Engineering Failure Analysis 160 (2024) 108233
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Engineering Failure Analysis
journal homepage: www.elsevier.com/locate/engfailanal
Failure analysis on abnormal cracking of the main landing gear
door of a civil aircraft
Tong-Hao Jiang , Zhen-Guo Yang *
Department of Materials Science, Fudan University, Shanghai 200433, PR China
A R T I C L E I N F O
A B S T R A C T
Keywords:
Landing gear door
Cracking
Failure analysis
Sandwich structure
Delamination
Among the myriad structural and moving components integral to civil aircraft, only a select few
confront the dual challenges of meeting aerodynamic and mechanical performance criteria, a
circumstance exemplified by landing gear doors. The failure of landing gear poses a substantial
threat to both the structural integrity and the safety of an aircraft during flight. Surprisingly, the
existing literature lacks an in-depth investigation into the failure of landing gear doors. This paper
addresses this gap by examining the abnormal cracking failure observed in the main landing gear
door of a civil aircraft. The purpose of this paper is to find out the root cause of the failure, so a
comprehensive investigation of the landing gear outer door is carried out. This investigation
employed various techniques, including computed tomography, three-dimensional scanning
analysis, macroscopic and microscopic morphology analysis using optical microscopy, threedimensional stereomicroscopy, scanning electron microscopy, and micro-area composition
analysis using energy dispersive spectrometry. Through repeated sorting and comprehensive
analysis of various failure forms and defects, the root causes of the main landing gear door failure
were conclusively identified. Evidence revealed that the crack originated at the upper edge corner
of the main landing gear door, precisely where abnormal contact and wear with the hinge joint
occurred. Targeted countermeasures were proposed from aspects such as design, manufacturing,
and installation, and the suggested improvements were adopted.
1. Introduction
The landing gear is an essential component for the takeoff and landing functions of civil aircrafts. During the flight, the aircraft
needs to maintain a specific aerodynamic shape, so the landing gear needs to be retracted inside the fuselage. At this time, landing gear
doors are necessary to protect the landing gear from exposure and maintain the aircraft’s aerodynamic shape. Therefore, landing gear
doors need to be able to open and close smoothly to meet the requirements of proper extension and retraction of the landing gear, while
also withstanding flight loads and maintaining the structural integrity of the aircraft [1,2].
The landing gear door mainly consists of panels and joints, with the door panel serving as the main load-bearing structural
component, typically made of a sandwich structure composite material. The sandwich structure generally comprises five elements: the
upper panel, the adhesive layer bonding the upper panel to the core material, the core material, the adhesive layer bonding the core
material to the lower panel, and the lower panel. Usually, carbon fiber or glass fiber reinforced composite materials with thin thickness,
high strength and stiffness are used as panels, while honeycomb or foam materials with small density and large thickness are used as
* Corresponding author.
E-mail address: zgyang@fudan.edu.cn (Z.-G. Yang).
https://doi.org/10.1016/j.engfailanal.2024.108233
Received 16 January 2024; Received in revised form 28 February 2024; Accepted 14 March 2024
Available online 15 March 2024
1350-6307/© 2024 Elsevier Ltd. All rights reserved.
Engineering Failure Analysis 160 (2024) 108233
T.-H. Jiang and Z.-G. Yang
core materials.
In previous studies, the failure of many components on the landing gear has been investigated such as nose wheel fork [3,4], cross
tube [5,6], brake reducer [7], cylinder support [8], support strut [9,10], hinge [11], axle [12], trunnion [13] and swinging lever [14].
Previous studies have shown that fatigue fracture is one of the main factors of landing gear failure. Infante et al. [3] presented the
fatigue fracture failure of a nose landing gear fork. The fracture was initiated by cracking in holes of the fork structure which are stress
concentration zones, and propagated from the upper surface of the fork under the cyclic tensile loads. Raković et al. [5] presented the
fatigue failure analysis of the aft cross tube made of aluminum alloy on the helicopter skid landing gear. The extrusion manufacturing
resulted in numerous surface defects at inner surface of aft cross tube, which were stress concentration locations and potential crack
initiation locations. Additionally, improper following of the prescribed procedures of parking and mooring by user manual led to
higher loads than the expected nominal static loads in the parking position. Diltemiz [8] reported fatigue damage of the main landing
gear cylinder support and the main cause of failure was the combination of a heavily loaded aircraft and hard landings. The fatigue
crack initiates in a highly stressed sharp corner and propagates under cyclic loads. The rough tool marks and shallow pits on the crack
initiation area are also contributing factors. Zucca et al. [11] investigated the fracture of main landing gear door hinge. The root cause
of failure was a local loss of coating, thus leading to the pitting corrosion of the back hinge and further fatigue fracture under normal
operative cyclic loads. Infante et al. [13] analyzed failure of trunnions of landing gears of military aircrafts. The failure occurred by a
fatigue process but in the presence of high stress levels at the crack initiation zone. Bagnoli et al. [14] investigated a fatigue fracture of
the swinging lever made of 7010 aluminum alloy. The fatigue originated from material defect on the external surface in correspon­
dence of the most stressed area.
In addition to fatigue breakage, overloading and corrosion are also possible causes of landing gear component failure. Jalaja et al.
[7] investigated the failure of the main landing gear brake reducer and the root cause of failure was the ductile load in the short
transverse direction. The failed reducer had grains oriented perpendicular to the loading, which imposed stresses in the weakest di­
rection. Hameed et al. [9] presented the failure analysis of a broken support strut of an aircraft landing gear. The root cause of this
fracture was the improper installation of the support strut, resulting in the dislodging of the left retaining pin. This localized con­
centration of stress, combined with the inability to effectively disperse or transfer the impact forces throughout the structure, caused an
overload situation. Freitas et al. [12] presented the failure of an axle from the nose landing gear during landing. The root cause of
failure was overloading due to both shear and bending stresses because the aircraft landed with the nose landing gear instead of the
main landing gear. Krstic et al. [6] reported the failure analysis of the fractured cross tube on the skid landing gear. The failure was the
result of stress corrosion cracking with a combination of AISI 4135 alloy steel, bending load and retained moisture in the flange area.
Turan et al. [10] investigated the cracked main landing gear strut bearing and the cracks were initiated by corrosion and assisted by
fatigue and the crack propagation was accelerated by corrosion.
However, the published literature on the failure of the landing gear and its components is metal material failure, not composite
material failure. With the increasing proportion of composite materials used in aircrafts and the trend of lightweight, the existing
Fig. 1. Location of the main landing gear left outer door.
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failure cases of metal components of landing gears can no longer be satisfied. Carbon fiber reinforced polymer (CFRP) sandwich
composites are increasingly utilized in collision-resistant parts due to their excellent mechanical characteristics [15–20]. In addition,
there is no detailed research on landing gear door failure in the existing literature. The existing literature on landing gear doors mainly
studied the aeroelastic stability [21], vibration characteristics [22], structural design and optimization [23–26]. There is a surprising
paucity of studies investigating the failure of composite components of landing gears, especially landing gear doors.
In this paper, the failure of the main landing gear left door of a civil aircraft was investigated. An abnormal cracking occurred
during a full side-slip test flight of a certain type of civil aviation aircraft at the maximum landing gear extension speed. The aircraft
first slid to the left for 2 min, then to the right for 2 min, reaching a speed of 280 knots (approximately 519 km/h), exceeding the
required 235 knots specified in the design specifications. This was a deliberate test flight conducted by the crew. The aircraft has a
tricycle landing gear arrangement and there are main landing gear doors on both sides of the fuselage. A single-side main landing gear
door consists of three panels, namely the inner door, middle door, and outer door. Post-flight inspection identified damage to the main
landing gear outer door, with the left side outer door experiencing abnormal cracking which was irreparable, as shown in Fig. 1. The
right side outer door, wing, and landing gear also sustained varying degrees of slight damage.
In this study, the failure analysis of the abnormal cracking on the left main landing gear outer door during a full sideslip test flight
was carried out. Our primary objectives were to ascertain the root causes of failure and propose targeted countermeasures. Specifically,
we aimed to determine whether the cracking resulted from factors such as excessive speed leading to a load exceeding design spec­
ifications, potential quality issues with the door itself, or undetected defects prior to the test flight. Through on-site investigation,
consultations with technical experts, information gathering, visual observations, as well as a series of tests on physical samples fol­
lowed by comprehensive analysis, the root causes of the abnormal cracking were ultimately identified. It is noteworthy that the
existing literature lacks in-depth investigations into the failure of landing gear doors, particularly those made of composite materials.
Given the critical role of landing gear doors in ensuring aircraft safety, the lack of comprehensive research on their failure forms and
causes is surprising and underscores the importance of our investigation. By bridging this gap in knowledge, our study contributes to
advancing the understanding of composite material failure mechanisms in landing gear components, thereby facilitating the devel­
opment of more robust and reliable aircraft systems [27–29].
2. Experimental methods
To identify the root causes of the failure, visual and optical examination was conducted carefully on the samples of the landing gear
outer door. The macroscopic appearance and microscopic morphology were observed using an optical microscope (OM, Zeiss Axio
Imager A1m), three-dimensional stereomicroscope (3D-SM, HIROX KH-7700) and scanning electron microscopy (SEM, Phenom XL
G2). Micro-area composition was measured by energy dispersive spectrometry (EDS, Thermo Fisher Scientific Super-X). Computed
tomography (CT, YXLON FF35) was used to determine the location and extent of damage inside the sandwich structure. Threedimensional scanning analysis was conducted by laser scanner (SHINING 3D FreeScan UE11).
3. Experimental results
3.1. Visual and optical examination
3.1.1. Landing gear door
First, a preliminary visual inspection of the failed main landing gear outer door was conducted. The appearance of the inner surface
Fig. 2. Inner surface of the main landing gear outer door.
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(caption on next page)
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Fig. 3. Damage to the inner surface of the landing gear door. (a) Delamination crack at the upper edge. (b) Crack propagation path on the inner
surface. (c) Crack on the rear edge extending from the inner surface to the outer surface. (d) Crack at the edge of the upper fixed sealing strip.
of the left main landing gear outer door was shown in Fig. 2. The main landing gear outer door mainly consisted of the door body, front
and rear hinge joints, tie rod joints, and sealing strips [30,31]. The sealing strips bolted to the upper edge were removed. There were
two hinge joints, one small in the front and one large in the rear, connected to the door body and one tie rod joint in the center. The
door body was a CFRP-foam sandwich structure. The panels or face sheets on both sides were carbon fiber reinforced plastic with
symmetric laminate, and the core material in the middle was closed cell foam. The panels and core material were bonded by adhesive
[32,33]. The hinge joint and the door were tightly connected with high lock bolts, and the bolt connection area was filled with potting
resin. The thickness of the foam at different locations within the door. The thickness of the foam core in the center area was about 20
mm, while in the hinge joint area the foam core is thinner, about 9 mm in the front side and 7 mm in the rear side, and there is no foam
sandwich in the edge area around the door. The transition area where the thickness of the foam sandwich varies has an inclination
angle of approximately 20◦ .
There were two damages on the inner surface of the outer door, namely crack A in the rear hinge joint area and crack B at the fixed
edge of the upper sealing strip. After removing the rear hinge, cracking and delamination were observed in the rear side hinge joint
area that the upper edge of the door panel, intersecting with the inner surface crack at the shoulder corner at the upper edge of the door
panel, as shown in Fig. 3(a).
The path of crack A started from the corner at the upper edge of the door, passed through the edge of the hinge joint area where the
foam core became thicker, followed the edge of the hinge joint downward, turned around the 7th bolt hole, and extended along the
lower edge of the hinge joint area to the rear edge. The crack extended continuously from inner surface to outer surface, as shown in
Fig. 3(b)-(c). The path of crack B aligned with the boundary of the internal structure with or without foam core. It penetrated through
the thickness of the door panel and represented the starting boundary of the area where the foam core becomes thicker, as shown in
Fig. 3(d). On the inner surface, crack A and crack B were not connected to each other.
The appearance of the outer surface of the left main landing gear outer door was shown in Fig. 4. The outer surface crack started
from the rear edge, passed through the 7th bolt hole, and extended forward to the upper edge, but it did not reach the upper edge, as
shown in Fig. 5(a). The crack A was not connected with crack B on the surface, as shown in Fig. 5(b). The crack A on the inner surface
and outer surface were mainly in the hinge joint area with thinner foam core or without foam core, and did not extend further into the
central region with thicker foam core. Comparing the paths and directions of crack A on the inner and outer surface, they were
generally similar but not entirely identical. This indicated that the crack A did not propagate by directly passing through and pene­
trating the foam core from one side to the other. In addition, the crack A on the surface remained continuous without interruption.
Due to the relatively large sample size, and in order to further investigate the internal damage near the crack A while ensuring the
integrity of the crack as much as possible, preliminary cutting was performed on the sample as the red rectangle shown in Fig. 4. The
preliminary cutting revealed that there were long continuous cracks on the cut surface of the foam core, indicating that the actual
internal damage had exceeded the rear hinge joint area. No cracks were observed on the cut surface when cutting away from the hinge
joint area, suggesting that the cracking was not caused by the cutting process. The cracks on the foam core were either located very
close to the interface between the foam and the adhesive layer or penetrated through the foam matrix, as shown in Fig. 6(a)-(c).
Metallographic specimens were taken from the uncracked foam core transition area and foam sandwich area away from the hinge
joints, as shown in Fig. 6(d) and (g) respectively. The foamless region shown in Fig. 6(e) has fewer interlayer pore defects, while more
interlayer pore defects could be seen in Fig. 6(f) in the bevel region. The staggered layup of the carbon fiber woven fabric was clearly
visible and these pores were mostly irregularly elliptical along the lay-up direction. Further magnified interlayer pores shown in Fig. 6
Fig. 4. Outer surface of the main landing gear outer door.
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Fig. 5. Damage to the outer surface of the landing gear door. (a) Crack propagation path on the outer surface. (b) Crack at the edge of the upper
fixed seal strip.
(h)-(i) were so concentrated and continuous in distribution that they were connected as interlayer cracks. The interlayer pores and
cracks between adjacent carbon fiber layups were very easy to become the weak link in delamination cracking and to accelerate
interlayer delamination crack propagation. As shown in Fig. 6(h)-(i),
3.1.2. Hinge joints
The rear hinge joint had already been removed when the failed landing gear door was received, while the front hinge joint was still
secured to the door using high-lock bolts. According to the design requirements, there should be a certain gap between the front or rear
hinge joints and the door panel. However, it was noticed that there was abnormal contact between the inner arc segment of the front
hinge joint and the upper edge corner of the landing gear door panel, as shown in Fig. 7(a). Furthermore, it was visible from a side view
that the front hinge joint near the gap had undergone significant deformation and twisting towards the outside at the front end, as
shown in Fig. 7(b). In combination with the obvious wear at the intersection of the upper edge crack and the inner surface crack
observed earlier, abnormal wear marks were also found at the corresponding position of the arc segment on the removed rear hinge
joint, with severe wear near the front edge (left side), as shown in Fig. 7(d). The wear marks on the hinge joint and the on the upper
edge corner of the door confirmed the abnormal contact and wear between them. In addition, it was visually evident that the rear hinge
joint had undergone significant deformation, as shown in Fig. 7(c) and (e), indicating that the joint had experienced abnormal loads.
Further observation of the abnormal wear on the rear hinge joint revealed two strip-shaped wear marks, with the most severe wear
at both ends, and the left side precisely corresponding to the wear on the door’s upper edge corner. Upon closer inspection, smaller
scratches were observed on the two parallel wear streaks, oriented perpendicular to the two wear streaks, i.e., in the vertical direction.
This indicated that the rear hinge joint moved up and down in the same direction as the fine scratches, as shown in Fig. 7(f)-(h).
Furthermore, the results of three-dimensional laser scanning size measurements showed that both the front and rear hinge joints had
undergone a certain degree of deformation, as shown in Fig. 7(i)-(j). The deformation at the point where the front hinge joint’s rear side
made abnormal contact with the door’s upper edge was only about 1.6 mm, and the joint was already in contact with the door on one
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Fig. 6. Crack of foam sandwich core on cutting surface. (a) Longitudinal cutting surface. (b) Transverse cutting surface. (c) Enlarged transverse
cutting surface. (d) Specimen at the foam core transition area. (e) Foamless region. (f) Bevel region. (g) Specimen at the foam sandwich area. (h)-(i)
Interlayer pores.
side of the gap. The deformation degree of the corresponding area on the rear hinge joint exceeded that of the front hinge joint, with a
maximum deformation of approximately 25 mm, indicating more significant deformation and confirming the abnormal contact and
wear between the rear hinge joint and the door.
3.2. Computed tomography
Subsequently, a computer tomography scan analysis was conducted on the critically important rear hinge joint area (similar to
Fig. 3(b)) in order to further determine the extent and degree of damage inside the sandwich structure.
Fig. 8(a)-(c) were the three sectional views corresponding to the red cross position (the same below). Fig. 8(a) showed the rear
hinge joint area of the landing gear outer door (viewed from the inside), with gaskets for bolts 2# and 7# having detached, and the
inner surface crack passing through bolt 7#. Observation in the thickness direction in the area revealed the continuous presence of the
crack A on the internal section, and the crack propagation path is essentially consistent with adjacent sections, indicating that the main
cracks on the internal and external surfaces extend continuously within the sandwich structure. In Fig. 8(b)-(c), it was observed that
the internal fragmentation of the potting resin near bolt 7# where the crack passed through, while the other bolt holes remained intact.
Observation along the front and rear direction on the section reveals severe fragmentation in the potting area near bolt 7#, and it
completely shattered and penetrated, as shown in Fig. 8(e)-(f).
Fig. 9(a)-(c) showed three sectional views corresponding to the red cross position. CT scanning revealed a significant area of foam
loss in the bevel zone, as shown in Fig. 9(a)-(b). Additionally, it was observed that the bolt hole 2# was near the edge of the potting
resin, indicating that there was deviation in positioning during the drilling process. As shown in Fig. 9(b), the upper edge area without
foam core was open and divided from the middle, with a larger opening on one side near the edge and a smaller opening on the side
with the foam core. Additionally, the inner side CFRP panel was noticeably deformed inward. Some foam residue remained in the
sectional view, but the foam was completely missing over a larger area. Fig. 9(c) showed delamination and cracking in the foamless
area along the corresponding rear edge. Since the range of CT scanning is limited, the actual area of complete foam loss exceeded what
was observed in the actual CT scanning area, and the internal foam had cracked in multiple places.
Computer tomography scanning also revealed abnormal internal foam loss. As shown in Fig. 10(a)-(c), another area with internal
foam loss was along the propagation path of crack A. However, neither the inner nor outer surface crack was separated and opened
wide enough to form a gap for foam fragments to drop out. The foam loss in this area was isolated, with no contiguous loss area nearby.
Since there was no way for the foam to break and fall out, the lack of foam here existed before the door cracked. There was no sig­
nificant change in the total thickness, but the internal foam core cracking was obviously missing. The maximum thickness of the
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Fig. 7. (a) The abnormal contact between the front hinge joint and the landing gear door. (b) The deformed and warped front hinge joint. (c)-(e)
Left, front and right side view of the rear hinge joint. (f)-(g) Two parallel wear streaks at the arc part of the rear hinge joint. (h) Abrasion wear trace
on the two streaks. (i)-(j) Three-dimensional size measurement results of front and rear hinge joint by laser scanning.
internal foam loss was about 3.8 mm in the thickness direction, which was more than half of the thickness of the original foam core.
In addition, the foam-free zone below the upper edge had completely delaminated and cracked. As shown in Fig. 11(a)-(d), the
carbon fiber panel without foam interlayer at the upper edge delaminated and cracked from the middle at different cross-section
positions near the end of the crack at the upper edge.
3.3. Macro and micro morphology analysis
Further cutting was performed after the internal damage of the hinge region was assessed using a CT scan, as shown in Fig. 12(a).
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Fig. 8. CT scanning of crack A at hinge joint area. (a)-(c) Three-section view corresponding to the center of the red cross. (d) Three-dimensional side
view. (e)-(f) Fragmentation of the potting area.
The approximate locations of the previously identified foam loss areas were also indicated in the figure. Further observation and
analysis were conducted on the samples marked with ①-⑤ on the inner surface.
Sample 1# was the shoulder corner transition on the upper edge of the door panel, precisely at the intersection of the upper edge
delamination (interlaminar matrix crack) and the inner surface crack. After cutting, the inner surface naturally separated along the
crack, as shown in Fig. 12(b). Sample 2# was the bolt hole 7# through which the surface crack passed diagonally, with the illustrated
section crossing the foam loss area near the potting resin, as shown in Fig. 12(c). Sample 3# was the foam loss area near the potting
resin, in line with the CT results, as shown in Fig. 12(d). Sample 4#was the rear edge where the surface crack extended from the inner
surface to the outer surface, as shown in Fig. 12(e). Sample 5# was the rear rounded area of the upper edge hinge joint area, which also
naturally separated after cutting. The loss of internal foam found in the CT scan was confirmed after cutting, as shown in Fig. 12(g)-(h).
Since the wear at the corner of the upper edge and the 7# bolt hole are the most likely locations of the crack source, the samples 1# and
2# were mainly analyzed.
3.3.1. Sample 1#
Sample 1# was located at a specific position where the upper edge delamination intersected with the inner surface crack, making it
a focal point of our observation and analysis. As shown in Fig. 13(a), abnormal localized wear at the corner of the upper edge was
observed, with wear present on both the inner and outer sides, but more severe on the inner side. The inner surface crack started from
the worn area at the corner of the upper edge, while there was no corresponding crack on the outer side. On the naturally separated
inner and outer side parts, carbon fiber woven fabric layers were visible at the upper edge without a foam core in all areas except the
worn region. The vertically braided carbon fiber tapes were complete and belong to the same layer, proving to be an interlaminar
fracture, as shown in Fig. 13(b)-(d). The bolt was pulled to the interface between the layered sections, proving to be crack opening
fracture, or mode I fracture. Additionally, the foam loss in the bevel area was confirmed again on sample 1#, corresponding to Fig. 13
(a). As shown in Fig. 13(e), upon closer examination of the abnormal, there was interlayer damage and localized delamination in the
wear area. Adjacent carbon fiber fabric layers separated with traces of localized delamination cracks. These confirmed that the di­
rection of relative movement of the wear was mainly along the Z axis, that is, perpendicular to the laminate. Usually delamination
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Fig. 9. CT scanning of crack A at foamless bevel area. (a)-(c) Three-section view corresponding to the center of the red cross. (d) Three-dimensional
side view.
Fig. 10. CT scanning of crack A at foamless area near the potting resin. (a)-(c) Three-section view corresponding to the center of the red cross.
fractures are a mixture of different fracture modes rather than pure fracture modes. Mixed-mode fracture surfaces are generally a
combination of the morphologies observed under pure mode fracture surfaces [34–37]. Both typical morphologies of mode I (scarps
and textured microflow) and mode II (cusps) could be seen in Fig. 13(f)-(g), indicating that it was a mixed mode I/II fracture [38,39].
Microscopic morphology of the wear zone on the upper edge was observed in detail under a scanning electron microscope. As
shown in Fig. 14, the wear area revealed interlayer damage to carbon fiber fabric layers and localized delamination between adjacent
layers, as shown in Fig. 14(c). The length of fractured fibers in the same layer was basically consistent, and the orientation of carbon
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Fig. 11. Delamination and cracking the foam-free zone below the upper edge. (a)-(d) Different cross-sectional locations have been delaminated
and cracked.
fibers in the same tape was generally parallel. The fracture morphology of carbon fibers exhibited mixed features of shear and slight
bending. Some fracture surfaces were divided into two or more regions with different characteristics, with one side nearly perpen­
dicular to the fiber’s length direction, having a few fragment residues on the top surface, while the other side formed an angle sloping
downward or a wedge shape with a clear boundary between them. The diameter of carbon fibers was approximately 7 μm.
Adjacent to the carbon fiber fabric layers, white single-layer fibers were observed, suspected to be glass fibers, with a diameter of
about 5 μm, as shown in Fig. 15(a). Some fibers exhibited clear kinkbands, with local variations in fiber orientation at the twisting band
locations. Kinkbands are a key characteristic of compression failure, and can be interpreted by considering the micromechanics of
compression. The fracture morphology of fibers is mainly in a hybrid microbuckling and shear form, as shown in Fig. 15(b)-(c)
[34,40–42]. The composition analysis of fibers was performed using an energy-dispersive X-ray spectrometer, as shown in Fig. 15(d)
and Table 1, confirming that these fibers were indeed glass fibers.
3.3.2. Sample 2#
Sample 2# was located near the bolt hole 7#, where a surface crack passed diagonally. Fig. 16 displayed the macroscopic
morphology of inner and outer surface and inner wall of the bolt hole 7# before cutting. In Fig. 16(a) and (d), both inner and outer
surface cracks propagated along a zigzag path with multiple sharp turns. The crack caused the surface coating to crack, and the local
carbon fiber to tear and expose. Surface cracks extended to the inner wall of the bolt hole, leading to delamination of the carbon fiber
panel and continuous cracking of the inner wall of the bolt hole in the potting resin.
The sample 2# was cut along the center of bolt hole #7, and cracking was observed between the potting area and the outer carbon
fiber panel, as shown in Fig. 17(a). The longitudinal section of bolt hole 7# was further observed under the SEM. Both sides of the
carbon fiber panel cracked through the ply. It was also noted that the cracks of the potting material turned at a large angle, expanding
from the direction parallel to the ply to the direction perpendicular to the ply and connected to the cracks of the carbon fiber panel
through the ply, as shown in Fig. 17(b). It was also noticed that there were white spherical additives and their fragments in the potting
resin, and the morphology was shown in Fig. 17(c). The EDS analysis was performed, as shown in Fig. 17(d)-(e) and Table 2. The results
indicated that the white spherical additives were hollow glass beads with a diameter of about 10–50 μm. The hollow glass beads were
added to attain low density, high specific strength and stiffness, low hygroscopicity and high thermal stability.
4. Discussion
Now the main results of the above failure samples are comprehensively analyzed and discussed to determine the root cause of the
abnormal cracking of the outer door of the left main landing gear.
Firstly, there are two major damages visible on the outer door surface (Fig. 2, Fig. 4): one is the crack A in the rear hinge joint area,
and the other is the crack B at the edge of the upper fixed sealing strip. In this failure analysis, it is crucial to compare and prioritize
between these two cracks to determine the initial crack. If the determination of which one is the initial crack is incorrect, the results of
the failure analysis will undoubtedly be unreliable and meaningless. From the point of view of stress analysis, the cracking in the rear
hinge joint area is subjected to abnormal extrusion forces generated by joint contact at the shoulder corner, while the edge of the fixed
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(caption on next page)
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Fig. 12. (a) The samples 1#-5# of crack A after further cutting and the position of foam loss. (b) Sample 1#. (c) Sample 2#. (d) Sample 3#. (e)
Sample 4#. (f) Sample 5#. (g) Foam loss near the potting resin (corresponding to Fig. 10(a)). (h) Foam loss in the bevel zone (corresponding to
Fig. 9(a)).
Fig. 13. (a) The abnormal wear at upper edge. (samples 1#) (b)-(c) Inner side and outer side parts. (d) Complete carbon fiber woven fabric layer. (e)
Interlayer damage at the worn area. (f)-(g) microscopic fracture surface.
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Fig. 14. (a) The abnormal wear at upper edge (corresponding to Fig. 13(c)). (b) Magnified view of Fig. 14(a). (c) Interlayer damage to carbon fiber
fabric layers and localized delamination. (d) Views of fractured carbon fiber ends. (e)-(f) Magnified views of Fig. 14(d).
seal strip is not normally a load-bearing area. From the perspective of crack propagation and damage areas, the path of crack prop­
agation in the rear hinge joint area on the inner and outer surfaces is not entirely consistent with that of the internal foam sandwich
layer (Fig. 3(b), Fig. 5(a)). In contrast, the latter exhibits crack with almost no foam sandwich layer at the crack location, and these
cracks completely penetrate in the thickness direction, with both sides showing a fully consistent propagation path (Fig. 3(d), Fig. 5
(b)). Therefore, the former (crack A) is the initial crack, and the latter (crack B) is the secondary crack.
The secondary crack location is likely to be caused by secondary damage from the collision of the fastening bolts of the fixed sealing
strip with other components of the landing gear. The secondary collision marks are similar to the impact marks at the end of the fixed
seal strip’s bolt, but the samples of the relevant bolts were not obtained for comparison, so this is only a preliminary hypothesis for the
secondary collision.
The main failure forms and damage types observed in the hinge joint area on the rear side of the outer door can be summarized as
follows, as shown in Fig. 18 [38,43]. The corresponding figure numbers are marked in the upper left corner.
• Cracks on the inner and outer surfaces of the outer door panel (Fig. 3(b)-(c), Fig. 5(a)).
• Interlaminar cracking at the upper edge of the carbon fiber composite panel (Fig. 3(a), Fig. 13(a), (d)).
• Local wear and indentations at the corner of upper edge (Fig. 13(e)).
• Fragmentation of the encapsulation resin near the 7# high-lock bolt hole (Fig. 16(a), Fig. 17(a)-(b)).
• Cracking of the foam core within the sandwich structure (Fig. 6(a), Fig. 6(b)).
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T.-H. Jiang and Z.-G. Yang
Fig. 15. (a) The morphology of the fiber layers at the wear zone of upper edge. (b)-(c) The fracture morphology of fractured fibers. (d) Position and
spectrum of EDS analysis.
Table 1
EDS analysis results of fibers at the upper edge of sample 1# (wt. %).
Site
C
O
Na
Mg
Al
Si
Ca
A
7.09
42.86
0.50
0.30
9.59
22.38
17.28
Fig. 16. The appearance of bolt hole 7# with crack A passing through. (a) Outer surface. (b)-(c) Inner wall. (d) Inner surface.
• Interlayer crack propagation along the foam and adhesive layer interface (Fig. 6(c), Fig. 6(i)).
• Inward warpage of the upper edge of the outer door panel (Fig. 9(b)).
• Wear at the circular segment of the rear hinge joint (Fig. 7(d), (f), (h)).
• Deformation in the middle part of the rear hinge joint (Fig. 7(e), (j)).
Next, the most important thing is to sort out and determine the sequence of these different failure forms and damage types.
Noticeably, at the shoulder corner on the upper edge of the door panel, there was a crack splitting along the interlayer direction
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Engineering Failure Analysis 160 (2024) 108233
T.-H. Jiang and Z.-G. Yang
Fig. 17. (a)-(b) The appearance of longitudinal section of bolt hole 7# after cutting. (c) Positions of site B and C in EDS analysis. (d)-(e) EDS spectra
of site B and C.
Table 2
EDS analysis results of the fragments at the potting resin in sample 2# (wt. %).
Site
C
O
Na
Al
Si
Ca
B
C
13.09
25.30
48.15
45.30
4.90
3.20
0.70
2.60
25.38
18.50
7.79
5.10
(Fig. 3(a)). Under normal circumstances, this area was not a load bearing area, and the delamination of the inner and outer sides from
the middle is extremely unusual. The appearance of the upper edge indicated that the delamination cracking originates from the
shoulder corner and propagated along the surface and inside (Fig. 13(a), (d)).
The wear morphology of the inner surface of the rear hinge joint (Fig. 7(f), (h)) and the corresponding position at the corner of the
upper edge of the landing gear door (Fig. 13(e)) strongly proved that the wear marks were caused by the reciprocating relative motion
of the rear hinge joint and the upper edge of the landing gear door. The formation of such wear marks confirmed that the buffeting
amplitude of the upper edge of the hatch door was too large during the test flight. It also proved that the assembly gap between the
hinge joint and the upper edge of the landing gear door panel was too small. If there was a slight lateral deflection when installing the
hinge joint, or the gap was not completely parallel, and the wear marks would be more serious on one side. At this time, the hinge joint
and the upper edge of the door form a lateral wear, which caused a local eccentric wear defect at the corner.
Between the rear hinge joint and the upper edge of the door panel, there was not only eccentric wear but also significant extrusion
forces, as evidenced by the mid-section bending deformation of the rear hinge joint and the inward bending deformation of the inner
carbon fiber panel at the upper edge (Fig. 9(b)). The rear hinge joint and the door panel had abnormal contact at half the thickness of
the upper edge corner, generating substantial extrusion forces. The contact position is located in the middle of the corner of the upper
edge, and it is also a local eccentric wear defect with significant stress concentration. The extrusion force, on one hand, caused
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Engineering Failure Analysis 160 (2024) 108233
T.-H. Jiang and Z.-G. Yang
Fig. 18. Main failure forms and damage types of the landing gear door and rear hinge joint.
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Engineering Failure Analysis 160 (2024) 108233
T.-H. Jiang and Z.-G. Yang
interlayer delamination on both the inner and outer sides of the panel at the upper edge, with the inner side undergoing noticeable
bending deformation. On the other hand, it tore the panel at the wear-damaged corner, and the crack continued to propagate from the
inner surface to the outer surface.
After that, internal delamination extended downward from the upper edge, starting with delamination cracking of the carbon fiber
composite panel without foam core along the upper edge. When it extended to the region with the foam core, it led to separation of the
foam core from the adhesive layer and cracking of the foam matrix. When the foam matrix cracks intersected with the interlayer
delamination, local foam chunks detached from the matrix (Fig. 6(a)-(c)). In reality, the internal damage area was much larger than the
surface crack area.
In addition, a significant presence of interlaminar and intralaminar voids in the carbon fiber panels was observed (Fig. 6(h)-(i)).
Although these defects may facilitate the delamination crack propagation within the carbon fiber panels, they were not the key reason
for failure. This was because the panel was subjected to the abnormal force perpendicular to the direction of the panel when it flut­
tered. Even if there were no interlaminar or intralaminar voids, it could hardly resist or prevent delamination cracking between the
layers. However, such defects can be significantly reduced by improving vacuum hot pressing molding process.
Localized foam loss was observed (Fig. 9(a), Fig. 10(a)-(b)), with foam missing near the bolting area without a clear detachment
path, indicating previously undetected manufacturing defects. However, in the chamfered area, the missing foam might have
completely shattered and then flown out from the open upper edge (Fig. 11(g)-(h)). However, these defect locations were not in the
crack initiation area, so they only contribute to internal delamination propagation and were not the direct or root cause of failure.
The fragmentation of the resin in the bolting area through which the cracks pass (Fig. 16(a), Fig. 17(a)-(b)) was due to increased
loading on the high-lock bolt after internal cracking along the interface of the adhesive layer. It was difficult to determine definitively
whether defects existed in the resin prior to causing the cracks to pass through. However, CT scans clearly showed that the bolt holes
were not located at the center, and some bolt holes were very close to the edge of the resin, indicating that the positioning or drilling
was not accurate enough.
Since detailed material and structure information was not available, EDS analysis was used to verify the basic type of uncertain
constituent materials when observing sample morphology with SEM. Extra glass fiber layer (Fig. 15 and Table 1) and hollow glass
beads in the potting resin (Fig. 17 and Table 2) were verified using EDS analysis. However, these were not related to the failure cause of
the landing gear door, so no further discussion is made.
In summary, the failure process of the main landing gear outer door can be described as follows. During a full sideslip flight test
with excessive airspeed, the laterally tilted outer door was subjected to excessive aerodynamic load and vibration amplitude. In
addition, insufficient assembly gap between the hinge joint and the upper edge of the outer landing gear door makes them abnormally
in contact. Subsequently, continuous wear and stress concentration occurred at the upper edge corner, leading to interlayer delam­
ination and cracking at half the thickness of the upper edge, with the cracked inner surface further exhibiting tearing deformation.
Internal cracks extended along the interface between the foam layer and the adhesive layer, as well as within the foam core. Surface
Fig. 19. Schematic illustration of the landing gear door failure.
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Engineering Failure Analysis 160 (2024) 108233
T.-H. Jiang and Z.-G. Yang
cracks propagated from the inner surface through the rear edge towards the outer surface. Some cracks intersected internally, causing
localized fragmentation of the foam matrix. Due to insufficient rigidity, the outer door underwent deformation, and the hinge joint also
deformed due to its relatively small height of the I-beam. The insufficient stiffness led to further aggravation of abnormal contact wear
and extrusion force between the upper edge and the hinge joint, and the crack propagation was accelerated, thus forming a vicious
cycle. Finally, the outer door cracked and failed due to damage to structural integrity. The schematic illustration of the failure process
was shown in Fig. 19.
5. Conclusions and countermeasures
In this paper, the abnormal cracking failure of the main landing gear door of a civil aircraft during a test flight was investigated. The
failure analysis study sorted out multiple failure forms involved and revealed reasons for the cracking of the main landing gear door
near the rear hinge. The conclusions are as follows:
1. The aircraft exhibited excessive speed and a large deviation angle of the landing gear door during a full side-slip test flight. The
reciprocating buffeting amplitude and deformation of the outer landing gear door under turbulent flow were too large, exceeding the
design-specified load capacity. This is the root cause of the abnormal cracking of the outer landing gear door.
2. The assembly gap between the hinge joint and the upper edge of outer landing gear door was too small. When the buffeting
amplitude of the outer landing gear door was large, it was in direct contact with the joint, resulting in additional extrusion force. This
leads to delamination of the carbon fiber panel due to compression which is another root cause of abnormal cracking of the outer
landing gear door.
3. The hinge joint was designed with unequal stiffness at the connection with the outer landing gear door. The stiffness at the
section change is insufficient, and the bending deformation of the joint is too large after abnormal stress, which is also an important
reason for the abnormal cracking of the outer landing gear door.
4. There are some internal defects in the sandwich composite structure of the outer landing gear door, such as the lack of foam, the
pores between the carbon fiber layers, and the insufficient local bonding strength between the foam and the adhesive layer. These
defects resulted in an overall decrease in stiffness, which is another significant cause of excessive deformation of the carbon fiber panel.
Based on the findings of the failure analysis study, targeted countermeasures were proposed from aspects such as design,
manufacturing, and installation to prevent similar failures from occurring in the future. The suggested improvements were adopted.
1. Control the additional load of the outer landing gear door in order to minimize the amplitude of the airflow buffeting and ensure
that the bending deformation remains within the allowable range specified in the design.
2. Adequate gap or clearance margin should be provided between the joint and the upper edge of the door to prevent any contact
between the door and the joint even when the buffeting amplitude is significant.
3. Optimize the shape of the hinge joint and enhance the height of the I-beam section in the hinge joint utilizing simulation techniques
to enhance the door’s bending stiffness.
4. Address manufacturing defects in the foam sandwich structure to minimize the sources of stress concentration and prevent the
structure from becoming a weak link in the crack propagation process.
CRediT authorship contribution statement
Tong-Hao Jiang: Writing – original draft, Methodology, Investigation, Data curation. Zhen-Guo Yang: Writing – review & editing,
Supervision, Project administration.
Declaration of competing interest
The authors declare that they have no known competing financial interests or personal relationships that could have appeared to
influence the work reported in this paper.
Data availability
The authors do not have permission to share data.
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