Effect Of Fuselage Frame And Tear Strap In Arresting A Two Bay Crack In Fuselage Structure

Effect of Fuselage Frame and Tear Strap in Arresting A Two-Bay Crack

in Fuselage Structure

Syed Faizus Salam Quadri

_{Amabdas Kadam}

1

_{P.G. Student}

_{Assistant Professor}

_{Department of Machine Design}

1,2

_{Visvesvaraya Technological University Regional Center Kalaburagi}

Abstract— Currently large transport airplanes are being developed with “Large damage tolerance capability” as a design goal. An important concept in the design of the pressurized fuselage of large transport aircraft is the provision of crack stopper straps to arrest the fast fracturing of a crack. Current study includes the role of the fuselage frame and crack stopper strap in the fail-safe design of the fuselage. As a first approximation a stiffened flat panel with a center longitudinal crack is considered. The strength of this cracked panel is investigated as a function of crack length in the absence of crack stopper straps, but only with the frame present. Crack stopper straps are then introduced at the locations of stiffeners perpendicular to the crack line and strength of the cracked flat panel is investigated as a function of crack length in the presence of crack stopper straps. The bulkhead dimensions and the thickness of the crack stopper straps are varied in the parametric study.

Key words: Fuselage Frame, Tear Strap in Arresting, Skin

I. INTRODUCTION

Due to the various accident happened in the 1950’s it caused the designers at that to time to modify the aircraft design specially fuselage which can withstand deterioration caused by fatigue cracking i.e., damage tolerant design philosophy. A strengthened strap on the inner lining of fuselage skin termed as tear strap/crack stopper strap is generally put to use. A tear strap is generally are made up of aluminum alloy, which is attached circumferentially to the inside of fuselage between the skin of fuselage and bulkhead which holds both of them altogether, the purpose of the tear strap is as its name implied to helps in arresting the crack propagation in the skin of fuselage.

II. OBJECTIVE

The aircraft fuselage is subjected to different loads and boundary conditions. Some modifications are made to the fuselage to withhold the loads and resist the failure, the structural integrity is checked.

III. GEOMETRICAL SPECIFICATION

Stiffened panel is of size 2625x1575 taken for analysis. In the stiffened panel there are seven (7) bulkhead and nine (9) longerons used as reinforcing member.

Fig. 1. Dimensions of Fuselage, Bulkhead and Stinger

IV. MATERIAL SPECIFICATIONS

The material considered used for the entire structure is Aluminum Alloy – 2024-T351 with the following properties

Property Aluminum Alloy – 2024-T351

Young’s Modulus 73Gpa or 7000kg-f/mm2

Density 2.77kg/cm3

Ultimate tensile

strength 490Mpa

Tensile strength: yield 350Mpa Fracture Toughness 99Mpa√m

Shear Strength 285Mpa

Fatigue Strength 140Mpa

Strength to weight ratio 177 kN-m/kg Table 1:

V. LOAD CASE

Stiffened panels are the most common structure where the crack is developed in an aircraft. Presently large transport aircraft implement the “Damage Tolerance philosophy” as the design criteria. Tear straps are used for arresting the fast propagating crack.

A difference of pressure of 8 psi (0.055158 N/mm2 or 0.005625 kg/mm2) is considered for the present load case. A hoop stress will be produced because of internal pressurization of the fuselage (cabin pressure). The tensile load acting on the edge of panel will be considered for analysis. In this case a flattened stiffened panel is considered as mentioned above.

Where,

p= cabin pressure r= radius of fuselage

t=thickness of skin of fuselage

VI. STRESS ANALYSIS OF STIFFENED PANEL

The formula for stress analysis of the fuselage is given with FEM perceptive is given below,

Theoretical Stress intensity factor KIC = C.F x K0

K0= σR√𝜋𝑎

C.F = Correction factor = 1−0.5 a b+0.326(

a b)

2

√1−a_b

σR = Remotely applied stress = P/A

a = Semi-crack length FEM Method

KIC = √𝐺𝐸

Where G = Strain energy released = 1

2∆a∆v F t

Where Δa= Length of the element Δv = Displacement

F = Force on the edge elements

t = thickness of plate E = Young’s Modulus The results are tabulated below

Crac k lengt h a Correction Factor(C.F .) Theoretic al KIC, Mpa√𝒎 Strain Energy Release d G FEM KIC, Mpa √𝒎

10 1.001 7.013 0.00669

9 6.99

20 1.004 9.945 0.01421 9.97 30 1.011 12.265 0.02167 12.31 40 1.02 14.289 0.02954 14.38 50 1.033 16.179 0.0379 16.31 60 1.05 18.015 0.04722 18.1815 70 1.07 19.829 0.05751 20.0641

80 1.1 21.792 0.06919 22.00

90 1.134 23.829 0.0827 24.06 100 1.175 26.026 0.12809 29.94

Table 2:

VII. BULKHEAD MODIFICATIONS

The bulkhead dimension of the middle part is increased from 50mm to 80mm to see the behavior of its and its role in arresting the crack. The values after the analysis are tabulated below.

Sl no Crack length a Strain Energy Release, G Stress Intensity Factor KIC

1 25 0.205 37.92

2 50 0.390 52.23

3 75 0.559 62.56

4 100 0.720 71.01

5 125 1.078 86.87

6 150 1.264 94.07

7 175 1.450 100.74

8 200 1.636 107.01

9 225 1.822 112.93

10 250 2.006 118.5

11 275 2.185 123.66

12 300 2.351 128.27

13 325 2.493 132.09

14 350 2.592 134.71

15 375 2.639 135.9

16 400 2.643 136.01

Table 3:

Fig. 1. Graph 1 KIC After Bulkhead Modification. The Residual Strength is as follows .

Semi Crack Length a in mm

Residual Strength (Mpa) of

Skin Bulkhead Rivets

25 10.41 18.29 57413

50 7.56 18.29 57413

75 6.31 18.13 57413

100 5.56 17.96 45930

125 4.54 17.34 25517

150 4.2 16.91 177665

175 3.92 16.49 12758

200 3.69 15.97 8505

225 3.5 15.36 5888

250 3.33 14.69 3588

275 3.19 13.97 1837

300 3.08 13.15 890

325 2.99 12.19 516

350 2.93 11.18 800

375 2.9 10.11 166

400 2.9 9.15 160

Fig. 3. Skin.

Fig. 4. Bulk Head.

Fig. 5. Graph 4 Graph showing Residual Strength.

VIII. TEAR STRAP INTRODUCTION

Lastly there are some modification made in the stiffened panel in obstructing the crack growth, for that two major modifications are made one being changing the dimensions of bulkhead and another be introduction of crack stopper in the stiffened panel. By introducing the cracks stopper will not only aid in arresting he crack but also increase the residual strength of the entire structure. The bulkhead changes are already stated in the geometrical specifications as well as the tear strap dimensions. The position of the tear strap is as shown in figure.

Fig. 6. Location of Tear Strap. The SIF values are

Sl no

Crack length

a

Strain Energy Release, G

Stress Intensity Factor KIC

1 25 0.132 30.44

2 50 0.285 44.68

3 75 0.396 52.64

4 100 0.505 59.47

5 125 1.610 106.15

6 150 1.828 113.15

7 175 2.044 119.6

8 200 2.256 125.67

9 225 2.463 131.3

10 250 2.657 136.39

11 275 2.831 140.78

12 300 2.972 144.23

13 325 3.030 145.64

14 350 2.959 143.92

15 375 2.352 128.32

16 400 1.628 106.77

Fig. 7. Graph 5 KIC after tear strap introduction. The values of residual strength is

Semi Crack Length a in

mm

Residual Strength (Mpa) of

Skin Bulkhead Rivets

25 127.22 22.04 346.39

50 86.67 20.78 343.79

75 73.56 20.66 340.23

100 65.12 20.51 334.77

125 36.48 18.98 316.76

150 34.24 18.46 302.58

175 32.38 17.96 286.35

200 30.81 17.34 268.29

225 29.49 16.63 248.01

250 28.39 15.97 225.59

275 27.51 15.13 200.75

300 26.85 14.17 174.78

325 26.59 13.24 170.62

350 26.91 12.12 162.99

375 30.18 10.99 22.41

400 36.27 10.16 8.65

Table 5: Residual Strength after Tear Strap .

Fig. 8.

Fig. 9.

Fig. 10. Graph 6 Graphs Showing Residual Strength after tear strap introduction.

IX. CONCLUSION

– Damage tolerance philosophy is most extensively used in the design process of aircraft to cut down the weight of the aircraft.

– Initially the stiffened panel is subjected to load with the basic configuration.

– The techniques FEA uses for analysis is Virtual Crack Closure method (VCCT)

– After that modifications are made to the stiffened panel by increasing the size of bulkhead. In t his circumstance, the SIF intensity increases after some increment i.e., 175mm crack length. – The residual strength also decreases of the

component that is skin, bulkhead and rivets, the values of rivets are extremely high which are impossible to have practically, and thus these values are just theoretical only for comparison. – Next, the alteration of stiffened panel is addition

of tear strap to the structure. The tear strap helps to arrest the propagation of crack as obvious from the table and graph. The sudden fall of graph ensures that the addition of crack stopper is helpful. The SIF values and graph are provided for proof

REFERENCES

[1] V.J. Anand and P.C. Arun Kumar, “Evaluation of Crack Arrest Capability in the Stiffened Panel of a Transport Aircraft” ICCOMIM, pg#299- 305, July 2012

[2] P. M. S. T. de Castro, S. M. O. Tavares, V. Richter-Trummer, P. F. P. de Matos, P. M. G. P. Moreira, L. F. M. da Silva, “Damage Tolerance Of Aircraft Panels” (Mecanica Experimental, Vol 18, pg# 35-46,2010