centrations. As provided in CMMs 32-00-05, 32-00-06,
damage. Chicken-wire cracking noted the pit forms, and 32-00-07.
in SOPM 20-10-04 indicates that the it damages the
chrome should be stripped and replated. shot-peened Landing gear truck fractures have
layer locally occurred in service because of corro-
If the plating runouts are blended or at the surface. sion on the inner diameter of the main
machined to remove the abrupt plating The pit then gear truck beam (figs. 2 and 3). These
edge, the techniques must be well con- grows through fractures may be caused by a combina-
trolled to avoid damaging the adjacent the compressive tion of degraded protective finishes on
base metal. Improper blending can layer, and the the truck inner diameter, poor drainage,
remove the required shot-peened layer change in resid- and contact with the corrosive chemi-
or create undercuts or grooves at the ual stress state cals in washing solutions or deicing
edge of the plating that can cause and the pit compounds. Truck fractures most often
cracking in service. geometry initiate occur at maximum ground loads such
stress corrosion as after fueling or during preflight taxi.
Several in-service fractures have cracking. This type of cracking most
been attributed to improper plating often occurs on surfaces that are both Figures 4 and 5 show a drag brace
technique, poor-quality plating, prone to corrosion and exposed to sus- from which corrosion was not removed
improper runout conditions, and base tained tensile stresses while in service, completely during overhaul. The part
metal damage caused by poor blending such as the lower surface of landing gear was subsequently shot-peened, and
or machining control. trucks, axles, and the surfaces of forward new protective finishes were applied
and aft trunnions. over the residual active corrosion.
Proper use of special plating tech- This resulted in crack initiation and
niques, such as conforming anodes and Corrosion pitting also can lead to propagation while in service and the
robbers, can control plating thicknesses fatigue crack initiation depending on eventual fracture of the component.
and runouts. This can reduce the possi- the component, the location of pitting,
bility of chrome chicken-wire cracking and cyclic loading conditions. In these Mechanical damage. Stress
and poor runout details. cases, the cracks can propagate to the concentrations can be created by
critical length and result in ductile mechanical damage that compromises
Plating into a transition (radius fracture of the component. The degree the protective finishes and alters the
transition or undercut) will create a of cracking tolerated before fracture compressive shot-peen layer. This
stress concentration that can cause varies by component, crack location,
crack initiation. For example, figure 1 and component loading conditions.
shows an outer cylinder clevis plated
into the lug transition. In service, To prevent excessive corrosion, thor-
fatigue cracking initiated ough visual inspections should be per-
formed on a regular basis to evaluate
1 CRACK INITIATION AT CHROME the condition of the protective finishes.
PLATE RUNOUT INTO RADIUS Damage should be repaired soon after
it is found. Touching up damage to ac-
FIGURE cessible enamel and primer in a timely
manner can prevent the formation of
corrosion pits and reduce the need for
excessive rework during overhaul.
Rework that requires low-hydrogen-
embrittlement (LHE) cadmium stylus
plating should be performed when the
component is not loaded.

When the component is removed

Second-Quarter 2003 — April AERO 27

3 CLOSE-UP VIEW OF CORROSION PITTING AT FRACTURE ORIGIN damage often is caused by improper
maintenance practices such as jacking
FIGURE adjacent to a jack pad or an inadvertent
impact with tools or ground-support
4 LOWER PORTION OF FRACTURED DRAG BRACE equipment (e.g., tow vehicles).
Although high-strength alloy steels are
FIGURE hard and resist dents, scratches, and
nicks, stress concentrations caused by
mechanical damage can dramatically
reduce the service life of a component.

High-strength alloy steel components
also can be damaged by mishandling
during shop rework (e.g., dropping,
impact), and in some circumstances, by
foreign object debris. Possible mechani-
cal damage to a high-strength alloy
steel component should be evaluated by
the operator and repaired as needed.

If the damage is local and widespread
deformations are not evident, repair may
be similar to that required for corrosion
and pitting. All deformed material must
be removed before refinishing; deformed
high-strength steel alloy components
must not be straightened. Contact
Boeing for assistance, if needed.

5 CROSS-SECTION VIEW OF RESIDUAL
CORROSION UNDER CADMIUM PLATING

FIGURE

28 AERO Second-Quarter 2003 — April

OVERHEATING COMPONENTS 6 CRACKED MAIN GEAR damage generally is shallow and can be
OUTER CYLINDER WALL removed by machining. After overhaul
Overheating of components can change operations are completed, the compo-
the original steel temper and mechani- FIGURE
cal properties of the affected area.
Overheating damage can be caused by nent is returned to service in accor-

■ Frictional heating while in service. dance with CMM requirements.

When grinding chrome to finish

dimensions, overheating the base metal

■ Abusive machining and grinding can create UTM and OTM formations
operations during manufacture
or overhaul. under the chrome. Figures 8 and 9

show a severe grinding burn on a main

landing gear axle that resulted in a

■ Exposure to high 7 NITAL ETCH INDICATIONS OF HEAT DAMAGE fracture. Similar grinding burns
temperatures during over- ON ID OF MAIN GEAR OUTER CYLINDER also have led to the fracture of
haul bake cycles. FIGURE flap carriage spindle journals

■ Unusual conditions such (figs. 10 and 11).
as refused takeoffs and
local fires. Any visible evidence of

chrome plate distress can

The degree to which the indicate the likelihood of base
mechanical properties are
changed depends on the tempera- metal heat damage. Figures 12
ture and duration of exposure.
and 13 show a grinding burn
Overheating can result in
overtempered martensite (OTM) that led to the fracture of a
or untempered martensite (UTM)
formations in the base metal. pivot pin. SOPM 20-10-04 and
Both conditions can be detected
by a temper etch (i.e., nital etch) CMMs 32-00-05, 32-00-06,
inspection of the base metal.
UTM indications show white and 32-00-07 provide guide-
during temper etch inspections
and often are found within lines that indicate when
patches of OTM, which show
dark gray to black during temper chrome must be removed
etch inspection. SOPM 20-10-02
provides details about the during overhaul.
inspection process and inter-
pretation of the results. Some heat damage is so

Heat damage generally is severe that the heat-treat con-
removed by carefully machin-
ing the base metal. Afterward, dition of material is altered in
another temper etch inspection
is done to ensure that the adjacent areas. This widespread
machining did not create more
heat damage. reduction in metal hardness

UTM formations may be (Rockwell-C hardness readings)
accompanied by heat-induced crack-
ing within these overheated areas that, may indicate that the compo-
if left in place, can propagate while
in service. Figures 6 and 7 show nent cannot be salvaged. Axle
service-induced heat damage on the
inside diameter of a main gear outer heat damage caused by a wheel
cylinder. This component developed
extensive frictional heat damage bearing fracture may lead to
in the upper bearing contact area
such a condition.

Shop procedures such as

magnetic particle inspection

and LHE cadmium stylus

plating can cause arc burns if

appropriate precautions are not

maintained during processing.

Figures 14 and 15 show a

as a result of improper clamp-up. fracture resulting from an arc burn

The heat damage led to cracking that developed during LHE stylus cad-

through the cylinder wall. Salvage was mium plating. (Note: In this article,

not possible. cadmium plating means cadmium-

Less severe friction-induced heat titanium or LHE cadmium plating.)

damage can be found on inner cylinders Overheating will not alter the heat-

during component overhaul. This dam- treat conditions of the base metal if

age, which occurs on a more frequent the temperatures are below the original

basis, is caused by vertical motion tempering temperature. However, the

against the lower bearing surfaces. This component still may require special

Second-Quarter 2003 — April AERO 29

8 FRACTURED MAIN LANDING GEAR AXLE 9 NITAL ETCH INDICATIONS OF HEAT DAMAGE
ON OD OF AXLE AFTER CHROME REMOVED
FIGURE
FIGURE

consideration (or rework) because of the enamel, primer, or baking, which is performed directly
chrome or evidence of cad- after stripping or plating operations dur-
■ Shot-peening may be compromised mium damage on the inner ing overhaul, effectively remove hydro-
(heated above 400°F). diameter of the axle may gen generated during these operations.
require the heat-damaged Processes that must be followed with
■ Cadmium embrittlement may occur component be removed relief baking include chrome, sulfamate-
(heated above 450°F with cadmium from service. nickel, and LHE cadmium plating; strip-
plating present). ping operations; and many nital etch
Overheating affects inspections. After hydrogen-generating
■ Chromate conversion coating may components to various operations, relief bake delay time limits
be degraded (heated above 400°F). degrees; in some instances, must be observed to ensure complete
only finish durability is hydrogen removal. In general, the best
■ Organic coatings or sealants may degraded. This may result practice is to initiate baking as soon as
crack or become brittle or discolored in a shorter than planned possible following a plating operation.
(wide range of temperatures). time between component
overhauls. Contact Boeing The delay time between plating
These situations often occur when for assistance with ques- completion and baking start typically is
components are tions about repairing or salvaging observed. However, when thick plating
high-strength alloy steel components deposits or multiple plating operations
■ Inadvertently overheated in an oven. that appear to have been damaged by are performed on a single component,
overheating. the total time between initial plating
■ Exposed to elevated temperatures start and baking start is a key factor
with some finishes intact or HYDROGEN EMBRITTLEMENT when determining the maximum delay
bushings installed. time allowed. For example, embrittle-
Hydrogen embrittlement occurs ment relief baking must begin 10 hr
■ Exposed to fire. when a high-strength alloy steel after sulfamate-nickel plating is com-
component absorbs hydrogen, which pleted or within 24 hr after plating
Residual cadmium often is left on is not removed in a timely manner
a part during overhaul processing to in accordance with the SOPM Second-Quarter 2003 — April
protect it from corrosion. The part is (e.g., embrittlement relief baking).
then stripped of all cadmium and re-
plated near the end of overhaul. Parts When hydrogen remains in a
with residual cadmium should not be component for an extended time, the
heated over 400°F during overhaul. microstructural damage that develops
significantly degrades the mechanical
Bushings should not remain installed properties of the steel. The infused
during overhaul unless retained by hydrogen migrates to areas of high
specific CMM requirements. Bushings stress (e.g., material internal stresses)
must be removed to permit a thorough and creates local microstructural dam-
inspection of the base metal and to age. When the component is installed
avoid bushing-to-bore interface degra- on an airplane, this internal damage
dation during bake cycles. Design fin- can lead to crack initiation and propa-
ishes are restored and new bushings gation, resulting in component fracture.
with design interferences and dimen-
sions are installed because bushing wear The elevated temperatures reached
limits do not apply during overhaul. during hydrogen embrittlement relief

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

30 AERO

10 FRACTURE THROUGH 11 GRINDING DAMAGE VISIBLE ON CHROME PLATE AT FRACTURE ORIGIN
FLAP CARRIAGE SPINDLE
FIGURE
FIGURE

begins, whichever results in the shortest 13 CLOSE-UP VIEW OF PIVOT
overall bake delay. PIN OD AT FRACTURE ORIGIN

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

12 FRACTURED PIVOT PIN

FIGURE

microscope view of a typical hydrogen embrittlement by cadmium can occur short and discoloration of the enamel
embrittlement crack where separation at temperatures below the cadmium or primer was minimal, the component
occurs along grain boundaries. melting point. These effects on the may be a candidate for salvage. Slight
Typically, hydrogen embrittlement base metal can begin to occur at 450°F, or no discoloration of the enamel or
cracks propagate rapidly once loads whereas the cadmium melting point primer may indicate the cadmium
are applied to the part. In some cases, is generally 610°F. The microstructural plating was not heated to the extent
internal residual stresses are sufficient- anomalies resulting from cadmium that cadmium embrittlement would
ly high to cause cracking even embrittlement can lead to component be suspected. Boeing can assist in
before the part is installed. fractures in service. this determination.

CADMIUM EMBRITTLEMENT Determining whether cadmium has AERO 31
migrated into the grain boundaries of
Overheating LHE cadmium or cadmium-plated, high-strength alloy
cadmium-titanium plated components steel components requires destructive
causes embrittlement of high-strength testing of the components. If these
alloy steel by cadmium, resulting components have been overheated,
in cadmium diffusion into the steel salvage may not be possible. However,
grain boundaries. Solid-metal if high-temperature exposure was

Second-Quarter 2003 — April

15IMPROPER FINISHING NITAL ETCH INDICATIONS OF ARC BURN
HEAT DAMAGE THAT LEAD TO FRACTURE

Improper application FIGURE

of protective finishes

during manufacture

or overhaul can lead

to finish degradation,

corrosion, and cor-

rosion pitting, which

can result in com-

ponent fracture while

in service (figs. 2

and 3, pp. 27–28).

Some cleaners and

chemicals may

accelerate finish

degradation and lead to corrosion. during overhaul (including removal of As a rule, if material removal exceeds

Operators should ensure that bushings and bearings in all structural 0.0015 in (or 10 percent of the Almen

cleaners and chemicals are tested components). This allows a thorough strip intensity), the surface should then

before use in accordance with Boeing inspection of the base metal (a primary be shot-peened to CMM requirements.

document D6-17487, Evaluation component overhaul requirement) and Exceeding shot-peen requirements is

of Airplane Maintenance Materials. ensures that all finishes, including the better than leaving areas without shot-

Testing to these requirements will LHE cadmium plating and conversion peening. All portions of a component

determine whether a cleaner or coating, are restored to the original that are to be shot-peened should first

chemical is detrimental to protective design requirements. This is addressed be completely stripped; no cadmium

finishes or base metal. However, in an all-model Boeing service letter residue should remain on the surface.

long-term exposure to the solution dated April 23, 2002, Overhaul

or material still may adversely of High Strength Steel
affect finishes. 16Components–Cadmium Strip VISIBLE CRACKS IN FLAP TRACK (ARROWS)

Required (e.g., FIGURE
757-SL-20-036-A,
14 FRACTURED ACTUATOR BEAM 767-SL-20-038-A,

FIGURE 747-SL-20-062-A).

Restoration of

the shot-peened

layer during over-

haul is important

to ensure that the

shot-peen com-

pressive residual

stresses are main-

tained or restored.

Removing or

Personnel must ensure that materials damaging the shot-peened layer can 17 INTERGRANULAR FRACTURE CAUSED
used for activities such as cleaning and reduce the protection that this com- BY HYDROGEN EMBRITTLEMENT
deicing conform to Boeing document pressive layer provides against fatigue
D6-17487 requirements and will and stress corrosion crack initiation. FIGURE

accomplish the intended task (verified Discontinuous shot-peening can lead

by the material provider or operator). to crack initiation at the tensile surface

Refer to the Aircraft Maintenance stresses adjacent to edges of abrupt

Manual for materials specified for air- compressive layer runouts (no fade-

craft cleaning and deicing. The CMM out). All reworked surfaces must be

specifies the materials for use in repair. shot-peened after removing material

High-strength alloy steel compo- damaged by corrosion, heat, and

nents should be stripped completely deformation.

32 AERO Second-Quarter 2003 — April

SUMMARY

High-strength alloy steels are used importance of maintaining component Operators should ensure that
widely in landing gear, flap track, flap finishes while in service, follow proper SOPM and CMM documen-
support carriage, and flap actuating proper rework practices, and use tation is used during overhaul and
components on Boeing airplanes. These Boeing-provided maintenance rework of high-strength alloy steel
high-strength materials provide signifi- procedures, planning guidelines, and components. The planning flowcharts
cant structural benefits and can result in checklists during scheduled main- in CMMs 32-00-05, 32-00-06, and
a weight savings. These parts often are tenance and overhaul processes. 32-00-07 are value-added guidelines
selected for placement in limited-space
Improper rework and overhaul for planning the rework of any
envelopes (e.g., wheel wells and wing practices may result in loss of finish, high-strength alloy steel component
trailing-edge support structures) corrosion, and damage to or alteration on a Boeing airplane.
because of their reduced profile or of the base metal, which may require
smaller diameters. unscheduled maintenance between
overhauls. The resulting damage also
With these benefits comes a need could precipitate crack initiation and
for airline personnel to exercise removal of the part from service.
precise care when reworking high- Removing corrosion and restoring worn
strength alloy steel components interfaces on a periodic basis are the
during scheduled maintenance and main emphases of high-strength alloy
overhaul. They need to understand the 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.

Editor’s note: The SOPMs and CMMs identified in this article can be ordered through the Data and Services Catalog. AERO 33

Second-Quarter 2003 — April

About the Authors

Ralph M. (Mike)
Garber has been a
structures engineer in
the aircraft industry
for 38 years, including
28 years as lead
engineer and the past
7 years as an Associate
Technical Fellow. His
primary responsibilities
have involved design and certification analysis of the wing, empennage, nacelle struts,
and landing gear with a focus on airline workshops and identifying design improvements to
increase durability and fatigue resistance. He is a licensed professional engineer and has
assisted in NTSB investigations and Boeing- and FAA-recommended structural changes.

34 AERO Craig Dickerson has been a
metallurgical engineer in the aero-
space industry for 18 years. As lead
engineer for the Boeing Materials
Technology Landing Gear Design
Center support group and former lead
engineer for the Fracture Analysis
group, he is involved with all aspects
of landing gear structure materials
and processes, including detail part manufacture, in-service
performance, component overhaul, analysis of parts returned
from service, and accident and incident investigations.

Second-Quarter 2003 — April

INCREASED USE OF TITANIUM identified for both steel and titanium avoid contamination from water
ALLOYS AND TUNGSTEN- alloy components in production and as and cleaning solutions.
CARBIDE COATINGS a substitute for chrome plating during
component overhaul. Some repair ■ If washing is done at the beginning
Boeing and the industry are working agencies and airlines are purchasing of a scheduled maintenance period,
together to develop thermal spray coat- equipment and preparing facilities for operators should not wait until the
ings to replace chrome plating. These HVOF coating application during over- end of the period to perform lubri-
coatings, which currently are used in haul as an alternative to chrome plating. cation — the elapsed time may not
some production and repair applications, be acceptable.
can be applied using the high-velocity MAINTAIN FINISH DURABILITY
oxygen fuel (HVOF) process. The two THROUGH PROPER WASHING, ■ When flushing or rinsing landing
coatings primarily in use are tungsten- CLEANING, AND FREQUENT gear assemblies, operators should
carbide-cobalt and tungsten-carbide- RELUBRICATION reduce spray pressure, ensure that
cobalt-chrome. the nozzle is at least 12 in from the
Properly restoring the finish of high- joints, and replace the corrosion
These tungsten-carbide coatings can strength alloy steel components to origi- preventive compound after washing.
be applied to steel and titanium alloys. nal design conditions during overhaul (See the multimodel maintenance
Steels can be either chrome plated or minimizes the effects of washing and tip, “Airplane High Pressure
coated with tungsten-carbide. Titanium cleaning solutions, solvents, and com- Washing,” May 18, 1999.)
alloys cannot be chrome plated but can pounds on the structure. Design-quality
be coated with tungsten-carbide through finishes are less likely to degrade in Impact of Aggressive Washing on
the HVOF process. With this coating, the service. In addition, frequent relubri- Finishes and Lubrication
lighter weight titanium can be used in cation of these components soon after
more landing gear and flight control washing protects finishes at lubricated ■ Short-term exposure to materials that
applications where high-strength alloy interfaces. Relubrication intervals are normally contact properly restored
steels would have been used in the past, specified in the Aircraft Maintenance finishes, such as solvents, should not
resulting in a weight savings. Manual (AMM) but generally are ad- cause premature degradation or loss
justed based on operator experience. of finish requiring repair or unsched-
Titanium alloys are being used more uled removal between overhaul.
often in the design of new landing gear Boeing continues to receive reports However, premature corrosion and
and flight control support structure. of premature corrosion from operators deterioration can occur in service
These materials, which are becoming that use pressure-washing techniques when water or foreign material enters
more readily available, exhibit higher on their airplanes. The following guide- joints as a result of spraying cleaning
strength-to-weight ratios than do steel lines will help operators maintain the solutions directly into joints. This
alloys. Titanium alloy components finishes of their high-strength alloy aggressive washing technique dis-
require less finishing, are more easily steel components through heightened places grease and negatively affects
maintained, are less prone to corrode in awareness and knowledge about key lubricated joints even though immedi-
service, and require less overhaul pro- aspects of airplane washing processes. ate relubrication will purge most con-
cessing than most high-strength alloy taminates from the lubricated cavities.
steel components. The durable, HVOF- Washing and Cleaning Techniques
applied tungsten-carbide coatings broad- ■ Most corrosion-related cracking and
ens possible titanium alloy applications. ■ Operators should avoid using fractures in service are aggravated by
high-pressure washing. aggressive washing techniques and
HVOF-applied tungsten-carbide coat- corrosive solutions. To help ensure
ings also provide multiple process bene- ■ When cleaning landing gear and other that finishes do not degrade pre-
fits when compared with chrome plating: mechanical, electrical, or hydraulic maturely between overhaul, operators
components, operators must follow should lubricate all greased bearings
■ Embrittlement baking is not needed the requirements in the AMM and cavities no later than 12 hr after
after application because hydrogen procedure Remove Material Around airplane washing. Relubrication and
does not infuse into the base metal. Sensitive Components (e.g., 747-400 replacement of corrosion preventative
AMM 12-25-01, p. 301, 2.A.[2]). compounds within this time period
■ When grinding the coating to the minimizes finish exposure to cor-
desired finish, overheating the base ■ After washing landing gear and con- rosive cleaning agents following
metal is less likely. trol surfaces, operators should com- airplane washing.
plete rinsing within the specified time
■ Applying coatings using the proper period. Rinsing must not be delayed. High-pressure washing
HVOF procedures and equipment can is detrimental and
save shop processing and flow time. ■ Operators should cover joints
without relubrication fittings to should not be used under
These benefits are driving more any circumstances.
tungsten-carbide applications to be

Second-Quarter 2003 — April AERO 35