Factor of Safety and Stress Analysis of Fuselage Bulkhead Using Composite

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Vol.04,Issue.32,

August-2015,

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Factor of Safety and Stress Analysis of Fuselage Bulkhead using Composite

Materials B. MANIDEEP

1, M. SATYANARAYANA GUPTHA

2

1PG Scholar, Dept of Aeronautical, JNTU, Hyderabad, TS, India, E-mail: [email protected].

2HOD, Dept of Aeronautical, JNTU, Hyderabad, TS, 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& 45

0 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.

Keywords: Safety and Stress Analysis, Composite Materials.

I. INTRODUCTION

A. 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 (for commercial 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.

B. 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.1 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,

B. MANIDEEP, M. SATYANARAYANA GUPTHA

International Journal of Scientific Engineering and Technology Research

Volume.04, IssueNo.32, August-2015, Pages: 6390--6396

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 .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.

B. Fuselage Detail Design 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.2. 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.3.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

Factor of Safety and Stress Analysis of Fuselage Bulkhead using Composite Materials


Volume.04, IssueNo.32, August-2015, Pages: 6390-6396

carry the load that the skin has given up. Tear straps are

located between each frame station Fig.4. 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.

Fig.4.Waffle doubler design instead of fail-safe straps.

Fig.5.Effect of sidewall frame depth.

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) as shown in Fig.5. Examination of the various

typical fuselage configurations of commercial airplanes

reveal that they are basically similar combinations of the

skin-stringer-ring(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.

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 as shown in

Fig.6. Frames also act as circumferential tear strips to ensure

fail-safe design

Fig.6. 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

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.




C. Fuselage Loads Loads affecting fuselage design can result from flight

maneuvers, landings or ground handling conditions as shown

in Fig.7. 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.

Fig.7.Fuselage divided into three sections.

Fore body: that portion of the fuselage forward of

the forward main frame.

Aft body: that portion of the fuselage aft of the aft

main frame; including the empennage.

Center body: that portion of the fuselage between

main frames.

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 fuselage to obtain the

gross weights and center of gravity locations on the c.g.

envelope.

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.

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.

External Pressures: External pressures on the fuselage,

other than in wing vicinity, are usually significant only

around protuberances as shown in Fig.8. 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.8.pressure distribution on fuselage.

Internal Pressures (Cabin Pressure): 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

Provide an 8000 ft. altitude cabin pressure

The max Flight altitude at 43,000 ft.

The pressures used with flight and ground conditions on

the airplane are:

TABLE I: Pressure Conditions On Flight And Ground

In addition, a pressure of 1.33*8.85=11.75psi is considered

to act alone.

II. INITIAL SZING

A. 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.

The thickness of stringer can be as same as the skin.

The width of flange is normally 16 times than its thickness

as shown in Fig.9.

Fig. 9.1initial sizing of Zed shape stringer.

B. Calculations of Skin Thickness Using Denis Howe’s

Methods Initial sizing of skin thickness using the condition of

pressurized skin:

(1)

Where

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 )

=allowable stress of standard value =100MN/m2

External pressure at 10,000 altitude is 0.2 bar (from

appendix) or 2.6550 104N/m2.

tp = 1.5 mm (approximately)

TABLE II: General Characteristics

TABLE III: Material Properties




Fig.10.fuselage structural assembly section view.

Fig.11. Fuselage Structural Assembly.

III. RESULTS & DISCUSSION

A. Results Before Modification

TABLE IV: Results of Static Analysis

TABLE V: Results of Modal Analysis

B. Results After Modification

Table VI: Results of Static Analysis

Table VII: Results of Modal Analysis

Fig.12.stress before modification.

The above table IV and V 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 as shown in Figs.10 to 12. 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 table V & VI

which are more satisfactory than the results before

modification. The analysis of results for different material

combination is shown in table VII.




Fig.13. stresses after modification.

IV. 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& 45

0). 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.

V. 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.

Documents

Factor of Safety and Stress Analysis of Fuselage Bulkhead Using Composite