Aircraft Structural Assessments in Data-limited Environments A Validated FE Method.pdf

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ENFOQUE UTE, VOL. 15, NO. 4, OCTOBER 2024, pp. 11-18, E-ISSN: 1390-6542
Aircraft Structural Assessments
in Data-limited Environments:
a Validated fe Method
Aun Haider1
Abstract — Aircraft operators often modify aircraft configura-
tions, install new equipment, and alter airframes to accommodate
this equipment, leading to operations in flight envelopes different
from original design profile. These modifications necessitate air-
frame structural assessments, which typically require compre-
hensive aircraft design data, often unavailable to operators. This
study aims to develop and validate a practical method for finite
element analysis (FEA) of aircraft structures in the absence of this
detailed design data. Focusing on a case study involving structural
analysis of an aircraft wing, this study presents assumptions and
idealizations used to develop 2.5D finite element (FE) model of the
wing. Fidelity of this model is established by comparing FE analy-
sis results with experimental data. Key validation metrics include
reaction forces, load distribution at wing-fuselage attachments,
and deformation at reference points on the wing under design
load. Comparison between FE analysis and experimental results
is carried out to substantiates accuracy of these geometric simplifi-
cations and idealizations of load-carrying behaviour of structural
members. Therefore, practicality of these idealizations in absence
of design data is demonstrated. This study offers a novel approach
for structural assessments of aircraft without relying on proprie-
tary design data. The validated method enhances capability of
aircraft operators to perform effective structural analyses, thereby
extending service life of aircraft with continued airworthiness.1
Keywords: Finite element analysis, Structural integrity, Redu-
ced scale model, Structural idealization, Experimental validation
Resumen — Los operadores de aeronaves a menudo modifi-
can las configuraciones de las aeronaves, instalan nuevos equipos
y modifican las estructuras de los aviones para acomodar estos
equipos, lo que lleva a operaciones en envolventes de vuelo dife-
rentes al perfil de diseño original. Estas modificaciones requieren
evaluaciones estructurales de la estructura del avión, que normal-
mente requieren datos completos de diseño de la aeronave, que a
menudo no están disponibles para los operadores. Este estudio
tiene como objetivo desarrollar y validar un método práctico para
el análisis de elementos finitos (FEA) de estructuras de aerona-
ves en ausencia de estos datos de diseño detallados. Centrándose
en un estudio de caso que involucra el análisis estructural del ala
de un avión, este estudio presenta suposiciones e idealizaciones
utilizadas para desarrollar un modelo de elementos finitos (FE)
2.5D del ala. La fidelidad de este modelo se establece comparando
1. Aun Haider. Email: aunbhutta@gmail.com, ORCID: https://orcid.
org/0009-0000-5279-2829, Institute of Aeronautics and Avionics (IAA) Air
University Islamabad, Pakistan.
Manuscript Received: 13/07/2024
Revised: 22/08/2024
Accepted: 31/08/2024
DOI:https://doi.org/10.29019/enfoqueute.1080
los resultados del análisis FE con datos experimentales. Las mé-
tricas clave de validación incluyen fuerzas de reacción, distribu-
ción de carga en las uniones ala-fuselaje y deformación en puntos
de referencia en el ala bajo carga de diseño. Se lleva a cabo una
comparación entre el análisis EF y los resultados experimentales
para corroborar la precisión de estas simplificaciones geométri-
cas e idealizaciones del comportamiento de carga de los miembros
estructurales. Por lo tanto, se demuestra la practicidad de estas
idealizaciones en ausencia de datos de diseño. Este estudio ofrece
un enfoque novedoso para evaluaciones estructurales de aerona-
ves sin depender de datos de diseño patentados. El método vali-
dado mejora la capacidad de los operadores de aeronaves para
realizar análisis estructurales efectivos, extendiendo así la vida
útil de las aeronaves con aeronavegabilidad continua. pp. 11-18
Palabras clave: Análisis de elementos finitos, integridad estruc-
tural, modelo a escala reducida, idealización estructural, valida-
ción experimental.
I. INTRODUCTION
A. Research Problem
STRUCTURAL integrity analysis is paramount for safe-
ty, maintenance, and operational readiness of aircraft [1].
Structural integrity of aircraft is essential to prevent catastro-
phic failures that could lead to loss of life and equipment [2].
Moreover, structural integrity directly influences the frequency
and cost of maintenance operations, as well as overall readiness
of aircraft for their intended missions [3].
One of most significant hurdles in maintaining and asses-
sing structural integrity of aging aircraft is the lack of access
to comprehensive design data. This issue is exacerbated when
either original equipment manufacturer (OEM) is no longer in
business or has shifted focus to newer products [4]. For aircraft
procured from foreign countries, the situation is often worse,
with operators finding it virtually impossible to obtain neces-
sary design data when technology transfer restrictions are in
place [5].
The need for structural assessments arises from modifica-
tions made by operators to accommodate new equipment or to
meet changing mission profiles [6]. These modifications can al-
ter flight envelope and resultant structural loads, necessitating
detailed analysis to ensure continued airworthiness [7]. Howe-
ver, unavailability of design data including CAD models, finite
element (FE) models, material properties, and external loads,
poses a significant challenge [8].

ENFOQUE UTE, VOL. 15, NO. 4, OCTOBER 2024, pp. 11-18, E-ISSN: 1390-6542 12
Therefore, operators often rely on CAD models as templates
for creating FE models [9]. This process is labour-intensive,
requiring extensive geometric cleaning and discretization to
produce a model suitable for analysis. FE model must be de-
tailed enough to allow for comparison with actual deformation
results, while coarse enough to ensure a quick turnaround of
numerical results [10].
Moreover, as mission profiles often deviate from design pro-
files, and aircraft capabilities remain under-utilized or over-ex-
ploited [11]. This situation is further complicated when OEMs
withdraw customer support at the end of contractual agree-
ments, focusing instead on newer products [12]. Operators may
also be forced to keep aircraft operational beyond design life
due to procurement restrictions. Consequently, most operators
of aging aircraft lack technical support from OEMs, making it
challenging to keep them airworthy beyond design service life
[13]. Therefore, to ensure the continued airworthiness of aging
aircraft, structural assessments must be conducted [14]. The-
se assessments require access to comprehensive design data,
which directly impacts fidelity of the analysis. Access to accu-
rate and detailed design data is crucial for developing reliable
FE models, conducting thorough structural assessments, and
ultimately ensuring safety of the aircraft [15].
B. Research Hypothesis
In the absence of detailed aircraft design data, it is hypothe-
sized that a reduced-scale finite element model developed using
appropriate material properties, structural idealizations, and
computationally inexpensive finite element assumptions, can
accurately represent structural behaviour of the aircraft.
C. Research Objectives
The objective of this research is to establish validity of the-
se finite element (FE) idealizations invoked for analysis of an
aircraft wing. These idealizations are intended to be highly
practical, particularly when detailed aircraft design data is una-
vailable. This study presents a practical method for finite ele-
ment analysis (FEA) of aircraft in absence of design data with
reduced computational costs.
D. Section wise Organization of Document
In this research, FE model of a wing isolated from the fuse-
lage is presented. This model is developed using idealizations
proposed in this paper. The structural behaviour of FE model is
validated through comparison with experimental data. A positive
correlation between FE results and experimental data validates
the proposed assumptions. Significance of these assumptions lays
in correct structural behaviour predicted by underlying FE model.
II. LITERATURE REVIEW
A. Existing Relevant Literature
An aircraft wing is a semi-monocoque structure designed
to resist and transmit aerodynamic forces to the airframe [16].
The wing is statically indeterminate due to redundant structu-
ral members. Therefore, resulting structural response of each
member depend on the stiffness of adjacent members [17].
Outer skin of the wing encloses three different types of struc-
tural members [18]. Beam-type structural members running
along the wing span are called spars. Longitudinal structural
members, which are considerably thinner compared to spars,
are referred to as stringers. The third type of structural member,
called ribs, is positioned along transverse chord direction [19].
Transverse ribs and longitudinal stringers are made from stam-
ped sheet metal, while spars are machined. These structural
members work together to support external aerodynamic and
inertial loads and transfer them to the airframe [20].
The skin transmits aerodynamic forces to both longitudinal
members (spars and stringers) and transverse members (ribs)
through plate and membrane action [21]. Along with longitudinal
members, the skin reacts to applied bending and axial loads. In
conjunction with transverse ribs, the skin reacts to hoop or cir-
cumferential loads due to internal pressurization. The skin also
develops shear stress that reacts to applied torsional moments [22].
Longitudinal members, including spars and stringers, pri-
marily resist bending and axial loads. They segment the skin
into smaller patches, which increases the buckling and com-
pressive failure stresses. They also help arrest crack growth in
the skin [23].
Transverse ribs maintain cross-sectional wing shape, distri-
bute concentrated loads and redistribute stresses around struc-
tural discontinuities [24]. Ribs also establish column length
for longitudinal members by providing end restraint, thereby
increasing buckling strength of these members.
B. Gaps in Existing Knowledge
Behaviour of wing and its structural members have been
explained in detail in existing literature. However, no general
guideline is available to FE analyst for developing wing mo-
dels for structural analysis [25]. Therefore, FE analyst tends to
use a variety of techniques ranging from simple beam model
to full scale 3D model with all installed components. Fidelity
and computational cost, thus, vary enormously between these
extremes [26].
C. Justification for New Research
In absence of comprehensive aircraft design data, a redu-
ced-scale finite element (FE) model is required that can deliver
high-fidelity results with quick turnaround time [27]. A redu-
ced-scale FE model is a simplified version of full-scale finite
element model, based on idealization of structural members.
This model is designed to accurately capture aircraft’s struc-
tural performance while reducing complexity. This approach
facilitates timely decision-making and ensures that structural
assessments are both accurate and efficient.
III. METHODOLOGY
Idealization of load-bearing behaviour of an aircraft wing
is presented for developing a reduced-scale finite element (FE)

13
model [28]. It involves simplifying geometry, using shell and
beam elements for thin-walled structures, applying averaged
material properties and focusing on representative load cases.
MSC Patran® and Nastran® are used for FE analysis of wing
model [29]. Validation of reduced-scale FE model against avai-
lable experimental data or benchmark case is carried out to
substantiate accuracy of these idealizations.
IV. FE ANALYSIS OF WING
A. Idealization of Wing
The wing of an aircraft is attached to fuselage at four diffe-
rent locations through spars, designated as Front Wall (FW),
Front Spar (FS), Main Spar (MS), and Rear Spar (RS) [30].
Only the placement and limited geometric details of the-
se structural members are available in maintenance manuals
(MM). This information is utilized to develop 2.5 D FE model.
Several assumptions regarding the load-carrying capacity of
the wing’s structural members have been made [31]:
1. Longitudinal stiffeners and spar flanges carry only axial
stresses.
2. Rib web, skin, and spar web carry only shear stresses.
3. Axial stress is assumed to be constant along cross-sec-
tion of each longitudinal stiffener (spars and stringers).
4. Shear stress is assumed to be uniform throughout the
web of ribs and spars.
5. Transverse frames (ribs) are considered rigid within their
own planes and have no rigidity normal to their planes.
The structural members of wing have geometric details,
including lightening holes [32], variations in thickness, cross-
sectional warp, and manufacturing artifacts like fillets, cha-
mfers, and radii. These features are not included in the reduced-
scale FE model. Following geometric simplifications have also
been carried out:
1. Using average thickness for structural members.
2. Assuming no warp in the cross-sectional shape.
3. Omitting fasteners such as bolts and rivets, with load
transfer between adjoining members ensured through
coincident nodes [33].
Various components installed inside the wing, such as fuel
transfer valves, hydraulic actuators, landing gear attachments,
and electrical ancillaries contribute to 30 % mass and inter-
nal volume of the aircraft wing [34]. These components do
not contribute to structural stiffness of the wing. Additionally,
flight control surfaces attached to the wing, including airspeed
brakes, leading edge flaps (LEF), trailing edge flaps (TEF), and
ailerons, are required for aeroelastic analysis. The present stu-
dy deals with static structural analysis whereby these ancillary
component and control surfaces do not add structural stiffness
and hence, are not included in the model [35].
The purpose of this reduced-scale FE model is to calculate
internal load distribution, load paths, structural deformation,
and free-body loads [36]. The model uses 0D mass and spring
elements, 1D beam elements, and 2D shell elements [37] arran-
ged in 3D space to mimic the wing structure. Fig. 1 presents the
illustration of wing and placement of internal members in wing.
Fig. 1. Wing Model
B. Finite Elements Selection
3D solid elements are often unsuitable for modelling thin-
walled aircraft structures due to the phenomenon of shear loc-
king [38]. This issue can be mitigated by selecting first order
2D elements with appropriate mesh density (element size).
Nastran Element Library recommends using shell elements
(CQUAD4) and beam elements (CBEAM) for plate and beam-
like structures, respectively [39]. For structures where cross-
section remains constant along the length, lower-order CROD
element can also be used as an alternative to beam elements.
CQUAD4 (linear 2D shell) elements is used to model aircra-
ft skin and webs of ribs / spars. Each node in a shell element
has 5 degrees of freedom (DOFs), while each node in a beam
element has 6 DOFs. Flanges of ribs and spars, which carry
axial loads, are modelled using beam elements. Stiffeners in
skin panels and stringers in aircraft wing are modelled using
1D rod element CROD. This modelling approach balances
computational efficiency with the need for accurate represen-
tation of the aircraft’s structural behaviour under various loads
[40]. Fig. 2 shows the finite element model of the wing with
outer skin removed.
C. Material Properties
In aircraft maintenance manuals, except for nomenclature,
material properties are often not provided. Due to lack of spe-
cific material details, properties available from open resources,
as listed in Table 1, have been considered for the analysis.
Fig. 2. Internal Members in FE Model

TABLE I
MATERIAL PROPERTIES FOR WING
Component Material E (GPa)
Poison
Ratio(µ)
Front Wall
Al 2000 series 73.1 0.33
Front Spar
Box Beam
Rear Spar
Ribs
Stingers
Main Spar Steel 196 0.3
Skin Al 7000 series 70 0.3
D. Systems of Units
MSC Patran Nastran is independent of unit system and the-
refore, consistency of units is the responsibility of FE analyst.
For current analysis, mm, kg, s unit system is used. So, defor-
mation output is mm and stress output is MPa.
E. Boundary Conditions
The wing is connected to a root beam which is attached to
four transverse bulkheads of the fuselage. These bulkheads are
modelled using beam elements (CBEAM) with a very high
stiffness of 1 GN/m which are fixed at aircraft centreline (CL).
Fig. 3 illustrates these boundary conditions applied on the wing.
Fig. 3. Applied Boundary Conditions on Wing
Use of very stiff beam elements to represent wing-fuselage
attachment offers two main benefits. [41] First, it ensures that
the deflection of the wing spar attachments remains minimal,
allowing the wing deformation from numerical results to clo-
sely match experimental data. Second, by applying the fixed
boundary condition away from the spar attachments, it helps to
avoid stress singularities at these attachment points.
F. Loads
Available experimental results were obtained by applying dis-
crete forces to the wing through hydraulic actuators, with applied
load set equal to design limit load for the wing. This load is si-
mulated in finite element (FE) model by applying nodal forces at
the locations corresponding to the actuators. Fig. 4 illustrates the
application of these nodal forces on FE model of the wing.
Fig. 4. Applied Load on Wing
G. Verification of FE Model
To ensure adequacy of finite element (FE) model, following
steps have been implemented [42]:
1. A mapped mesh approach is used for development of FE
model of the wing.
2. More than 95 percent of shell elements are quadrilate-
ral. Triangular shell elements are employed only in mesh
transition regions to maintain model consistency.
3. Edge length for shell elements is set to 50 mm. Mesh
density is adjusted to ensure at least four shell elements
along beam cross-sections of ribs and spars.
4. Coincidence of nodes between adjoining structural
members is enforced to ensure effective load transfer
throughout the model.
5. Using default settings in MSC Patran, mesh quality
checks (including taper ratio, skewness and warp) are
conducted and confirm no errors.
6. Shell normal and beam orientations are verified to ensu-
re consistency within FE model.
7. No duplicate elements or free edges are present in FE
model.
These measures collectively ensure that FE model accurately
represents wing structure to produce reliable simulation results.
H. Static Structural Analysis
Finite Element analysis has been performed at design load.
Pre- & Post processing is performed in MSC Patran while MSC
Nastran is used as solver. Deformation field of wing under
applied load is given in Fig. 5. Von-Mises equivalent stress in
wing at design load is given in Fig. 6.
Fig. 5. Wing Deformation Field

15
I. Validation of FE Results
Validation of the finite element (FE) results has been per-
formed using experimental data to substantiate the proposed
assumptions and idealizations for the FE analysis of the wing.
The available experimental data includes:
1. Reactions (Forces and moments) measured at the wing
attachments.
2. Measurements of wing deformation at various locations
along the span under design load.
By comparing these experimental data points with the results
obtained from FE analysis, accuracy and reliability of underlying
assumptions and idealizations of the model are assessed.
Fig. 6. Wing Stress Field (Outer Skin Removed)
V. DISCUSSION
A. Interpretation of Results
Table 2 presents comparison of reaction (forces and mo-
ments) at wing attachment for FE and experimental results.
Both experimental and FE results corelate because maximum
percentage difference between these results is less than 6 %.
TABLE II
COMPARISON OF REACTION FORCES
Attachment
FE Results
Experimental
Results
%Age Difference
Force
(N)
Moment
(N.m)
Force
(N)
Moment
(N.m)
Force Moment
Main Spar 79354 124704 83520 129950 4.99 4.04
Front Spar 37578 58431 39810 61120 5.61 4.4
Rear Spar 33643 52334 35350 54870 4.83 4.62
Front Wall 2791 4263 2915 4435 4.25 3.88
Total 153366 239732 161595 250375 -- --
Table 3 gives load distribution among wing spars from FE
and experimental results. Both methods predict that main spar
takes 52 % load, front spar takes 24 % load, rear spar takes
22 % and front wall takes 2 % load, approximately.
Comparison of deflection field of wing for FE and experi-
mental results have been carried out. Front wall, front spar and
rear spar run from wing root to wing tip. Fig. 7 shows the mo-
nitor points for which experimental deformation of wing under
design load is available.
TABLE III
LOAD DISTRIBUTION
Attachment
FE Results Experimental Results
Force % Moment % Force % Moment %
Main Spar 51.74 52.02 51.68 51.9
Front Spar 24.5 24.37 24.64 24.41
Rear Spar 21.94 21.83 21.88 21.92
Front Wall 1.82 1.78 1.8 1.77
Total 100 100 100 100
Fig. 7. Monitor Points on Wing
Fig. 8 shows the comparison of deformation of along Front
Wall, Front Spar and rear spar for FE analysis and experimental
results. Deformation field of wing available from both studies
corelate with each other.
Fig. 8. Comparison of Deformation along Wing Spars

B. Research Questions and Hypothesis
Validation of FE results confirms that the idealized beha-
viour of wing structure can be accurately assumed. It has been
demonstrated that geometric simplifications do not signifi-
cantly impact the deformation field of the aircraft structure.
Additionally, it has been validated that the fasteners can be
excluded from FE model, with load transfer effectively faci-
litated through coincident nodes. Effects of surface and heat
treatments on the mechanical behaviour of structural members
can be disregarded in FE model without compromising accu-
racy. Use of candidate material properties, rather than exact
material specifications, is acceptable for modelling structural
members in FE model. These findings provide comprehensive
answers to key research questions, substantiating the initial hy-
pothesis that such assumptions are valid for aircraft structural
analysis. The study demonstrates that idealized behaviour and
simplifications can be reliably used in FE modelling without
adversely affecting accuracy of results.
C. Placement of Results with Existing Literature
These findings are unique within existing literature, addres-
sing the limitations of both low-fidelity analysis and compu-
tationally expensive methods. It has been demonstrated that
useful and accurate results can be achieved by implementing
these idealizations and assumptions with minimal compu-
tational cost. This approach provides a practical solution for
analyzing aircraft structures, especially when detailed design
data is unavailable.
VI. CONCLUSION AND RECOMMENDATIONS
Comparison of reaction forces, load distribution, and defor-
mation field of the wing with experimental results validates the
methodology for development of FE model. This validation con-
firms that the following idealizations are useful for FE analysis:
1. Idealized behaviour of structural members in the wing
can be assumed for static structural analysis. Longitudi-
nal stiffeners and spar flanges carry only axial stresses.
Rib web, skin, and spar web carry only shear stresses.
2. Geometric simplifications can be made in FE model,
omitting manufacturing features like fillets, chamfers,
and small cut-outs.
3. In absence of specific material nomenclature, reasona-
ble material properties can be used for components, and
effect of surface treatments on mechanical behaviour of
structural members can be ignored.
4. Fasteners can be excluded from FE model, with load
transfer effectively handled through coincident nodes.
5. Far-field boundary conditions can be applied to isola-
te structural members from the global assembly. This
boundary condition can provide an accurate approxi-
mation of their structural behaviour.
These validated assumptions streamline modelling process
and ensure accuracy of FE model without requiring exhaustive
detail, thus facilitating efficient and reliable structural analysis.
A. Contributions of Present Research
This research offers a methodology for developing a redu-
ced-scale FE model of aircraft structures. The reduced-scale
model can be developed using information accessible to air-
craft operators, without relying on detailed design data. This
model is particularly advantageous for achieving a quick tur-
naround of results during design iterations and modification
phases. This approach enables effective structural analysis and
decision-making, even the absence of proprietary design de-
tails, thereby supporting maintenance and modification efforts.
B. Benefits and Limitations of Proposed Solution
The proposed solution offers accurate predictions of defor-
mation and load transfer paths within the aircraft wing. Ex-
tension of this methodology for the development of FE model
of aircraft fuselage is also required to establish its robustness.
Further studies are also necessary to verify the stress results
obtained from the reduced-scale finite element model. These
additional investigations will help ensure the reliability and ac-
curacy of stress distributions, providing a more comprehensive
validation of the methodology.
C. Potential Applications
This research holds significant potential in the field of aerospa-
ce engineering. By employing proposed idealizations, a reduced-
scale finite element model can be developed, which effectively
captures the structural behaviour of the aircraft. In absence of de-
sign data, this model offers high-fidelity results at minimal com-
putational cost, making it a valuable tool for structural analysis,
design optimization, and modifications in aerospace applications.
D. Future Lines of Research
It is recommended that a strain gauge survey of the complete
wing under design load be conducted to verify the fidelity of
2.5D FE model for stress calculation. This experimental vali-
dation would ensure that the model accurately represents the
stress distributions within the wing structure, providing a more
comprehensive assessment of its reliability and accuracy.
ACKNOWLEDGMENTS
The author acknowledges the facilitation of his department at
Air University for providing all the resources for this publication.
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