WON-SHIK CHU AND SUNG-HOON AHN*School of Mechanical and Aerospace Engineering
Building 301, Room 1205, Seoul National University
San 56-1, Shinlim-dong Kwanak-ku, Seoul, Korea, 151-742
(Received April 9, 2004)(Revised July 10, 2004)(Accepted July 18, 2004)
ABSTRACT: As composite materials are gaining wide acceptance in aircraftstructures, the repair of damaged composites is becoming an important issue.Composite repair, however, has been performed more or less in a manual fashion,an integrated repair system is desirable.In this paper, an Internet-based advisory service (Repair Advisory Service, RAS)
for composite repair is proposed to increase the efficiency of the repair process.Based on the client–server architecture, RAS offers time-effective sharing of engi-neering knowledge between remote locations, easy maintenance and update of theapplication software, and ubiquitous access to the services. The web browser is usedas the user interface, which provides an easy access to the service and softwarefriendly environment. The RAS system provides integrated design and repair toolsfor repair technicians and engineers as follows:
(1) A design software for scarf repair and lap repair estimates the tensile failureand shear failure of repaired structures.
(2) A structural repair manual is provided in the PDF document for viewing, anda web search engine can search the document.
(3) A fabrication module of RAS generates a CNC toolpath to cut repair patches.(4) A CNC-based prepreg cutting machine has been developed to implement an
automated composite repair.
A scenario for the web-based remote repair was proposed, and a test repair wasperformed on an aircraft wing. The RAS service is open to public and available athttp://fab.snu.ac.kr/ras.
KEY WORDS: composite, repair, patch, internet, world wide web (WWW),automation, NC code.
INTRODUCTION
AS COMPOSITE MATERIALS are gaining wide acceptance in aircraft structures, the repairof damaged composites is becoming an important issue. The issues in composite
*Author to whom correspondence should be addressed. E-mail: [email protected]
Journal of COMPOSITE MATERIALS, Vol. 39, No. 9/2005 827
repair include high repair cost, material interchangeability, low temperature-curablerepair prepreg, water ingression, and structural integrity. To address these issues,an international committee called Commercial Aircraft Composite Repair Committee(CACRC) was formed with the objective to develop standards on composite repair [1].Although a number of researchers have studied on repair techniques [2–10], mechanicalproperties [11,12], and mathematical modeling [13–15], a large portion of the earlierwork was focused on manual repair that inherently requires labor-intensive fabricationprocesses. The following subsection describes a typical repair process of a damagedcomposite.
Repair Process
The repair process starts from the identification of the damage. The type of the originalcomposite and the size of the damaged area are identified. Field technicians either proceedwith the standard repair procedure or consult the repair engineers for further supervision(this procedure may vary depending on the airlines).
Figure 1 schematically shows the scarf repair and the lap repair techniques, whichare typical methods of repairing composites. Usually a small damage, less than 4 in. (about10 cm) in the maximum damaged area, is repaired according to the Structural RepairManual (SRM) written by the Original Equipment Manufacturers (OEM) of the aircrafts[16,17]. For the larger damage, engineers’ decision on the repair process may be required.If the damage is too severe and large to handle by the airline’s repair shop, the damagedcomponent ought to be shipped to the OEM for an advanced repair.
Although repair materials and techniques have been advanced, field repair still involvesmanual processes such as SRM look-up, fabric cutting, resin impregnation for handlay-up, and heat-blanket curing. Thus, the quality of the repaired part depends somewhatupon the craftsmanship of the individual repair technician.
In this paper, an Internet-based advisory service for composite repair is proposedto assist repair engineers and technicians with networked design and fabrication tools.The traditional repair may be expedited by applying the automated process and sharingof repair information (Figure 2).
Figure 1. Schematic geometries of scarf repair and lap repair [12].
Based on the client–server architecture, RAS takes advantage of distributed computingsuch as the time-effective sharing of engineering knowledge between remote locations, easymaintenance and update of the application software, and ubiquitous access to the servicesprovided.
Communication of RAS
In order to share information effectively throughout the Internet, the three-tier client–server architecture was applied to the RAS system [18] (Figure 3). Having web browsersas its user interface (1st tier), accessibility to the service is maximized. Thus, engineers andtechnicians in different repair shops in different cities can access a common RAS server(2nd tier) via the web browser to obtain information that is located in several modules (3rdtier). For numerical calculations, the MATLAB web server was applied [19]. A test versionof RAS uses the system shown in Table 1, but its communication system can be modifiedand customized for the airliner’s network environment. For example, if ‘intranet’ isavailable, important business information may be transferred between different businessunits in a more secure manner.
To integrate the network system, Active Server Page (ASP) [20] and Common GateInterface (CGI) [21] were used as the middleware located between the client and thecomponents at the server.
Functionality of RAS
Several modules were implemented in RAS to assist composite repair, and thefunctionality of each module is as follows:
� PDF-based SRMIn RAS, the OEM’s SRM is provided in the PDF (portable document format) [22]
for easy viewing and searching by the web browsers instead of looking up paper versions
Figure 2. An Internet-based communication for repair.
of the manuals (Figure 4). The search for SRM was implemented by the Googlesearch engine [23]. If a faster and easier search for specific information is desired, theformat of SRM may be further revised to enable more graphical navigation. For testingpurposes, the SRM of wet lay-up repair for Boeing’s 777 [16] was temporarily linked to thewebsite.
� Material DataThe properties of the base material and the repair material are stored as a material
database. Current RAS contains test data of Hexcel F593 and Dexter Hysol EA9396, bothreinforced with 3k70 plain weave fabric. To predict failures at the hot/wet condition aswell as at the room temperature/dry, material properties measured at hot/wet conditionwere also stored in the database. The hot/wet condition used in the data is 82�C/100%humidity. The material database was established using Microsoft Access with the webconnection [24].
Figure 3. Communication architecture of the RAS system.
Table 1. Specification of the web-based system.
Component Specification
Computer Pentium4, 2GHzOS MS Windows 2003 ServerMiddleware ASP/CGINumerical calculation MATLAB 6.5Database MS Access 2000
� Repair designFor the case where strength design of repair is necessary, failure prediction modules
were implemented in the RAS. The failure loads and modes of the lap repair and the scarfrepair are calculated by MATLAB codes [12]. Figure 5 shows the web-based designinterfaces for the scarf repair and lap repair. Each web page reads the geometric param-eters of a repair type, stacking sequences of base and repair laminates, and environmentalconditions of repaired composites such as room temperature/dry or hot/wet.
Once the parameters are transmitted to the RAS server via the Internet, numericalcomputations for the given repaired structure are performed. The results are shown to theuser as another web page (Figure 6). Based on the failure results, engineers can evaluatethe load carrying capacity of various repair designs. The details of the numerical modelsare discussed in ‘‘Strength Design’’.
� Automated FabricationUsing the RAS, the repair procedure for cutting prepreg into the designed dimensions
may be automated. An automated prepreg cutting machine was developed and tested toprovide higher efficiency than the traditional manual process. More description is given in‘‘Automated Fabrication’’.
� InventorySince most resins for wet lay-up repair consist of two-parts epoxy that has a limited shelf
life (6–12 months), the materials will expire if not used within that period [25]. In addition,various kinds of repair substances such as honeycomb, vacuum bagging material, andprepreg are used for composite repair. The web-based inventory enables engineers tokeep track of the amount of usable repair material, and share the information with the
Figure 4. An example of web-based SRM for a scarf repair (Boeing 777 typical sanding instructions [16]).
purchasing department for stable supply. The current RAS just shows the idea of theweb-based inventory, not the functional one.
STRENGTH DESIGN
The design of repair geometry can be obtained by strength models or models basedon damage tolerance. In the RAS, strength models are applied to predict the desired
Figure 5. Web-based user interfaces of RAS. (a) Scarf repair; (b) Lap repair.
strength. The laminate and the repair patch consist of fiber-reinforced plies whichbehave in linearly elastic manners. The failure models are based on the maximum straincriterion for tensile failure while the shear failure model is applied for interfacial failure[13]. The shear failure model was modified from the Hart-Smith models [14,15] toaccommodate the anisotropic property of the individual composite layer. For the wetlay-up repair, there is a thin resin layer between the laminate and the repair patch, andwe treat the resin layer as an ‘interlayer’ which exhibits elastic–perfect plastic shear
Figure 6. Web-based failure prediction from RAS. (a) Scarf repair; (b) Lap repair.
behavior. The shear strains at the elastic limit and at plastic failure are �ef and �pf ,respectively (Figure 7).
Although the Ahn model [13] may fail for highly anisotropic materials, the modelrepresents reasonably well the typical repair parts made of woven fabric laminates, whichhave less anisotropic properties.
Model of Uniform Lap Repair
The model for the double-sided uniform lap repair is illustrated in Figure 8. The lay-upof the laminate/repair patch assembly is symmetric. In order to model such a symmetricrepair, it is sufficient to consider only one-half of the repaired laminate.
The repaired geometry subjected to a tensile load may fail either by tensile failureof the base laminate, by tensile failure of the repair patch, or by shear failure of theinterlayer.
TENSILE FAILURE OF THE LAMINATE OR REPAIR PATCHFor a symmetric laminate and repair patch, the in-plane strains at failure are
Laminate Repair patch
ð"Lx ÞF ¼ aL11F ð"Rx ÞF ¼ aL11F
ð"Ly ÞF ¼ aL21F ð"Ly ÞF ¼ aL21F
ð"Ls ÞF ¼ aL61F ð"Ls ÞF ¼ aL61F
ð1Þ
where "x, "y, and "s are the off-axis in-plane strains. The aijs are the components of thecompliance matrix. The superscripts L and R refer to the laminate and the repair patches,respectively.
Figure 7. Illustration of the shear stress–shear strain relation of an elastic–perfect plastic interlayer.
p, q, r, and s are constants which must be determined from the boundary and continuityconditions.
BOUNDARY AND CONTINUITY CONDITIONS
– At the ‘free’ ends of the laminate and the repair patches, the axial load is zero.
NL ¼ 0 at x ¼ 0
NR ¼ 0 at x ¼ dl
ð8Þ
– At those ends of the laminate and the repair patches where loads are applied, the load(per unit width) NL in the laminate is equal to the applied load (per unit width) P, andthe load (per unit width) NR in each repair patch is equal to P/2.
NR ¼P
2at x ¼ 0
NL ¼ P at x ¼ dl
ð9Þ
d�
dx¼ �
PaR112hi
at x ¼ 0 ð10Þ
d�
dx¼
PaL11hi
at x ¼ dl ð11Þ
– At the locations where the elastic and plastic regions meet (x¼ xp1 and x¼ xp2) theshear strains in the elastic and plastic regions are equal (with the value �ef) and arecontinuous. Correspondingly, we have
�e ¼ �p ¼ ð�ef Þ at x ¼ xp1,
d�edx
¼d�pdx
x ¼ xp2
ð12Þ
The locations xp1 and xp2 are unknown.
The system of equations with boundary conditions and continuity condition was solvednumerically to provide tensile failure loads of the laminate and repair patches as well asthe interlayer failure as a function of the lap length (Figure 6).
Model of Scarf Repair
A symmetric composite laminate repaired by the scarf technique was modeled (Figure 9).When such a repaired laminate is subjected to a tensile load, the base laminate may fail in
tension, the repair patch may fail in tension, the interlayer may fail in shear, the laminatemay fail in tension while the interlayer fails in shear, or the repair patch may fail in tensionwhile the interlayer fails in shear. The results fed back to the user include the failure loadand the mode (Figure 6).
INTERLAYER SHEAR STRAINThe shear strain is calculated as follows.
dðkNLÞ
dx� k� ¼ 0 Laminate
dðkNRÞ
dxþ k� ¼ 0 Repair patch
ð13Þ
kNL (¼ kNLx ) and
kNR (¼ kNRx ) are the in-plane loads (per unit width) inside the laminate
and inside the repair patch, respectively. k�(¼ k�zx) is the shear stress in the interlayer. Thesuperscript k refers to the kth overlap segment.
In the elastic shear region, the shear strain and shear stress are
k� ¼ k�e and k� ¼ Gk�e ð14Þ
k�e ¼kp sinhð�xÞ þ kq coshð�xÞ ðElastic regionÞ ð15Þ
where � is defined as
k�2� �
¼G
hi2k�L
11 þk�R
11
� �ð16Þ
In the plastic region, the shear strain and the shear stress are
kp, kq, kr, and ks are constants, which must be determined from the boundary andcontinuity conditions.
BOUNDARY AND CONTINUITY CONDITIONS
– At the inside edge of the repair patch (x¼ 0) the axial load (per unit width) is zero in thelaminate, and is equal to the applied load (per unit width) P in the repair patch. At theoutside edge of the repair (x¼ xK) the axial load (per unit width) is zero in the repairpatch, and is equal to the applied load (per unit width) P in the laminate
1NL ¼ 01NR ¼ P
at x ¼ 0 ð19Þ
KNL ¼ PKNR ¼ 0
at x ¼ xk ð20Þ
dð1�Þ
dx¼ �
p1�R11
hiat x ¼ 0 ð21Þ
dðK�Þ
dx¼
pK�R11
hiat x ¼ xk ð22Þ
– At the edge of each overlap segment (at x¼ xk) the shear strain in the interlayer iscontinuous and the in-plane loads are equal and opposite on the left and right sidesof the segment. It is assumed that the load is transmitted only by continuous layers(Figure 10).
Figure 10. The equations, boundary and continuity conditions for calculating the shear strain in the interlayerwhen the interlayer is perfect plastic near the x¼0 and x¼ xK ends [13].
Accordingly, at the edge of each segment, the following continuity conditions areapplied:
kþ1� ¼ k � ð23Þ
kþ1NL ffi kNL at x ¼ xk ð24Þ
kþ1NR ffi kNR ð25Þ
The above continuity conditions for the kth segment can be expressed in terms of k� asfollows:
kþ1� ¼ k� ð26Þ
dðkþ1�Þ
dx¼
dðk�Þ
dxk�� þ kP� at x ¼ xk ð27Þ
where k�� and kP� are defined as
k�� ¼kþ1�L
11 þkþ1�R
11k�L
11 þk�R
11
ð28Þ
kP� ¼ �P
hi
k�L11
kþ1�R11 �
k�R11
kþ1�L11
k�L11 þ
k�R11
� �ð29Þ
– At the locations where the elastic and plastic regions meet (x¼ xp1 and x¼ xp2) the shearstrains in the elastic and plastic regions are equal (with the value �ef ) and arecontinuous. Correspondingly, we have
k� ¼ k�p ¼ ð�ef Þ at x ¼ xp1,
dðk�eÞ
dx¼
dðk�pÞ
dxx ¼ xp2
ð30Þ
The locations xp1 and xp2 are unknown, a priori, and must be determined from thesolutions of the equations summarized.
FAILURE LOADTo determine the applied tensile load under which either the laminate or the repair
patch fails, the in-plane loads in the laminate and the repair patch are calculated asfunctions of axial position x
The off-axis in-plane strains, as a function of x in the laminate and in the repair patch are:
Laminate Repair patchk"Lx ðxÞ ¼
k�L11
kNLðxÞ k"Rx ðxÞ ¼k�R
11kNRðxÞ
k"Ly ðxÞ ¼k�L
21kNLðxÞ k"Ry ðxÞ ¼
k�R21
kNRðxÞ
k"Ls ðxÞ ¼k�L
61kNLðxÞ k"Rs ðxÞ ¼
k�R61
kNRðxÞ
ð32Þ
– If failure occurs at x¼ 0 end of the repaired laminate, the failure is considered to haveoccurred in the repair patch.
– If failure occurs at x¼ xK end of the repaired laminate, the failure is considered to haveoccurred in the laminate.
– If failure occurs in the laminate or the repair patch at one of the overlap segments, thenfailure is considered to have occurred due to a combination of laminate (or repair patch)failure and interlayer failure.
AUTOMATED FABRICATION
Software
To achieve more automated composite repair, a fabrication process was developed. Asshown in Figure 11, the RAS interface reads from the user dimensions related to the scarf
Figure 11. User interface for the automated patch fabrication.
repair. Currently, concentric circular cutting pattern is available in the RAS. The user canspecify the diameters of repair patches and the number of repair plies. The RAS generatesG&M codes promptly at the server so that the engineer can download the code from theweb page.
Figure 12 shows an example of a G&M code generated by the fabrication moduleof RAS.
Hardware
Among various prepreg cutting processes, three-axis Computer Numerical Control(CNC) was used in RAS. Thus the G&M codes are the standard machine control languageused in RAS.
A list of G&M codes generated by the web-based system is uploaded to the CNC cuttingmachine (Figure 13). The prepreg cutting machine was developed by modifying the cuttingknife from the original micro endmill [26]. Figure 14 shows a repaired aircraft wing usingthe RAS system. The repair prepreg is automatically tailored into a series of circularpatches (Figure 15c). The tailored prepreg patches were laid on the damaged area by arepair engineer. This automated cutting machine assists repair engineers who layout theprepreg patches into the damaged area.
SCENARIO OF PROPOSED REPAIR PROCESS
After damage inspection is done (Figure 15a), a repair engineer accesses the RASwebsite from his/her computer (Figure 15b). Using the RAS, the repair engineer cansearch and view the relevant documents for the specific part to be repaired. If necessary,the load carrying capacity of the repaired composite can be estimated in a few secondsusing the failure prediction modules (Figure 15c). Once the engineer makes a decision onthe types of repair, he/she goes forward to remove the damaged area by routers and othermechanical tools (Figure 15d). While machining is in progress, the repair prepreg can be
Figure 12. An example of G&M code generated by RAS.
cut into the patch geometry by an automated cutting machine. The engineer can applythe repair patches and vacuum bagging layers to the cleansed surface of the removed area.While heating up the heat blanket for curing, the engineer can start another repair processqueued on his/her work list.
Figure 14. An example of the repaired wing using the RAS system. (a) Damaged part of aircraft wing;(b) Prepared surface by machining, sanding and cleaning; (c) Automated prepreg cutting; (d) Repair ofdamaged part.
Using the RAS, the process of composite repair can be expedited by the web-basedsimulation and by the automation of the prepreg cutting process. The overall result will bereduced time and cost of composite repair (Figure 15e).
DISCUSSIONS
RAS is a newly proposed tools and repair process. Considering the practicalapplications of RAS in commercial airliners, the following issues are worthy ofdiscussion:
. Prepreg vs. wet lay-up: Although prepreg for repair purpose, such as Hexcel’ M20, hasbeen developed, wet lay-up is still a popular repair process. Prepreg is ready to be cut bythe ply cutter while the wet lay-up requires impregnation process before cutting. Sincethe impregnated fabric is more difficult to cut than the prepreg, the application ofautomated cutting could be not applicable.
. Cost of the cutting machine: The CNC-based ply cutter was developed at a low price.For small repair stations or for small-size airliners, even the cost of the cutting machinemay be not trivial. The economic benefit from automated repair using the cutter is to beconsidered by comparing it with the manual repair. The number of repairs completedin the repair station per certain time frame and the time saved by having the automatedcutting system are the main parameters to be considered.
. Authorization: Some repair stations do not have certified engineers to modifystandard repair geometry, but merely follow OEM’s SRM. In such a case, design-ing or modifying a repair geometry that is different from SRM cannot getauthorized.
Figure 15. Expedited repair process using the RAS system.
Emphasizing the bright side of the web-based information exchange, RAS has thefollowing advantages over the current practice of composite repair:
. Searching information via web is gaining more ground than looking up thick paperversion documents. Web-based SRM and other in-house documents can increaseefficiency of the repair process.
. Knowledge acquired from engineers’ experience is a valuable asset for any businessorganization, especially for a manufacturing company [27]. The web is such an effectivetool to store and share the knowledge base of manufacturing processes, here repairingof the composite. Memos, tips, industrial news, and process knowledge stored in RASare possibly reformulated into a knowledge map that can be useful in the trainingof newly hired engineers as well as to provide experienced engineers with checklists ofstandard repair process.
CONCLUSIONS
The web-based RAS has been developed to provide software tools that might be usefulin composite repair. The RAS provides process knowledge as well as promotion of anautomated repair process. The repair engineers may utilize RAS for fast repair design,and for prepreg cutting system resulting in saved time for composite repair. Since itis developed to prove the concept, RAS is open to public and available at http://fab.snu.ac.kr/ras/. A couple of components in RAS are under development, and morematerial properties will be included in the database. Broad feedback from field techniciansand engineers as well as researchers working on composite repair may expand andimprove the functionality of RAS.
ACKNOWLEDGMENTS
This research was supported by the Korea Research Foundation (KRF 2001-003-E00075) and Brain Korea 21. The authors thank Mr. Young-Soon Kim andMr. Seungyong Kim of Korea Aerospace Industry (KAI), Mr. Jungwon Han,Mr. Hyung-Ik Kim and Mr. Jung-Hoon Choi of Korean Air for their comments andsuggestions.
REFERENCES
1. Bardygula, L.O. (1996). CACRC: Progress and Plans, Composites ‘96 Manufacturing and ToolingConference, Society of Manufacturing Engineers, 229–239.
2. Myhre, S.S. (1985). Advanced Composite Repair-Recent Development and Some Problems,In: Brown, H. (ed.), Composite Repairs, SAMPE Monograph (1), pp. 14–25.
3. Stone, R.H. (1985). Development of Repair Procedures for Graphite/Epoxy Structureson Commercial Transports, In: Brown, H. (ed.), Composite Repairs, SAMPE Monograph (1),pp. 26–39.
6. Lee, F., Brinkerhoff, S. and McKinney, S. (1991). Low Energy Cured Composite Repair System,In: Hamermesh, C.L. (ed.), Composite Repairs, SAMPE Monograph (2), pp. 24–35.
7. (1993). D-53900, Repair Procedures for Composite Structures, Boeing Commercial Aircraft Co.
8. Wegman, R.F. and Tullos, T.R. (1993). Adhesive Bonded Structural Repair, Part III Repairof Composite, Honeycomb Cored, and Solid Cored Structures, SAMPE Journal, 29(6): 8–12.
9. Shyprykevich, P.G., Springer, S. and Ahn, S.H. (1995). Standardization of CompositeRepairs for Commercial Transport Aircraft, Composite Repair of Aircraft StructuresSymposium, pp. 1–24.
10. Ahn, S.H., Springer, G.S. and Shyprykevich, P. (1996). Composite Repair with Wet Lay-upand Prepreg, Composites ‘96 Manufacturing and Tooling Conference, Society of ManufacturingEngineers, pp. 211–226.
11. Chesmar, E.F. (1996). Metal Bond and Composite Repairs: Similarities and Differences,Composites ‘96 Manufacturing and Tooling Conference, Society of Manufacturing Engineers,pp. 281–290.
12. Ahn, S.H. and Springer, G.S. (1998). Repair of Composite Laminates – I: Test Results, Journalof Composite Materials, 32: 1036–1074.
13. Ahn, S.H. and Springer, G.S. (1998). Repair of Composite Laminates – II: Models, Journalof Composite Materials, 32: 1076–1114.
20. Buser, D., Kauffman, J., Llibre, J.T., Francis, B., Sussman, D., Ullman, C. and Duckett, J.(2000). Beginning Active Server Pages 3.0, Wiley Publishing, Inc., pp. 9–126.
21. Birznieks, G., Guelich, S. and Gundavaram, S. (2000). CGI Programming with perl, O’reilly &Associates, 2nd Edn., pp. 65–83.
22. Padova, T. and Rosenbaum, S. (2003). Adobe Acrobat 6 PDF bible, John Wiley & Sons, pp. 169–222.
23. http://www.google.com.
24. Prague, C.N. and Irwin, M.R. (1999). Microsoft Access 2000 Bible, John Wiley & Sons, pp. 101–292.
25. Armstrong, K.B. and Barrett, R.T. (1998). Care and Repair of Advanced Composites, Societyof Automotive Engineers, Inc., pp. 19–110.
26. Bae, K.M., Lee, S.U., Lee, B.C., Gang, G.S., Kim, H.U., Hong, Y.J., Jin, Y.S., Kim, J.C.,Park, J.H., Park, S.M., Park, S.H., Baek, C.I. and Ahn, S.H. (2003). PaperMill – A LayeredManufacturing System Using Lamination and Micro Endmill, Transactions of the Society ofCAD/CAM Engineers, 8(2): 153–157.
27. Ahn, S.H., Bharadwaj, B., Khalid, H., Lieu, S.Y. and Wright, P.K. (2002). Web-basedDesign and Manufacturing Systems for Automobile Components: Architectures and UsabilityStudies, International Journal of Computer Integrated Manufacturing, 15(6): 555–563.
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