Factor of safety and stress analysis of fuselage
bulkhead using composite materials
1
B.Manideep, 2M.Satynarayana guptha
1
M.tech student of JNTU, Hyderabad, TG-India,manideepbalusani.9@gmail.com
2
HOD, Aeronautical Dept, JNTU Hyderabad, T.G-India
ABSTRACT
The fuselage is an aircraft's main body section that serves to position control and stabilization
surfaces in specific relationships to lifting surfaces, required for aircraft stability and
maneuverability. It holds crew and passengers and cargo. In single-engine aircraft it will also
contain an engine, although in some amphibious aircraft the single engine is mounted on a pylon
attached to the fuselage. Along with different types of loads, pressure loads are to be considered
as very important which can be overcome by the skin and structural loads overcome by the help
of bulkhead and other structural members like stringers, formers etc. the main theme is to
provide safety with high reliability which is measured by the factor of safety. Now a day the
structural strength has been improving using different composite materials. The composite
material mainly reduces the weight of structure and increases ability to with stand at high load
operating conditions. An analysis involves with the finding of deformations and finding stresses
at general load conditions with different combination of aluminum alloys (AL 2024 T4, AL 6061
T6) and composite materials (Rein forced carbon fiber with 900& 450 orientation) for skin and
fuselage structural members respectively, and involves with modification of design according to
the analysis of results to improve the factor safety and reduce the stresses.
1.1 FUSELAGE CONFIGURATION
Ground rules:
 Analyze the difference in various
structural arrangements in terms of:


Producibility
Structural efficiency

Weight

Provisions for ducting and control cables

Effect on fuselage outside diameter

Noise attenuation
airplanes)
 Establish the impact of a future version
on the structural arrangement.
 Compare the fuselage structural
arrangement with other aircraft that are
applicable.
 To obtain a minimum structural wall
depth.
(for
commercial
1.2 General Requirements
General requirements namely, to transport a
payload, people, at subsonic speeds in a
comfortable environment. This dictates that
the fuselage must be pressurized to provide
a comfortable environment for the
passengers. The figure below shows the
most efficient pressure-carrying structure
that has a cylindrical cross-section with
spherical end caps.
The efficient pressure structure is
compromised (fig. As shown in 2) to satisfy
the aerodynamicist. The aerodynamicist is
compromised by operational requirements
for visibility. It must be obvious by now that
compromise is the name of the game. From
a structural standpoint, passengers are the
worst possible payload. Every cutout or
opening in a pressurized structure is a
compromise. Fig 3 shows the effect of
passenger requirements. They must have
doors to get in, windows to look out, and
more doors to load food for their comfort,
more doors to load their luggage, and still
more doors so they can get out in a hurry in
emergencies.
Finally, after providing space for a
weather radar antenna and large cavities for
the wing and landing gear, the fuselage
configuration is established, not completely
but in profile anyway. As can be seen in
fig.4, the only detail remaining of the
original efficient structural pressure vessel is
the closure bulkhead at the aft end. And
space considerations have compromised this
to a curvature flatter than a true spherical
shape.
With configurations established it is
possible to determine the primary structural
requirements. The primary flight loads
applied to the fuselage are shown in fig
below. They are lift, thrust, and pitching
moments applied by the wing, and
maneuvering tail loads from empennage.
These loads should combine along with
pressure in the proper combinations. It
should be pointed out here that these loads
represent only flight and environment
conditions. Secondary loads or those
loadings associated with the function or
utility of the airplane. Impose additional
structural requirements on the fuselage
structure. Some of the more significant of
those secondary loads come from large
equipment such as galleys, passengers and
their seats, baggage, and cargo, which must
be restrained not only for normal flight
conditions, but in most cases for extreme
loadings imposed under crash conditions.
The
final
consideration
in
configuration design is the cross section for
passenger transport which is predominantly
influenced by passengers. Fig shows double
lobe cross section to
Fig 1.1 Typical double lobe cross-section
Satisfy passengers on the upper lobe and use
enough room for cargo volume at the lower
lobe. This cross section is a common one for
narrow body transport of 5 to 7 abreast
seats; more than 8 abreast usually go to a
single circular cross section.
1.2 FUSELAGE DETAIL DESIGN
A) Skin and Stringers:
The largest single item of the fuselage
structure is the skin and its stiffeners. It is
also the most critical structure since it
carries all loads due to shear,fuselage
bending,and torsion and cabin pressure. This
primary loads are carried by the fuselage
skin and stiffeners with frames spaced at
regular intervals to prevent buckling and
maintain cross-section.
NOTE: that there has been very little
change in the basic structural concept, since
the earliest metal stressed skin airplanes,
These skin or stiffeners combinations have
proved over the years to be light weight,
strong structure that is relatively easy to
produce and maintain.
The most efficient structure is the
one with the least number of joints or
splices, therefore skin panels are as large as
possible, limited only by available mill
sizes. Stringers, being rolled from strip
stocks are limited in length by
manufacturing techniques.
Single lap splices are typically used for the
longitudinal skin joints. This is the lightest
design and does not impose a severe
aerodynamic penalty on a subsonic airplane.
The transverse splices, those normal to the
air stream, are flush-butt splices because a
lap step here would have an appreciable
effect on boundary layer turbulence and
drag.
Stringers splice locations are
established by another set of rules, since the
skin and stringers are working together; they
should both be spliced at the same location.
This maintains the relative stiffness of the
stringer or skin combination, which is
desirable from a fatigue stand point.
Fig 1.3 Fuselage transverse splice of skin
and stringer
The fuselage cross section in the
ideal shape is of a true cylinder such as the
L-1011,DC-10 and A300 transports, and the
cabin pressure loads are carried by hoop
tension in the skin with no tendency to
change shape or induced frame bending.
The two most common fail-safe
design concepts are breaking the component
down into several small overlapping pieces
where, if one fails, its loads can be carried
by adjacent parts, or utilizing a restrainers or
fail-safe strap that will contain a failure with
incontrollable limits. The latter method is
applied in the skin of transport design where
the critical design loading is pressure.
Fig 1.4 Fail-safe strap located between
skin and frame
The above fig shows a typical skin,
stringer, fail safe strap, and frame
attachment. The tear strap, which is riveted,
spot wielded or bonded to the skin, is sized
such that a skin crack mill stop when it
reaches the strap and the strap can carry the
load that the skin has given up. Tear straps
are located between each frame station
Fig.9. It has been proved by the test that a 20
to 40 inch crack can be sustained without a
catastrophic failure. In the areas where the
skin thickness is determined by bending
loads, the stress level from hoop tension is
low enough that fatigue is not critical.
The discussion of skin, stringer, and
fail-safe pressure design has been somewhat
short compared to its importance. Most
thought, study, and testing has gone into this
phase of the structure than any other because
poor design details in this area are uniform
giving.
Utilization of the airplane power
over shorter route segments means that
fatigue is a primary design consideration.
The lower operating pressure differential
means the minimum gage skins could be
used in the hoop tension areas if a
satisfactory design for fail safe crack length
control could be developed.
Fig1.5: Waffle doubler design instead of
fail-safe straps
Shown are fuselage skins with
doublers which are bonded to the skin to
extend fatigue life. Minimum thickness
(0.036 inch) skins have a waffle pattern
doubler bonded on them, a plane product is
the elimination of the stringer joggle where
the stringers have to step over individual
fail-safe straps (or tear straps).
Examination of the various typical
fuselage configurations of commercial
airplanes reveal that they are basically
similar combinations of the skin-stringerring(frame) structures, with the interior trim
line (one inch greater than the frame
depth)providing an overall cabin wall
thickness. Ring or frame spacing is in the
order of 20 inches and stringers spacing
varies between 6 to 10 inches. Commonly,
the passenger transport fuselage sidewall
(window and door area) design replaces
stringers with heavier thickness skins so that
a quieter cabin can be obtained and the skin
fatigue stress can be reduced because of
cabin pressurization cycles. In the sidewall
region, frame depth can be kept to a
minimum (provided that adequate working
space is not a problem) because no
significant concentrated loads are involved.
Above and below the side wall region,
ample space is available to profile for
increased frame depth as required. Another
advantage is to reduce fuselage diameter to
save structural weight and less fuselage
frontal area to reduce aerodynamic drag with
the
same
internal
width
across
constraint on the expansion of cabin shell.
Owing to their film sine conventional cabin
frames, whose main function is the
preservation of the circular shape against
elastic instability under the compressive
longitudinal loads, have little constraint on
the radial expansion of the shell.
Fig 1.6 Effect of sidewall frame depth
b) Frames and floor beam:
Fuselage frames perform many diverse
functions such as:
 Support shell compression/ shear
 Distribute concentrated loads
 Fail-safe (crack stoppers).
They hold the fuselage cross section
to control shape and limit the column length
of longerons of stringers. Frames also act as
circumferential tear strips to ensure fail-safe
design
Fig 1.7 Typical floor beam arrangement
In addition they distribute the
external and internal loads onto shell,
redistribute
shear
across
structural
discontinuities, and transfer loads at major
joints. Heavy cabin frames and bulkheads
represent extremes in the matter of radial
1.3 FUSELAGE LOADS
Loads affecting fuselage design can
result from flight maneuvers, landings or
ground handling conditions. Fuselage loads
are primarily a problem of determining the
distribution of weight, tail loads, and nose
landing gear loads. Weight distribution is
important because a large part of fuselage
loads stems from the inertia of mass items
acted upon by accelerations, both
translational and rotational. Tail loads,
which are generally quite large, contribute
heavily to bending the aft portion of the
fuselage. Dissymmetry of tail loads causes
significant aft-body torsions. In the same
sense that tail loads affect the aft-body,
loads acting on the nose landing gear will
contribute significantly to net loads on the
forward portion of the fuselage, the fore
body. As implied above, the various portions
of the fuselage can be most critically loaded
by completely different flight or handling
conditions. For expedience of analysis the
airplane is divided into three sections and
each of these is analyzed separately. Of
course, eventually the structure must be
considered for effect of loads carrying
through from one section to the other. For
discussion here, the fuselage will be divided
into sections in the same manner as is
generally applied in analysis. For a typical
airframe, the fuselage is divided into three
sections.
fuselage to obtain the gross weights and
center of gravity locations on the c.g.
envelope.
Fig 1.8: Fuselage divided into three
sections
1. Fore body: that portion of the fuselage
forward of the forward main frame.
2. Aft body: that portion of the fuselage aft
of the aft main frame; including the
empennage.
3. Center body: that portion of the fuselage
between main frames
1.3.1 Distribution of Weight
The fuselage weight distribution
consists of the fixed weight of the structure
and equipment, and the removable load. The
removable load in military types is relatively
small and concentrated. Passenger and cargo
airplanes, however, are required to carry
loads of varying quantity and location in the
fuselage. Because of the various possible
arrangements of the cabin for different
customers, the removable load (cargo and
passengers) is considered to include such
items as seats, galleys, lavatories, etc. An
arbitrary floor loading is then used for
transport airplane design. This arbitrary
loading is selected to envelope all possible
variations of cabin loading. The floor
loading for this equipment plus passengers
runs approximately 45 lb. per sq. ft.
Baggage
weight
is
satisfactorily
approximated by a maximum of 20 lb. per
cubic ft. of cargo space. The arbitrary
loading used is then distributed in the
1.3.2 Fore Body Loads
The fore body loads incurred during
symmetrical flight are obviously in a vertical
direction only; there are no side loads. This
vertical direction is normal to the airplane
reference axis, or in the "z" direction. These
vertical loads are determined simply by
multiplying the weight loads by the load
factor. Vertical air loads are generally
neglected in fore body loads calculations
except for wide body fuselage or their effect
on local structure. Neglecting them is
generally conservative because they are in
the direction to relieve inertia loads. Also,
they are usually small relative to net loads.
Another reason contributing to their
omission is the fact that accurate
distributions are difficult to determine
because of irregularities in the fuselage
profile. Good distributions would have to be
determined by wind tunnel pressure
measurements. The cost involved is usually
not
warranted.
Assumed,
linearized
distributions are sometimes used, but it
should be borne in mind that the probable
inaccuracy, coupled with the relative
magnitude of the air load, results in a load
component of questionable reliability. Side
loads (in the y direction) are caused by side
and yawing accelerations and air loads
incurred during unsymmetrical maneuvers.
Here the air loads make up a large part of
the net loads and therefore cannot be
neglected. As reasonable a distribution as
possible is estimated based on best available
data. Critical fore body loadings may also be
experienced from application of nose
landing gear loads. Design loadings might
arise from landing or application of main
wheel brakes during taxiing, particularly
unsymmetrical brake application.
1.3.3 Aft Body Loads
Aft body vertical flight loads are a
critical combination of inertia loads and
horizontal tail balancing loads. The
horizontal tail loads are determined for the
various conditions on the V-n diagram and
center of gravity locations. Since the
distribution of weight in the fuselage as well
as the tail loads are a function of c.g.
location, the problem is one of determining
the critical combinations. Lateral loadings
result from application of air loads acting on
the vertical tail in combination with side
inertia loads. Air loads on the fuselage aft
body is generally neglected both in the
vertical and side directions. In this case the
air loads are not necessarily relieving;
therefore it is not conservative to neglect
them. However, they are generally quite
small, and their distribution in the
unpredictable flow behind the wing is
impossible to determine.
1.3.4 External Pressures
External pressures on the fuselage,
other than in wing vicinity, are usually
significant only around protuberances. In the
area of the wing, the pressures on the wing
are carried onto the fuselage.The pressure on
the fuselage will be of the order of
magnitude of the pressure on the wing.
Fig 1.15: pressure distribution on fuselage
The fuselage internal pressure
depends on the cruise altitude and the
comfort desired for the occupants. Pressure
differential may be readily determined from
the altitude charts as shown if the actual
altitude and the desired cabin pressure are
known. Fuselage pressurization is an
important structural loading. It induces hoop
and longitudinal stresses in the fuselage
which must be combined with flight and
ground loading conditions. The important
consideration for establishing the fuselage
design pressures is the cabin pressure
differential or present in altitude and the
fuselage is designed to maintain.
Example: Assume an airplane
(a) Provide an 8000 ft. altitude cabin
pressure
(b) The max Flight altitude at 43,000 ft.
The pressures used with flight and ground
conditions on the airplane are:
Conditions Max positive
(burst)pressure
Flight
8.85psi
Max
negative
(collapsing)
-0.50psi
Ground
1.0psi
-0.5psi
Table: 1.1 Pressure conditions on flight and
ground
In
addition,
a
pressure
of
1.33*8.85=11.75psi is considered to act
alone
2.0 INITIAL SZING
.
1.3.5 Internal Pressures (Cabin Pressure)
2.1 Stringer thickness initial sizing by
denis howe’s method
The method to estimate the initial size of
stringer section in Denis Howe’s book is:
 The pitch of stringers is between 1.5
and 5 times the stringer height, it is
suggested that in the initial design
phase, this value can be assumed 3.5.


The width of stringer flanges can be
estimated 40% of the stringer height.
=allowable stress of standard value
=100MN/m2
External pressure at 10,000 altitude is 0.2
bar (from appendix) or 2.6550 104N/m2
(1 × 105 − 2.6550 × 104 ) × 1.975
𝑡𝑝 =
100
tp = 1.5 mm (approximately)
The thickness of stringer can be as
same as the skin.
The width of flange is normally 16 times
than its thickness.
Crew
Capacity
2-3 crew
members
150 passengers
Cabin length
27.51m
Fuselage width
3.95m
Height
11.76m
Maximum takeoff
weight
162900lb
Wing span
35.80m
Fig 2.1initial sizing of Zed shape stringer
2.2 Calculations of skin thickness using
Denis Howe’s methods:
Initial sizing of skin thickness using
the condition of pressurized skin:
𝑡𝑝 =
Table 2.1: General Characteristics
∆𝑃. 𝑅
𝜎𝑝
Where
∆𝑃 =P2 –P1
P2 is the internal pressure of 0.1bar
P1 is the external pressure to be calculated
R - is the radius of the skin = 3.95m ( from
table )
Fig 2.2 fuselage structural assembly section
view
Fig 2.3 fuselage structural assembly
Materials
Ex
(Pa)
Pois
on’s
ratio
Density
(Kg/m3
)
Yield
strength
(MPa)
Al alloy
2024 T4
73.1*
10^9
0.33
2780
324
Al alloy
6061 T6
68.9*
10^9
0.33
2700
276
Carbon
Fiber
Reinforced
Polymer
900
Carbon
Fiber
Reinforced
Polymer
450
44.9*
10^9
0.05
1760
474
11.2*
10^9
0.33
1750
149
Combinatio
n2
(al 6061
T6,corbon
fiber 900)
Combinatio
n3
(al 2024
T4,corbon
fiber 450)
Combinatio
n4
(al 6061
T6,corbon
fiber 450)
3.912
244.49
1.128
3.692
244.27
1.326
3.917
244.29
1.129
Table3.1 Results of Static analysis
Al
6061
T6
Al
2024
T4
Mode 1
68.631
Frequencies in Hz
Mode 2 Mode 3
69.918 99.62
Mode 4
119.28
67.61
68.87
117.51
98.138
Table3.2 Results of modal analysis
Table 2.2Material properties
3.0 RESULTS & DISCUSSION
3.2Results after modification
3.1Results before modification
Deformati Vonon
misses
(In mm)
stress
(in
MPa)
Combinatio
3.686
244.48
n1
(al 2024
T4,corbon
fiber 900)
Factor
of
safety
1.325
Deformation Von- Factor
(In mm)
misses
of
stress safety
(in
MPa)
Combination
2.351
175.21 1.849
1
(al 2024
T4,corbon
fiber 900)
Combination
2
(al 6061
T6,corbon
fiber 900)
Combination
3
(al 2024
T4,corbon
fiber 450)
Combination
4
(al 6061
T6,corbon
fiber 450)
2.494
175.26
1.574
2.347
173.73
1.864
2.491
173.84
1.587
Table 3.3 Results of Static analysis
The above table 3.1&3.2 showing
results for before and after modification of
design. By observing the data and stress
distribution it can be understood that the
stress on the skin are less, that means it has
more strength than required, so we can
decrease the skin thickness to some extent
and we can reduce the structural weight.
Second one is the bulkhead structure has
shown more deformations than skin so in
order to increase the strength we have to
increase the thickness of the bulkhead
structure. As a result of modification the
final results were tabulated in table3.3&3.4
which are more satisfactory than the results
before modification. The analysis of results
for different material combination is shown
in table 3.3
Frequencies in Hz
Al
6061
T6
Al
2024
T4
Mode 1
51.725
Mode 2
56.76
Mode 3
99.345
Mode 4
105.68
50.956
55.91
89.44
97.86
Table 3.4 Results of modal analysis
Fig 3.2 stresses after modification
4.0 Conclusion
The analysis of model was done with
four different material properties, (for outer
skin Al alloy 2024 T4 and 6061 T6, for
bulkheads carbon fibre 900& 450).
Fig 3.1 stress before modification
The main aim of this project is to
minimize stress. To reduce this stress three
methods had considered they are
 Design modification in model
 Material change in model
 Design modification and material
changes in model
The initial analysis before design
modifications has given best results for al
2024, carbon fiber 450 with von-misses
stress 244 MPa, and factor of safety 1.5.
After design modification by
comparing results of several above
mentioned
combination
of
material
properties, the modified design with 3mm
thickness (outer skin) ,45mm width(bulk
head) having materials Al 2024 T4, Carbon
fiber with 450 orientation has given best
results with von-misses stress 173.73Mpa
and high factor of safety 1.86 compared to
other materials
The Value result of Von-Misses
Stresses from the analysis is far than
material yield stress so the design is safe
with a factor of safety 1.86.
And the model performing very less
vibration compare to other model’s material
properties with minimum frequency of
50.956 Hz and maximum frequency of
97.867 Hz.
References:
1. Krakers, L.A., Multi-disciplinary design optimization of aircraft fuselage structures, PhD
Thesis, to be published, 2007.
2. Sullins, R.T., Smith, G.W., Spier, E.E., Manual for structural stability analysis of sandwich
plates and shells, NASA report,CR-1457, Langley, 1969
3. Heckl, M., “Vibrations of point-driven cylindrical shells”, Journal of the acoustical society of
America, Vol 34, nr 10, 1962.
4. Morse, P.M., Vibration and Sound, McGraw-Hill, New York, 1948.
5. Cremer, L., Heckl, M., Ungar, E.E., Structure born sound, Springer Verlag, Berlin, 1973.
6. Rocca, G La, & Tooren, MJL van, “Development of Design and Engineering Engines to
support multidisciplinary design and analysis of aircraft.” Proceedings Design Research in the
Netherlands, 2005.