Technical Paper - Ensuring Airworthiness for C-130J-30 Aft-Plug Upper BL61 Longerons

8/12/2019 Technical Paper - Ensuring Airworthiness for C-130J-30 Aft-Plug Upper BL61 Longerons

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Ensuring Airworthiness for C130J30AftPlug Upper BL61 Longerons

Dr Hugh Stone, Australian Aerospace (AA) C130J Engineering

1. Abstract This paper describes work done to ensure the ongoing structural integrity of aft plug upper BL61 Longerons in

RAAF C130J 30 aircraft. The work arose out of finding multiple instances of damage in the RAAF fleet. Whilst the damage itself was not fatigue related, the repairs required an assessment of their life as some of the damage

details were well outside the bounds of normal good practice structural design. The main aspects of the work include the development of a detailed FEM model to recover bearing and bypass loads in the joint; the

development of a suite of maximum damage models to bound the fleet wide problems; and the use of damage

tolerance analysis to calculate inspection intervals for the fleet, based on these maximum damage models and conservative spectrums. This has now led to alterations to the structural inspections for these components. This

safetyby inspection approach has resulted in total cost savings estimated at 3700 man hours of work, or 62%, in comparison to longeron replacement.

1 Introduction During depot level maintenance for a C130J 30 aircraft in 2010, severe damage to holes in the Aftplug upper BL61 longerons was discovered. The damage consisted of a mixture of oversized, elongated or double drilled holes, often with low edge distance. A picture of one of the damaged longeron ends on the first aircraft is shown in Figure 1. The

damage led

to

longeron

replacement

and

the

issuing

of

an

STI

(Service

Bulletin)

to

inspect

other

aircraft.

As

a result

of this inspection the majority of RAAF aircraft were found with similar instances of damage. Full longeron replacement is expensive as it is hard to procure parts and requires a large number of man hours. To obviate the need for full longeron replacement across the fleet, the possibility of cleaning out the damaged holes by use of oversize fasteners or freeze plugs was investigated. As the degree of over sizing was extensive (increases of 6 oversizes: i.e. 6/64ths: 0.250 0.344), and remaining edge distances were often less than 1D, a Damage Tolerance Analysis (DTA) was conducted to determine appropriate inspection intervals for the repaired structure.

Figure 1 Damage at FS817E R/H and 737E L/H on 1st aircraft: the Longerons were replaced in this case. As can be seen damage consists of double drilled holes, which require significant over sizing to clear them out.

Aircraft Airworthiness and Sustainment Conference, Brisbane 2012 1 of 14


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The main features of the analysis were the development of a robust model of the joint which would allow the extraction of accurate bearing and bypass loads at the fastener locations; the use of simplified Severity Factor calculations to winnow the different damage conditions down to a limited set of critical conditions; and the use of the DT tool AFGROW to model crack growth in the longeron outstanding flange for the critical conditions under conservative spectrums. Although the analysis received the OEMs final blessing, it was performed entirely at AA,

without any special tools or OEM data above what is available to the RAAF. The analysis found that the repairs were acceptable, provided that slight enhancements to the longeron inspection regime were promulgated. This has resulted in the saving of thousands of hours of effort in changing longerons on multiple aircraft.

2 Description of Structure The C130J 30 Hercules is a 164,000 lb GW aircraft powered by four RollsRoyce AE2100 engines with a 132 7 wingspan. The basic structure of the fuselage consists of 6 major longerons and a structural cargo floor, (See Figure 2). The external skin of the aircraft is only considered to transmit shear and pressure loads and is assumed ineffective in compression. Unlike many aircraft the Hercules does not have a multiplicity of stringers around its fuselage which leaves the longerons and the cargo floor the only structure available to resist primary fuselage bending loads.

Figure 2: Fuselage Structure of C130J 30 aircraft showing primary structural elements for resisting fuselage bending moments, (base picture from www.flightglobal.com)

The location of damage observed on RAAF aircraft is in the aft 80 plug. The plug longerons are joined to the continuing structure fore and aft via back to back double tension fittings. These in turn are joined to the longeron via 5 x 0.250 HiLoks through the longeron outstanding flange as well as 14 x 0.164 HiLoks through the longeron skin flanges. At the joint the longeron is a Tshaped extrusion with the middle leg of the T pointing inwards and the top of the T attached to the aircraft skin. The two tension fittings nestle in either side of the T as shown in Figure 3.



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3 Description of Damage The damage consisted primarily of oversized, elongated or double drilled holes, often with low edge distance. The location of damage observed on RAAF aircraft was at the 5off 0.250 HiLoks in the longeron outstanding flange. The damage location can be seen in Figure 3. Although damage was observed in both the fittings and the longeron the fittings were relatively easy to replace. This paper will only focus on the longeron.

Figure 3:

FS817E

Upper

BL61

Longeron

Tension

Joint,

with

Lower

Tension

Fitting

Removed

for

Clarity.

(Note

that

the

Longeron

skin flange is not visible in this picture as it sits outside the skin of the aircraft.)

Besides the damage shown in Figure 1, another example of damage is also shown in Figure 4. In this case the damage consists of oversize holes with low edge distance.

Figure 4: Further damage observed in this case oversize holes and low edge distances. (The nominal configuration is 0.250 holes with 0.370 edge distance).



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Repairs to the longerons consisted of the use of oversized fasteners and freeze plugs to clean out the damaged holes. This paper will not consider the causes of the damaged holes: rather it will focus on the determination of an inspection regime for the repaired structure.

4 Sorting Through Damage Cases A multi stage process was used to sort through the different damage scenarios encountered to arrive at a general inspection regime for all aircraft. Rather than developing individual inspections for each aircraft, it was decided that a general inspection would be more appropriate. This simplified both the analysis and the implementation of the inspection regime and saved significantly on engineering time to determine appropriate inspection intervals. The first part of arriving at a common interval for all aircraft was to find a worst case damage scenario that could be used to set the inspection regime for the fleet.

This was done using a hierarchical assessment as shown in Figure 5. For all damaged aircraft, the damage at each

location in the aft plug (four possible locations: BL737E L&R, BL817E L&R) was initially assessed qualitatively. If the damage was obviously non critical in comparison to other aircraft, no further action was taken other than the development of a normal static repair. If the damage could not be ruled as non critical from this eye ball assessment, it was further analysed using a detailed FEM Model (See Section 6) to obtain fastener loads and bypass stresses in the longeron. The fastener loads were then used with a Niu1 Severity Factor style calculation to obtain a peak stress at each hole in the joint. These peak stress results were then compared and only the largest set aside for DT analysis using AFGROW.

Figure 5: Flow Chart for damage assessment



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5 Maximum Damage Models At the time of the analysis, many aircraft had not been inspected. It was thus thought appropriate to try and bound the worst case damage we would expect to see based on the initial aircraft inspected. Inspection intervals for such a worst case damage model could then be used as a bounding solution for the fleet. The worst case damage models focussed on the following factors.

Edge distance. Low edge distance causes higher stress concentrations, higher factors and reduced crack growth life in the ligament between the hole and the edge. The minimum edge distance used in the maximum damage models was 0.9D.

Maximum fastener sizes used in a repair before plugging. Larger fasteners in the outstanding flange attract load away from the skin flange of the longeron. Thus large fasteners (particularly in positions 2 thru 5 as shown in Figure 6) cause higher bypass loads at the lead fastener. The largest fasteners installed without freeze plugging the hole were 2nd oversize 5/16 or 0.344 diameter. (Baseline fasteners were .) Any holes requiring cleaning out greater than this were freeze plugged and had the nominal fastener installed through the plug. This is considered less severe than the large fastener condition provided edge distances are not less than 0.9D (based on plug diameter.) This is because with a plug and nominal fasteners, there is less load attraction and the bearing stresses are lower due to the larger plug.

As later aircraft were inspected, provided their damage fell within the maximum damage bounds, no further work was necessary.

Figure 6: Maximum Damage Model "Sad Case 2". Model Sad Case 1 has the same geometry except fastener hole #1 is 0.250" at an edge distance of 0.225

6 FEM Modelling of Joint to Obtain Bearing and Bypass Loads A significant amount of effort was expended to obtain an accurate FEM model of the tension fitting joint. This was required to obtain a good understanding of the bearing bypass loads at different fastener locations and thus find the critical damage conditions across the fleet. The main features of the model are as follows.

The models used plate elements to represent both the fitting and longeron;

Only a symmetric half model was used taking advantage of the symmetry of the joint; Fastener flexibility was based on standard flexibility equations;



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Contact constraints between both the longeron flanges and the fitting flanges were modelled using the NASTRAN Implicit Non Linear solver ;

Contact constraints with the back to back fitting were also modelled as a rigid symmetric constraint this was important as it effectively shifts the offset moment produced by the tension bolt.

Pre load for the tension bolt was modelled but not significant. The model was given length equal to the local frame spacing (10) at which simple supports were assumed.

A variety of pictures taken from the model are shown in Figure 7 thru Figure 9.

After the eyeball assessment, five individual damage conditions (together with the nominal condition and two Sad Case maximum damage scenarios were analysed using the FEM model to obtain fastener loads. Although fastener positions varied somewhat between aircraft, no positional variation was used in the models. This made it easy to update models for different fastener configurations: only the fastener properties needed to be varied * .

Figure 7: Picture of basic joint model using PATRAN "Display Shell Thickness" utility function. (Note: The modelling was done with plate elements NOT solid elements the thickness has been shown here for display purposes only to confirm correct separation of element mid planes and to relate actual solid parts to the 2Dplate representation shown in the following figures)

* Although fastener positions were not varied in the FEM loads models, actual edge distances were used in determining stress concentration factors for Severity Factor calculations and in performing DT Analyses.



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Figure 8: Basic Longeron and ( Model) Fitting Meshes (separated). Fastener elements joining longeron and fitting highlighted

Figure 9: Von Mises Stresses (maximum of both sides of plates) on a x10 deformation plot basic constraints indicated loading is in the xdirection via a constant stress along the LH edge of the longeron equivalent to load of 11628 lbf. This is half

the joint load. Fitting is restrained at the right hand end by the tension bolt.



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The effects of different modelling options on fastener loads for the nominal fastener configuration (fasteners 1 thru 5 are 0.250 HiLoks) is shown in Table 1. Besides showing the effects of the different contact constraints this table also shows the effects of some minor corrections made during the development of the model. As can be seen, the fastener loads change significantly with the contact constraints. This causes a significant change in the bearing and bypass loads at fastener 1. Comparing Model 6 with Model 1, the bearing load at the lead fastener is 8% lower in model 6 but the bypass load is 39% higher. It was also noted that significant differences existed between the bolt loads from the FEM model and loads given in the OEM stress notes 4, which used a simple bolt group analysis. Whilst this was expected, (given that the stress notes are primarily concerned with ultimate strength) it shows that the stress note values could not have been used as the basis for the DTA.

Model: 1 2 3 4 6

No Yes Yes Yes YesNo No Yes Yes YesNo No No Yes YesNo No No No YesNo No No No Yes

1 1607 1652 1424 1421 14872 1170 1286 1144 1141 12063 829 1056 990 990 10384 592 936 936 937 9765 452 923 989 992 1021

(6,7) 994 981 1057 1058 1016(8,9) 876 849 902 903 874

(10,11) 831 776 812 812 792(12,13) 841 736 769 769 752(14,15) 921 742 781 781 758(16,17) 1107 812 867 867 832(18,19) 1410 965 1044 1044 990

11630 11714 11715 11715 11742

Outstanding Flange Fastener FxLoads

Skin Flange Fastener Fx Loads

Longeron / Fitting ContactFitting / End-Pad ContactTension Bolt Pre-Load (18710 lbf)Flange Height CorrectedUnfastened Flange Present

Table 1: Comparison between different FEM models: the reason that the summed loads for models 2, 3, 4 and 6 dont sum to ~11628 lbf is because there are minor xdirection contact loads between the fitting and longeron that sum to between 86 lbf

and 112 lbf.

Results from sorting through the 5 worst cases of aircraft damage are shown in Table 2. These results are derived from running the FEM models with the fastener properties set to match the particular damage scenario. (As mentioned, nominal fastener positions were always used). The bearing and bypass stresses have been calculated from the fastener loads obtained in the model and the peak stresses have been obtained by applying appropriate stress concentration factors. See Niu1 for further details. As can be seen the actual damage configurations only exhibit peak stresses ~6% greater than the nominal condition, while the maximum damage cases (Sad Cases 1 and 2) show up to 14% peak stress increments.



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Nominal

FlangeBypassStress

BearingStress

Bearing /Bypass

Ratio

Peak

Stress

Nominal D e e/D bypass

(psi)

bearing

(psi)

BB Ratio

max (psi)

Hole #1 0.250 0.370 1.480 7860 15141 1.926 54503Hole #2 0.250 0.370 1.480 5621 12267 2.182 41676Hole #3 0.250 0.370 1.480 3696 10548 2.854 32057Hole #4 0.250 0.370 1.480 1889 9906 5.245 24845Hole #5 0.250 0.370 1.480 0 10351 INF 19400

A/C #1817E L/H D e e/D bypass

(psi)

bearing

(psi)

BB Ratio

max (psi)


A/C #2817E R/H D e e/D bypass

(psi)

bearing

(psi)

BB Ratio

max (psi)

Hole #1 0.328 0.518 1.579 8413 14453 1.718 54513Hole #2 0.328 0.532 1.621 5722 11236 1.963 39586Hole #3 (Frz-Plug) 0.410 0.518 1.263 4207 5065 1.204 24548

Hole #4 0.328 0.411 1.253 2105 8775 4.168 23599Hole #5 0.313 0.425 1.358 0 9214 INF 17012

A97-Sad CaseSad Case 1 D e e/D bypass

(psi)

bearing

(psi)

BB Ratio

max (psi)


Sad Case 2 D e e/D bypass

(psi)

bearing

(psi)

BB Ratio

max (psi)


Table 2: LimitLoad fastener bearing and bypass stresses from FEM analysis showing the two worst aircraft configurations (Labelled A/C #1 and #2) and "Sad Case" configurations. The maximum stress shown is the peak hole stress based on

applying appropriate bypass and bearing stress concentration factors 1,5 to holes.



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7 Damage Tolerance Analysis (DTA) Based on the results from the FEM analysis of fastener loads, a series of DT analyses were conducted to determine crack growth lives in the repaired structure. Per MILA83444, 0.050 x 0.050 circular quadrant primary cracks were assumed, with 0.005 x 0.005 secondary ones. For all cases shown in Table 2, the primary crack was placed towards the ligament. Two additional cases were also run, using the Sad Case 1 and 2 loads but with the positions of the primary and secondary cracks reversed. Such cases can be more critical in low edge distance holes as the large primary crack is not wasted in growing in a thin ligament . These cases were labelled as Sad Cases 1 (alt) and 2 (alt). The DTA tool used was AFGROW v53. Although this allows for primary and secondary cracks to be grown simultaneously, the solution databases in AFGROW do not allow appropriate coverage for 0.005 continuing damage cracks in the present structure. Hence classic AFGROW models were used as outlined in Figure 10 to cover the primary crack growth and continuing damage up to ligament fracture. The new AFGROW v5 crack with aslot models were then used for the final phases (III and IV) of growth until fracture. Although the presence of the skin flange of the longeron may be expected to retard crack growth during phase IV, this was conservatively ignored. This solution technique is summarized in Figure 10. The crack growth solution phases are shown in Figure 11 .

Figure 10: AFGROW solution Technique (see Figure 11 for labeling of crack growth phases)

Wasted in this sense is meant in the context of producing the critical or shortest crack growth life. If the larger crack is placed opposite the ligament then effectively while it grows the ligament will crack through anyway. If the

situation is reversed, the continuing damage crack opposite the ligament takes an extremely long time to grow as it is growing from a relatively shallow slot and does not increase in size much during the time the primary crack breaks through the ligament.



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Figure 11: CrackGrowth Scenario for Longeron. Note that the crack growth phase labelling here follows that of the OEM. Phase I is the ligament corner crack growing until it transitions to a through crack or causes ligament fracture; Phase II (not

shown) is the ligament through crack growing until ligament fracture; Phase III is the corner crack opposite the ligament growing until it transitions to a through crack; and Phase IV is the non ligament through crack growing until failure. Most

models in the present analysis dont have a Phase II segment because the ligament corner crack, causes ligament fracture

before it transitions to a through crack.

The cases shown were augmented by two extra Sad Case conditions in which the same loads were used but the edge distances were increased to maximum (0.57) rather than minimum values. This was done because large edge distances cause the slot formed at the end of ligament break through to be larger, which in turn significantly reduces the life during the latter stages of crack growth. Given that the proposed inspection regime for the repairs is to perform an Eddy Current Surface Scan (ECSS) along the exposed edge of the longeron flange, such conditions will give the lowest inspectable life as it is not possible to find a crack prior to ligament break through (without mandating Bolt Hole Eddy Current [BHEC] inspections). These cases were labelled Sad Cases 3 and 4.

Having determined the load share in the joint and knowing the geometry and material of the longeron the key

unknown required to run a DTA is the load spectrum. Unfortunately Australian Aerospace and the RAAF do not have access to OEM spectrum data used in the C130J 30 damage tolerance analysis 2. Furthermore that analysis does not give crack growth curves at the present damage location. Rather the aft plug longeron fitting locations are currently visually inspected every three years based on a DTA of the BL20 fitting at FS737 (the control point for this region). One solution to the problem of setting an inspection interval on the repairs would have been to perform a relative analysis using the existing inspection interval as a baseline. By assuming that the aft plug upper BL61 longerons are equally critical as the control point and running a DTA with a nominal spectrum for both the damaged and undamaged configurations a crack growth life knockdown factor could have been developed. This would have resulted in a very burdensome inspection regime because the extreme damage compared to the nominal configuration would have led to a significant knockdown factor. The reason this solution technique is of limited use

The RAAF DTA2 gives material da/dN curves for the 7075 T6 extrusion material used for the longeron. This OEM data was used in preference to other data sources.



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Figure 12: DTA Results

9 Cost A significant amount of engineering effort was expended on the development of the inspection requirements for the aft plug longerons. This was mainly done by AA Engineering but also involved review by the OEM. The most significant instances of damage were confined to 6 aircraft, ignoring the original aircraft on which the longerons were changed. An approximate labour costing of the present Safety ByInspection solution versus a full replacement solution is as follows, (based on ball park costings only).

Safety ByInspection Cost includes:

Development of Engineering Safety By Inspection intervals; Development of static analysis for repairs on 6 aircraft; Freeze plug and O/S repairs as well as fitting replacement on 6 aircraft Extra Inspections of all 12 aircraft over the next 21000 Hrs at 3000 hour intervals

Total = 2300 Hrs

Alternative cost of replacing 6 sets of discrepant longerons and fittings (assuming some learning as we go on):

Replacement of both Longerons on 6 aircraft: ~2*6*500 = 6000 Hours

Thus the nett saving is 3700 Hrs (~62% hour saving compared with replacement). The true saving is greater as the above figures take no account of aircraft down time and the impacts this would have on RAAF capability during an extensive longeron replacement effort. To obtain this true cost, consideration would have to be given to the costs of aircraft down time as well as the direct labour costs involved.



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10 Conclusion The use of DTA tools and indepth analysis has allowed Australian Aerospace and the RAAF to see a significant cost saving through the implementation of a Safety byInspection program on repairs to aft plug upper BL61 longerons on C130J 30 aircraft. Due to the extensive nature of damage found across many aircraft in the fleet, the only other option was full longeron replacement. It could be argued that in the present case a DT analysis was not required as the locations of damage were not existing DTA locations and the stresses were relatively low. As against this, the damage conditions were severe and well beyond what would be normally accepted in repair designs. Given the extent of the damage, it is unlikely a convincing case could have been made to accept the designs as is. Although the DT analysis performed was extremely conservative, it has allowed us to demonstrate that it is acceptable to repair the longerons in situ with a combination of oversize fasteners and freeze plugs. This has resulted in a considerable time saving over repairing the 6 most damaged aircraft via longeron replacement.

11 References 1. Niu, M.C. Y,, Airframe Structural Design, Conmilit Press, 1988.

2. LM Report A61A08A626, RAAF C130J 30 Damage Tolerance Assessment, May 2005 3. AFGROW v5.01.05.16, , LerTech Inc., 26/July/2010 4. LG95ER0038, C130J and C130J 30 Center Fuselage Stress Analysis, May 2000 5. Petersons Stress Concentration Factors, 2nd Edition, W.D. Pilkey, John Wiley, 1997

Technical Paper - Ensuring Airworthiness for C-130J-30 Aft-Plug Upper BL61 Longerons

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