NASA Contractor Report 4620 Composite Chronicles: A Study of the Lessons Learned in the Development, Production, and Service of Composite Structures Louis F. Vosteen Analytical Services & Materials, Inc. • Hampton, Virginia Richard N. Hadcock RNH Associates • Huntington, New York National Aeronautics and Space Administration Langley Research Center • Hampton, Virginia 23681-0001 Prepared for Langley Research Center under Contract NAS1-19317 November 1994 https://ntrs.nasa.gov/search.jsp?R=19950010444 2020-03-10T21:46:26+00:00Z
Transcript
NASA Contractor Report 4620
Composite Chronicles: A Study of the LessonsLearned in the Development, Production, andService of Composite Structures
Louis F. Vosteen
Analytical Services & Materials, Inc. • Hampton, Virginia
Richard N. Hadcock
RNH Associates • Huntington, New York
National Aeronautics and Space AdministrationLangley Research Center • Hampton, Virginia 23681-0001
Prepared for Langley Research Centerunder Contract NAS1-19317
Structural Design, Analysis, and TestDesign and Certification RequirementsStructural DesignStructural AnalysisStructural Test
Cost Considerations in Design
Design R&D NeedsLessons Learned
Materials & ProcessesMaterialsProcessesJoints and AttachmentsLessons Learned
22234
55
7111112
1314
1515161717
Page
Quality Control 23General 23
Nondestructive Inspection 23Effects of Defects 23
Lessons Learned 23
Supportability 24General 24
In-Service Damage & Repair 24Lessons Learned 25
Conclusions and Recommendations 25
Manufacturing and ToolingManufacturing
ToolingLessons Learned
Appendix 27
Military Aircraft 27Commercial Transport Aircraft 35General Aviation 40
Remotely Piloted Research Vehicles andDrones 43
Helicopter Applications 44Observations and Conclusions 48
18 References 51
18
20 Acknowledgments 5422
LIST OF FIGURES
FigureNumber Title
1. Organization of structures Design Build Team.2. Current FAA damage tolerance design requirements.3. Revised damage tolerance requirements.4. Example of process description from shop manual.
LIST OF TABLES
TableNumber Title
1. Civil aircraft in use by US military and the certificating country and agency.
2. Examples of moldform tooling used on some past and current programs.
Composite components on military aircraft.(a) 1960-1979(b) 1980-1995Composite components on commercial transport aircraft.Boeing composite component suppliers.Airbus composite component suppliers.Composite components on business aircraft.Composite components on private, trainer, competition, RPV, and drone aircraft.Composite components on Helicopters.
iii
m-1.
m-2.
A-3.A-4.
A_5.A-6.A-7.
Page388
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COMPOSITE CHRONICLES: A STUDY OF THE LESSONS LEARNED
IN THE DEVELOPMENT, PRODUCTION, AND SERVICEOF COMPOSITE STRUCTURES
Louis F. Vosteen
Analytical Services & Materials, Inc.
and Richard N. HadcockRNH Associates
INTRODUCTION
The development of advanced fiber
composites in the 1960's brought aircraftdesigners a new material option comparable tothe introduction of aluminum some 40 yearsearlier. Carbon fibers, with moduli and
strengths comparable to steel and a densityhalf that of aluminum, created visions of 50%
weight saving for airframe structure. Althoughsuch weight savings have been achieved on afew specific components, the added weightassociated with load introduction, the need to
satisfy multiple design conditions, design andproducibility requirements that usually requirea balance of in-plane properties, accessibilityfor maintenance, inspection, and damagerepair, and production cost constraints havemade weight savings of 15-30% a more realis-tic and achievable goal.
As typically happens with the introduc-tion of new technologies, advanced fiber
composites have had their share of difficultiesalong with many notable successes. A fewprograms involving utilization of compositeshave experienced unforeseen problems andpremature failures during development testing.Many more have been very successful, haveprovided significant weight savings, andservice experience has been excellent. Thou-sands of safety-of-flight components are in
production and are providing excellent serviceon more than forty different US and foreignmilitary and civil fixed-wing aircraft andtwenty different helicopters. Most of thesecomposite components, as well as componentson technology development and demonstratoraircraft, are identified in the Appendix to this
report.
The major issues today are associatedwith the materials, manufacturing, and repaircosts and not with structural performance. Thestructural problems and failures that haveoccurred were primarily caused by deficien-cies in the detail design of joints, cut-outs, anddiscontinuities, designs that did not makeproper allowance for the lack of ductility andanisotropic mechanical properties of thecomposite materials. Other problems wereassociated with communications between
engineering and manufacturing personnel,
especially when the people and facilities werelocated hundreds of miles apart. In mostinstances, design changes corrected the struc-tural problems and most companies are nowusing the "Concurrent Engineering" approachwith collocated design-production teams toimprove communications.
Given the problems and failures thathave occurred, and are still occurring, one
must ask whether the problems experiencedare similar and inherent in the nature of ad-
vanced composites or whether new problemscontinually arise because of lack of technicalunderstanding of the materials and theirbehavior. Furthermore, when programs aresuccessful, is there an underlying reason thatshould be recognized, understood, and appliedin the future?
In an effort to find answers to these
questions, NASA Langley Research Centercontracted with Analytical Services andMaterials to conduct a study of past compositeaircraft structures programs and determine thelessons learned during the course of thoseprograms. The study was focused on findingmajor underlying principles and practices thatexperience showed could have a significant
The various organizational issues, needsand barriers associated with transfer of com-
posites technology were discussed at many ofthe meetings. There was general agreementabout the needs, but some differences of
opinion about barriers, some of which are
dependent on company organization and varyfrom company to company.
ORGANIZATIONAL BARRIERS
When advanced composite materials firstbecame available during the late 1960's, andwhen the Air Force, Navy and NASA began tofund the first advanced composites develop-ment programs, there was considerable lack ofsupport and skepticism of the projected ben-efits and usefulness of composites for aircraftstructures by industry.
The companies which were most suc-cessful had a few senior corporate executiveswho became champions for composites. Thesecompanies assigned some of their most ca-pable people to composites developmentprograms and set up teams to work with andrespond to the Government's initiatives. Eventhough significant advances were made by theearly 1970's, upper management in somecompanies were still raising concerns aboutrisks and questioned the predicted weight
2
saving potential of composites. Twenty years
later, there are still some senior corporateexecutives and engineers who feel that com-posites have no place in airframes. A senior
government representative suggested that topmanagement education in composites is morenecessary now than ever. Management shouldunderstand that problems invariably arise incomposites structures programs, but these
problems have generally been resolved satis-factorily.
ORGANIZATIONAL NEEDS
"Concurrent Engineering", whereby anew product or system is developed jointlyand concurrently by a team composed ofdesigners, stress analysts, materials and pro-cesses, manufacturing, quality control, andsupport engineers, as well as cost estimators,
has generally become the accepted approach toimprove the quality and performance andreduce the development and production costsof complex systems.
Most of the aerospace companies haveimplemented concurrent engineering ap-proaches in one form or another. Boeing, forexample, has formed "Design Build Teams"(DBTs) for development of the Boeing 777.The DBT hierarchy for a typical major struc-tural component is shown in Figure 1. TheBoeing structures teams are composed of tento twenty people from the various engineering,manufacturing, quality control, and costestimating departments. Where necessary,teams also include people from Boeing pro-curement and from subcontractors and suppli-ers. The teams, each of which are collocated,
are responsible for producing a final design,cost estimates, production planning, etc. of thestructural components-and subcomponents. Alldrawings and interfaces with other teams andsubcontractors are made using the Boeingcomputer network.
Lockheed, GD, Northrop, Vought andother companies have all implemented similarsystems for new programs. The Grumman"Task Teaming" approach also subdivides thedevelopment effort into tasks and then collo-cates small multidiscipline teams to performthe tasks. McDonnell Douglas versions are
Many of the people interviewed impliedthat almost all of the critical advanced com-
posites technology developed in the US hasbeen developed by industry in a research and
Program
Sytstem Integration Teams
• Led by management
• Not Collocated with other Teams
Other
Integration TeamsCollocated
Multidiscipine Team Members• Structures: Body, Wing,
& Propulsion
• Avionics, Flight Controls,
Subsystems, etc.
development (R&D) environment by a small
number of R&D people funded by DoD,NASA, or Independent Research & Develop-ment (IRAD) programs. Very little of thecritical technology has actually been devel-oped by universities, research centers, orgovernment laboratories. The fact that thetechnology was developed in a R&D environ-ment has caused most companies to havemajor problems with the transfer of the tech-nology and experience to people working onproduction programs. In four Navy and fourcommercial aircraft programs, major compos-ites design and producibility problems were
caused by the lack of composites experience ofthe people working on the programs.
One solution to problems in technologytransfer among R&D and production groupshas been to assign the experienced R&D
people to the production program. The produc-tion program gained from their expert knowl-edge and the less experienced people working
r
Empennage DBTs
Collocated
Multidiscipine Team Members
Horizontal & Vertical Stabilizer
Structural Design
Analysis
Manufacturing/ToolingMat's & Processes
QC
Suportability... etc.
• .J
Figure 1. Organization of a structures Design Build Team.
* Programs that were cited as having suffered from a lack of experienced composites personnel included Navy
programs for the V-22 wing and fuselage, A-6 wing, AV-8B wing, and A-12. Civil programs included Lear Fan,
Starship, Boeing 777-300 elevator, and L-1011 vertical satabilizer.
on the program benefited from the "expert's"experiences. However, other R&D and lessimportant production programs sufferedbecause the number of experienced composites
people was limited. In other companies, thissituation was ameliorated by assigning experi-enced engineers and manufacturing people towork with, and learn from, composite special-ists on R&D programs. Experienced metaldesign and manufacturing engineers can betrained in the critical elements of compositesdesign and manufacture in a relatively short
time. Some suggested that the training couldtake as little as a few weeks if the engineers
experienced in metals design are workingclosely with experienced composites engineerson composites programs. Many of these"composites production program experts"were later assigned to management or engi-neering positions on other production pro-grams or were transferred back into R&Dwhere their "real world" experiences from theproduction program added to their capabilities.
Additional problems have occurred whenthe design approach, structural analysis meth-ods, materials, processes, tooling, etc. devel-oped under R&D programs have been appliedto full-scale structures. R&D program budgets
have generally constrained the size and com-plexity of demonstration components, and insome cases, failed to identify real worldproblems. Many R&D designs were verystructurally efficient and demonstrated weight
savings, but were not producible withoutmajor design changes.
Very little of the technology developedby one company under an R&D contract wasever transferred to other companies via techni-cal reports, presentations, lectures and courses.Composites technology, which is highlyinteractive and must cross many disciplinaryboundaries, was best accomplished by havingexperienced and inexperienced people workingtogether. On programs that involved a team oftwo or more companies, technology wasreadily transferred between companies byhaving people from each company working
together, i.e., collocated.
The transfer of technology and know-how into and out-of some classified military
programs has been a major problem, particu-larly when one of the programs has specialrestrictions and security requirements. In one
program, for example, the components for anunclassified program were produced in arestricted area. The restrictions even precludedthe Government's program representativesfrom visiting the area to see first-hand howtheir components were being made.
LESSONS LEARNED
a.
b.
c.
d°
e.
f.
Composite structures technologyrequires far more interaction thanconventional aluminum structures
technology.
Structural design, certification, andtest requirements as well as materials,processes, manufacturing, tooling,quality control, product support, andcost issues must be addressed andunderstood from the start.
The production and operational costsof composite structures must becompetitive with counterpart metalstructures. Weight savings are abonus, but cost is the driver.
The "Concurrent Engineering" ap-proach, using small collocatedmultidiscipline teams, resolves prob-lems up front and is being used by themajority of US aircraft companies. Byresolving problems early, the costsand time spent on rework, modifica-tions, and changes are reduced.
Small companies and compositesR&D organizations of large compa-nies with engineering and productioncapability located in the same facilitysuccessfully practiced "ConcurrentEngineering" before the term wasinvented.
The "Concurrent Engineering" ap-proach enables designers to becomefamiliar with manufacturing and QCtechnologies, capabilities, and prob-lems, and vice versa. Designers
4
g.
h.
i.
j.
should spend time on the shop floor in
the production facility.
There is no substitute for compositesexperience. Experienced engineersand technicians are a very valuable
commodity and are a key towardassuring program success.
Composites technology is transferredby people working together and notby reports, presentations, and lectures.
Problems have been caused by short-
ages of composites-experiencedpeople. Many companies have senttheir key people to outside tutorialsand training courses and have imple-mented in-house composites trainingcourses for engineers, technicians andproduction personnel. Some of themost effective courses are taught in-house by experienced companydesign and manufacturing engineers.
Companies that have engineering andproduction facilities located in thesame area generally appear to havehad fewer problems and have re-solved problems more rapidly thancompanies with facilities that are farapart geographically.
STRUCTURAL DESIGN,
ANALYSIS, AND TEST
DESIGN AND CERTIFICATION
REQUIREMENTS
tion test programs. Some structural compo-nents designed to meet the requirements of one
agency had to be extensively redesigned tomeet the requirements of another. As a result,
design, certification, and production costsincreased enormously, and, because the com-
posite weight savings and cost projectionswere based on a different set of requirements,
the weight and performance targets for theaircraft were not achieved, adversely affecting
aircraft capability, competitiveness and price.Most of the government and industry peopleinterviewed felt that the basic requirementsshould be uniform and that some of the re-
quirements cause unnecessary weight and costpenalties to composites structures.
Twenty-two different US civil aircrafttypes, which were certificated under FARrequirements, are in service with the US AirForce, Navy, and Army. An additional tenforeign aircraft types, that were certificatedunder country-of-origin or JAA requirements,are in service with the US Air Force, Army,and Coast Guard. Five of the six JPATS
contender aircraft are derivatives of foreigntrainers, and one of these has a composites
airframe. The Slingsby T-3A trainer, whichwill soon enter Air Force service, is almost
entirely made from composites. A list of theseaircraft, their country of origin, and the certifi-cating agency is given in Table 1. Except forthe AV-8 and the trainers, all the aircraft are
used for transport or as electronic surveillanceplatforms. The AV-8A was essentially thesame as the British Aerospace (BAe) GR. Mk3 Harrier and was certificated by the British
military. The AV-8B, also designated the GR.Mk 5, was a major redesign of the AV-8A byMcDonnell Douglas and BAe and was certi-fied to US Navy standards.
Composite structural design and certifi-cation requirements were identified as a majorconcern at twenty-eight of the thirty-five
meetings with industry and government orga-nizations.
The differences among the design andcertification requirements for compositestructures specified by the FAA, the US AirForce, and the US Navy, have caused major
design problems and duplication of veryexpensive material and component certifica-
Certification to foreign standards is
causing some problems for the US services,which have to decide if the original certifica-
tion procedures and tests are satisfactory byUS military standards. The standards also haveto be reviewed by the maintenance organiza-tions, which have to inspect and repair damageon the basis of the original civil or foreigncertification requirements.
The current FAA damage tolerancerequirements were criticized as being overly
Table 1.- Civil aircraft in use by US military and the certificating country and agency.
severe by many of the people interviewed.
These requirements specify that a structurewith barely visible impact damage (BVID)
must be capable of sustaining design ultimate
flight loads in the most adverse temperature/humidity environment throughout the life of
the aircraft. (Federal Aviation Regulations
(FAR); FAA Advisory Circular, 1984) Pres-
surized fuselages with BVID must havesufficient residual strength to withstand the
combined effects of critical ultimate flight
6
loads in combination with normal operating
internal pressure and external aerodynamic
pressure. Residual strength must be established
by component or subcomponent tests, or by
analysis supported by test evidence. The
effects of temperature, humidity, and other
environmental factors that may result in
material property degradation must be ad-
dressed in the damage tolerance evaluation
(FAA Advisory Circular, 1984; Evaluation of
Composite Structure, Fed Reg, 1986).
The design maneuver limit load factorsfor Commercial Transport Category Airplanes
weighing more than 50,000 Ib are +2.5/-1.0g.(FAR Part 25.337) The current FAA require-ments call for BVID residual strength capabil-
ity of load factors of +3.75/- 1.5g with BVID.Airbus certification requirements call for onlylimit load (+2.5/-1.0g) residual strength capa-
bility.
has also barred the use of polyimide- andbismaleimide-matrix materials.
The three agencies also have differentrequirements for simulation of low energy
impact damage and different test procedures todetermine the fatigue strength and residual
strength of components or the full-scalefatigue test article.
Boeing Commercial Airplane Group(Dost, 1993) discussed a revised approach thatis similar to metal practices and USAF re-quirements for composites. The current FAA
requirements are shown in Figure 2. Anapproach proposed by Boeing is outlined onFigure 3. Boeing suggests reducing the designload requirement for non-visible and barely-visible impact damage to a level between limitand ultimate that would account for "Real
World" damage scenarios and limit loadconditions based on the maximum load perfleet lifetime. The current requirements for
easily visible damage would not be changed.
The US Air Force damage tolerance
requirements specify that the structure must becapable of carrying the maximum load themember might encounter during a specified
inspection interval or, for noninspectablestructure, during a lifetime. This load (Pxx) isdefined as a function of the specific degree of
inspectability in a given inspection interval.(Damage Tolerance Criteria, AFSG-87221 A)The changes to the commercial aircraft re-
quirements recommended by Boeing wouldbring them into line with the USAF require-ments.
The US Navy composite requirementsare, in some respects, more severe than eitherthe FAA or the USAF requirements. As anexample, the structure must withstand designultimate load, adjusted for the effects ofenvironment and material variability, with
clearly visible damage. The Navy specifies noyielding at ultimate load (versus limit load andmaximum expected flight load, respectivelyfor FAA and USAF requirements).
The Navy requirements currently pro-hibit use of honeycomb sandwich structure andaramid/epoxy. Because of corrosion caused bya mix of aviation fuel and salt water, the Navy
STRUCTURAL DESIGN
General. - The most successful compos-ite development programs have invariablyused an integrated engineering team approachto design, development, and production inte-gration. The team is usually structured toinclude personnel from design, analysis,materials processing, tool design, productionengineering, quality control, and, in someinstances, costing. Ideally, the team is collo-cated to facilitate communication.
The integrated team approach of itselfcannot guarantee success. The team membersmust possess the skills and experience neededto work with composites. Personnel withlimited experience will not be able to copewith the complexities inherent in compositesdesign and application.
In the past, the personnel involved in the
design of composites for a production programoften came from the R&D elements of the
organization. This should not be consideredunusual because the R&D organizations werethe first to work with and develop an under-standing of composites. There is also benefit inhaving engineering personnel cycle betweenproduct development programs and R&Dbecause exposure to the "real world" of prod-uct development makes them more cognizantof the constraints and specific requirements ofa production environment and could helpengineers formulate more focused reseai'chefforts.
Another mark of most successful pro-
grams is the implementation of a building-block approach to development. This approachimplies that the design, fabrication, and test ofmajor structural elements are taken in steps sothat potential problems with the design, either
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E
ET(3rr"0
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(3
,_Unlikely Loads for Damage Scenarios I
with High Probability of Occurance i
iUnlikely Scenarios withUltimate I More ProbableDamage Load Cases ]
structurallyor from thefabricationperspec-tive, areuncoveredat theearliesttime andwith thesmallestinvestmentin toolingandtestcomplexity.Each"block" buildson theknowl-edgegainedin thepreviousstep.Programmanagementmustunderstandthatcompositesareacomplexmaterialsystemandtheireffectiveuseis not a straight-forwardandsimpleengineeringeffort. New applicationswill requiremoretimeandeffort thanconven-tionalstructures.Thetime andotherresourcesrequiredto implementabuilding-blockap-proachmustbeapartof theprogramplan.Thebuilding-blockapproachhasbeenshowntoreducerisksand,in the long run,keepoverallprogramcoststo aminimum.
Theadvantagesof usingcompositeshasbeendemonstratedin manyprograms.Successhas,in somecases,led to overenthusiasm.Designersbenton makingthemostuseofcompositesin a newdesignoftentry to usecompositeswherethereis nobenefitto doingso.Oneexampleis to usecompositesfor manysmallpartssuchasclips andbracketswhenmetalswoulddo thejob effectivelyat muchlowercostandlittle, if any,additionalweight.Compositesshouldbeusedonly wherecarefulandthoroughstudiesshowaclearbenefit.Thestudiesmustincludetheavailabilityandcostof facilities,thecompany'sexperienceandability to implementtheir ideas,andthetradeoffof benefitwith risk.
Designersoftenoveroptimizetheirdesignsin aneffort to obtainmaximumweightsavings.Weight savingsshouldnotbemea-suredonapartby partbasisbutshouldbeassessedglobally. Highly optimizedstructuresdonot leaveroomto accountfor manufactur-ing discrepancies,inherentdefects,andinevi-tableloadincreases.As programrequirementschangeor difficulties arise,designersshouldhaveanacceptablefall-backposition.To thegreatestextentpossible,thedesignermustunderstandall of therequirementsatthestartof thepreliminarydesign.Theseincludethecustomer'sspecificrequirementsandorconstraints.Thecustomermust,afterall, bewilling to acceptadvancedcompositesandhavetheinfrastructureto supportthem inserviceincluding inspection,maintenance,andrepaircapabilities.
Companieshavefoundthattheproblemsinvariablyoccurduringtheearlyphaseofproduction.Theseproblemscanbemoresuccessfullyresolvedif designengineersareavailableatthemanufacturingfacility duringthefirst two to threemonthsof production.
Detail Design. - The building blockapproach is an excellent method for develop-ing and validating the details of the design.The final design must be validated at fullscale, however. In all aspects of the design, butin particular the design details, the designermust pay attention to producibility and dimen-sional tolerance requirements. Specifyingunnecessarily restrictive tolerances can driveup costs. On the other hand, maintaining closedimensional tolerances at mating surfaces cankeep assembly costs down and avoid thepotential for assembly-induced damage. Agood working knowledge of manufacturingprocesses or, if possible, experience in amanufacturing facility would not only behelpful to designers but may be essential if thecosts of composites are to be competitive withmetals. The design and the method of con-struction must be worked together, particularly
if some automated production methods arebeing considered. Again, the importance of theinput of manufacturing, tooling, and quality
10
Although it is well known that out-of-
plane loads should be avoided in compositedesign, failures still occur because of the
inadvertent introduction of out-of-plane loads.Joints, structural discontinuities, and otherareas of stress concentration must receive
careful attention. Bolted joints are the pre-ferred method for the introduction of out-of-
plane loads.
Generally, joints will be the criticalelement affecting component strength. Thedesigner must rely on a combination of de-tailed structural analysis and full-size elementtests to verify the design. Because jointstrength is dependent on actual laminatethicknesses and bolt size and placement,subscale tests cannot be used to verify a jointdesign.
Overall Design. - Composites havealways held the potential for reducing the partcount through the cocuring of large assem-blies, and a number of companies have hadsuccess with designs that minimize the numberof parts that have to be joined in separateassembly operations. The larger the part,however, the more likely that the part will
become complex and consequently the toolingmay also become complex and costly. Thebenefits of reduced assembly costs must beweighed against the possible increase intooling costs and the risks involved in curinglarger parts.
Early in the development of composites,designers tended to avoid post-buckled, stiff-ened skin designs. Experience has since shown
Theinspectabilityof structures,bothduringproductionandin service,mustbeconsideredin thedesign.Designingforinspectabilityincludesconsiderationof thetypeof equipmentthat will beavailableto thefield inspectionunits.As notedabove,largemonolithicstructureswith reducedpartcountaredesirable.This approachmustbe temperedwith theneedto inspectcritical interfacesthatcouldgetburiedwithin thestructure.Remov-ablefastenersshouldbeusedin areaswhereaccessto the interior is neededfor inspection.
STRUCTURALANALYSIS
Finite Element Analyses (FEA) have
been applied successfully to most compositedesigns. In regions of high stress gradients,such as around cut-outs and at ply and stiffenerdrop-offs, a fine mesh must be used. Except atcut-outs and discontinuities, general laminatetheory and the modeling of the laminate as anorthotropic plate are satisfactory. Atdiscontinuities, out-of-plane loads are intro-duced and a 3-dimensional FEA must be used.
The problem in many cases has been theidentification of a discontinuity. A number ofmajor structural failures have occurred duringdevelopment programs because the discontinu-ity was not adequately addressed in the analy-sis. In most all cases, a post-test analysis of thecritical region showed clearly the inadequacyof the design and guided a satisfactory rede-sign. Generally, detailed analyses should beperformed whenever there is an abrupt change
in load path or the load in the composite isbeing taken out or introduced through a fitting.Several designers noted the importance of
designing and analyzing the details first andthen filling in the design of the spaces be-tween.
Although teaming has become verycommonplace in today's environment, prob-lems can and often do arise when companiesworking on the same project are using differ-
ent analysis tools. An obvious solution is forall team partners to use the same analysistools. But, difficulties can still arise when teammembers are not familiar with the analysis
tools imposed by the team leader or when teammembers do not want to share their "propri-
etary" codes. When planning joint programs,the planners must be aware of the need forcoordination of analytical activities as well asproviding for the usual physical interfacecontrols.
STRUCTURAL TEST
Several techniques and criteria arecurrently in use for dealing with the effects oflow energy impact damage (LEID) and thedefinition and generation of barely visible
impact damage (BVID). Several organizationsfelt that techniques for generating damage,
As notedabove,subscalecomponentsarenotusefulfor evaluatingjoint designs.Theevaluationof designdetailsandthedemonstra-tion of damagetoleranceshouldalsobedoneat full size.
As discussedearlier,awell plannedtestprogramis an integralpartof thebuilding-blockapproachto development.Certificationtestsshouldbeapartof the initial plan.Expe-riencehasshownthatcompositesdonotpresentahigh technicalrisk whentheprogramincludesa 2-3yearbuilding-blockvalidationprogram.All of thestructureshouldbestatictested.Whenthestructuresatisfiesall thestatictestconditions,experiencehasshownthatfatiguewill not beaproblem.Techniquesfor takinginto accounttheeffectsof moisture,temperature,impactdamage,etc.during thestatictestingremainsa problem.An approachthat atleastonecompanyhasfoundsuccessfulis toevaluatetheeffectsof temperatureandmoistureatthecouponandsubcomponentlevel andmakecomparisonsbetweenpre-dictedandmeasuredperformance.Theseratiosarethenappliedin thedesignof thefull-scalecomponentto providethemarginsneeded.
COST CONSIDERATIONS IN DESIGN
There was nearly universal agreementthat costs and not weight savings have becomethe major driver in the application of advancedcomposites. For some military applications,improved performance, including stealth, canstill dictate the need for composites. But,aff, ordability remains a major design con-straint, even for military aircraft.
12
The need to achieve both weight savingsand retain affordability presents a challenge to
the designer. The real cost of a weight-opti-mized structure is often much higher thanestimated. Production costs, and particularlyassembly costs, should drive the design. Inorder to achieve a cost-effective design, an
integrated design and manufacturing team isessential. Often, slight changes in design canlead to part simplification and reduced manu-
facturing costs with little or no weight penalty.Designers must understand the ramifications
of design details on producibility and beprepared to explore non-traditional approachesin order to reduce costs. As noted in the
section on Assembly, monolithic, integralstructures can be used to reduce part count andassembly costs. Part commonality is another
effective method for reducing costs. Oftenslight changes in design can lead to signifi-cantly more part commonality. Designers mustunderstand that not everything on a so-called
composite airplane needs to made of compos-ites. Small parts, for example, can be veryexpensive when made of composites byconventional methods and metal may be themost cost-effective choice.
The high material costs associated withcomposites, about 25 times that of aluminum,means that manufacturers need to minimize
the amount of scrap. Again, close coordinationof design and manufacturing and an under-standing of the manufacturing methods avail-able can heIp to minimize scrap. In a produc-tion environment, for example, one manufac-turer was able to maintain a buy-to-fly ratio ofabout 1.1. High material costs are one reasonthermoplastics have not been used in any greatquantity. Also, the compressive strength ofthermoplastic composites has not been as goodas that of thermosets. Except in some specialapplications, the manufacturing cost savingsprojected for thermoplastics have not beenrealized.
New US and foreign tactical aircraft aremaking extensive use of composites for bothprimary and secondary structures. For thesenewer aircraft, composites make up about 20to 35% of the airframe weight. Regardless ofthe type of material used for the airframe, thecosts of military aircraft are considerablyhigher than those of commercial transports.
Basedoncurrentprogramcostestimates,thefly-awaypricesof somenewermilitary aircraftstructuresrangefrom $590/kg($1300/1b)fortheMcDonnellDouglasC-17and$680/kg($1500/Ib)for theLockheedF-22to about$1800/kg($4000/lb)for theNorthropB-2B.Thesepricesreflect the limited productionratesandquantities.In contrast,thepricesofcommercialturbojetandturboproptransportsrangefrom $90-140/kg($200-300/1b).(Aero-spaceFacts& Figures,1992-93;Hadcock,1985,1989;McCarty, 1990;Harris,W., 1993)Althoughit may appearthatthehighercostsassociatedwith compositesmightbeeasiertoacceptwithin currentoverallstructurecosts,compositeswill still haveto "buy theirway"ontotheaircraftby demonstratingimprovedperformanceat anacceptablecost.
Thecostsassociatedwith assemblyofmechanicallyattachedcompositejoints areveryhigh andrequireclosetolerancesatfaying surfacesandrigid control of thethick-nessesof thepartsto bejoined.Unlessmatchedtoolingor machiningis usedto matchfaying surfaces,liquid and/orstructuralshim-ming mustbeusedto preventthe introductionof out-of-planeloadsin thecompositepartsatthejoint duringassembly.Thehighcostsofspecialfasteners,their installation,thecontrol,inspectionandmeasurementof thethicknessesof thecompositeparts(toassureproperfas-tenerselection),andshimmingaretheprimaryreasonfor minimizing mechanicaljoints in thedesign.Thesecostissuesaredescribedin moredetail in thesectiononAssembly.
At present,theexperienceof the inte-grateddesignteamisthebesttool availableforkeepingcostsdown.Tools thatcanhelpthedesigneroptimizeon costsaswell asweightwouldbeausefuladditionto designtechnol-ogy.
DESIGN R&D NEEDS
Opinions on current R&D needs werequite varied. Some felt that most R&D pro-grams contributed little to the real problemsencountered in the application of composites.They felt that most research was unfocusedand failed to address the problems and issuesassociated with production developmentprograms. Some felt that research, rather thanfocusing on long range needs, should alwaysbe tied to a development program and focus onnear-term (3 to 5 year) needs and solutions.Others stated that a R&D program shouldconcentrate resources on full-scale compo-nents with real design features rather than oncoupon/design allowables programs.
Because fewer resources are going intonew (particularly, military) aircraft develop-ment, the present time was viewed by some asa good time to concentrate R&D on the devel-opment and improvement of overall compos-ites technology. This development wouldinclude integration of design concepts, mate-rial forms, and manufacturing methods.
Some time has passed since a forum hasbeen held involving Government, Industry,and Academia to specifically discuss presentand future R&D needs. One of the first forums
was Project Forecast sponsored primarily bythe Air Force and held in the early 1960's.This was followed in 1972 by Project Recastwhich served as a guide for composites re-
search through the 1970's. The most recentforum, conducted under the auspices of theNational Research Council, was chaired byProfessor James W. Mar and was completed inabout 1986. A major recommendation of thecommittee was for NASA to institute a bold
new initiative in composites aimed at thedevelopment of technology that would reducethe cost of composites. The present AdvancedComposites Technology (ACT) Program was adirect result of the Committee's recommenda-
tions. It is now appropriate to once againassess the status of composites, the near-termoutlook, and future R&D needs. The presentinterest in high speed transports and results ofstudies that indicate the need for composites toachieve the weight targets essential for eco-nomic viability are more reasons to address the
most appropriate course for composites R&D.
13
LESSONS LEARNED
14
a.
b.
c.
d.
e.
f°
g.
h.
Design and certification requirementsfor composite structures are generallymore conservative than for metal
structures.
There are no reported aircraft acci-dents involving failure of primary orsafety-of-flight composite structure.
Design and certification of compositestructures are expensive and costs aremuch higher than for metal structures.The effort and costs associated with
design and certification have oftenbeen underestimated.
All design and certification require-ments must be thoroughly evaluatedand understood at the start of the
program. Certification requirementsof the FAA and the various services
can differ and require different ap-proaches.
Design and certification test datagenerated under a military aircraftprogram has only rarely been transfer-
able to a commercial aircraft programor a program sponsored by anothermilitary department and vice versa.This practice has resulted in consider-able duplication of effort.
Successful programs have madeeffective use of integrated develop-ment teams that include personnelexperienced in design, analysis,
materials and processes, tooling,quality control, production, and cost
analysis.
Experience gained in R&D programsdoes not readily transfer to productionunless the people with the R&Dexperience participate actively in theproduction development.
Successful programs have used abuilding-block approach to develop-
ment. Program managers with priorcomposites experience usually under-stand the necessity of realistic sched-
i.
j.
k.
m.
n.
ules that allow a systematic develop-ment effort.
The use of a basic laminate familycontaining 0/90/+45 plies with aminimum of 10% of the plies in eachdirection is well suited to most appli-cations, generally assures fiber domi-
nated laminate properties, and simpli-fies layup and inspection.
The number of mechanical jointsshould be minimized by utilizinglarge cocured or cobonded subassem-blies. Mechanical joints should berestricted to attachment of metal
fittings and situations where assemblyor access is impractical using alterna-tive approaches (see also, the sectionon Assembly).
Large, cocured assemblies reduce partcount and assembly costs. If thecocured assembly requires overlycomplex tooling, however, the poten-tial cost savings from low part countcan be easily negated. Producibilitymust be a key consideration in thedesign.
Structural designs and the associatedtooling should be able to accommo-
date design changes associated withthe inevitable increases in designloads.
Standardization of techniques forinducing impact damage and assess-ing its effects would eliminate confu-
sion and permit direct comparison oftest data and transfer of results to
other programs.
Designing for producibility is gener-ally more cost effective than optimi-zation for weight savings.
MATERIALS & PROCESSES
MATERIALS
Nearly all composites engineers holdstrong views on material selection and proper-ties determination and how these have influ-
enced composite structures development. Mostagreed that a unified approach was needed fordetermining material properties. Earlier efforts
by NASA and continuing coordination of testmethods by SACMA are beneficial and permitdirect comparison of test data. Many wouldlike to see uniform material specifications andstandards for a few selected material systemsand suggested the Government could take thelead in defining these. There were others whobelieved just as adamantly that standard
specifications were not needed. Companiesadd to the problems of myriad specificationsby writing their own and asking materialssuppliers to show that their product will meetthose specifications. The cost to qualify a newmaterial and generate design allowables can beas much as $3-5 million.
The current Air Force Manufacturing
2005 program is looking at the potential costsavings in using commercial specifications formany of the systems they procure includingcomposites. Other efforts at standardization ofmaterials and processing specifications areunder way through the auspices of the AircraftIndustries Association (AIA) and the Great
Lakes Composites Consortium. Because of the
large number of material systems in use,standardization of repair materials wouldbenefit the user by reducing the number ofexpensive material systems that have to bestored and periodically monitored to assuretheir viability. Out of date materials must oftenbe scrapped, thus adding to the cost of com-
posite maintenance. As more and more com-posites go into service, the problem of main-taining replacement materials for repair andrebuilding of older structures is of increasingconcern. Generic composite materials that canbe used with a broad range of composite
materials are urgently needed for repair.
Development risks can be reduced by notlocking into a single fiber/matrix system,particularly if the fiber/matrix system is
relatively new. Preferably two material sys-tems that will meet all requirements should beavailable at the start of a product developmentprogram. Unduly high risks are incurred if amaterials development program is undertakenin conjunction with the product development.If some materials development is underwayduring the preliminary phases of a program,there should be a specific cut-off point early inthe program so that design can proceed withknown (and not projected) performance val-ues.
The Navy has had bad experience with
polyimide- and bismaleimide-matrix systemsin service. The combination of aviation fuel
and salt environment degrades the matrixmaterial. The Navy now considers thesematerials unacceptable on Navy aircraft.Bismaleimides are proposed for use on theF-22 in areas where there is no contact with
aluminum and moisture is not a problem. Theneed to isolate carbon fiber composites fromaluminum or steel because of galvanic interac-tion has been long recognized and has beendealt with effectively by using an adhesive
layer and/or a thin glass-fiber ply at fayingsurfaces.
Continuing concern over the effects of
low energy impact damage has led to thedevelopment of toughened epoxy systems, andthere is a strong tendency for designers to usethe latest and "best" material systems. In manyinstances, an untoughened system can do thejob reliably at a much lower cost than a tough-ened system. For example, the Navy has hadgood in-service experience with a well-charac-terized, untoughened epoxy system that hasbeen on the market for many years.
Material costs can be a significant por-tion of overall component costs. In addition tothe cost of the material alone, the part designalong with the method of construction canaffect the amount of scrap and the buy-to-flyratio. At a buy-to-fly ratio of 2:1, for example,carbon/epoxy material purchased at $16/kg($35/Ib) is actually costing $32/kg ($70/1b) on
the part, not including processing and lay-upcosts. Buy-to-fly ratios ranging from 1.1:1 to2.2:1 were quoted by interviewees for ongoingproduction programs. With the cost of alumi-num at less than $2.3/kg ($5/11o) and emerging
15
advancedalloyssuchasaluminum-lithiumprojectedto cost$5 -$7/kg($10-15/lb),thecostof compositematerialis still a significantfactorin their effectiveapplication.Costsassociatedwith incomingmaterialinspection,handling,andstoragearehigherfor compos-itesthanfor metals.Inspectionsdoneby thematerialsupplier,providedtheycanbecerti-fied to thesatisfactionof thebuyer,shouldnothaveto be repeatedatthemanufacturingfacility. Quality control costsusuallycanbereducedfurtherwhenautomatedprocesscontrolsareused.
Noneof thepeopleinterviewedwhohavehadexperiencewith thermoplasticswereoptimisticabouttheir viability for extensiveuseonaircraft.Currently,with costsrangingfrom $55-80/kg($120-175/lb),thermoplastic-matrixcompositesareveryexpensivecom-paredto mostthermosets.Furthermore,prop-erties,especiallyin compression,makethemof questionableusefor primarystructure.AlthoughconsiderableR&D fundinghasbeenappliedto thedevelopmentof thermoplastics,manyproblemsremainunresolvedandsomequestiontheefficacyof continuedGovernmentfundingfor their development.
Processing depends on material, configu-ration, tooling, and the fabrication methodbeing used. There are, however, some generalcomments that can be made and would applyto most materials and applications.
Process controls must serve two pur-poses: they must assure consistent propertiesand also provide dimensional control onthickness and overall geometry. The processcontrols must be compatible with the tooling.Tool design can affect heat-up rates, forexample, and dictate the extent to whichtemperature can be controlled. Materials that
permit a broad processing window can allevi-ate some of these problems, and also problems
16
associated with batch processing various partsin an autoclave.
Several techniques have been usedsuccessfully to control part thickness andoverall part quality. These include:
1. Material specifications that affectresin flow during cure.
2. Improved control of resin content inno-bleed resin systems.
3. Use of higher than normal pressureduring cure.
4. Frequent compaction cycles, particu-larly of thick parts.
. Intermediate partial cures for thickand/or large parts when material out-time could be a problem.
6. Post cure of all parts.
Continuing process control and processmonitoring are required during production toassure that neither the process nor the material
is changing. Tag-end specimens can be used tocheck the processing of every part. Verifica-tion of the process should include tests thatcheck the critical structural properties of alaminate, particularly for primary structure.Any proposed changes in processes duringproduction must receive careful evaluation andvalidation before being approved. Experiencedprocess engineers must be available to workwith designers in the early phases of theprogram and with manufacturing personnelonce the part starts into production. Processingproblems that occur during development andin production can often be categorized andrelated thus leading to some generic proce-dures and solutions that can be applied onfuture programs. Careful records must be keptand thorough assessments made in order togain the full benefit from past experience.
Engineering drawings and Materials &Processes specifications tend to be highlydetailed and complex. Although the detail isnecessary for thorough documentation, it is noteasy to follow during the fabrication processon the factory floor. Several companies have
developedhandbooksthatdescribethefabrica-tionof apart in astep-by-stepprocess.Engi-neeringdrawingsarereplacedby severalsuccessiveperspectiveviewsor cross-sectionsof apartshowinghow eachply or elementisto beinstalled,anddetailedmaterialspecifica-tionsarereplacedby shortnarrativesdescrib-ing theproceduresto be followed ateachstepof thefabrication.An example of a fabricationprocess sheet is shown in Figure 4. The de-scriptive manual aids in the inspection processas well as the fabrication.
A number of techniques were suggestedfor reducing fabrication costs. Some of theseincluded:
1. Use of no-bleed prepreg and adhesive
prepreg systems.
In some cases, however, secondary
bonding might have to be used. For example,if cocuring a large complex part with internaltooling, some portion of the part might pur-posely be left unbonded during the initial cureto permit access for removal of the internaltooling. If secondary bonding is used, greatcare must be taken to assure near perfect fit-upof the faying surfaces. Close-tolerance ma-chining of faying surfaces may be required
prior to bonding.
In all bonding and cocuring operations,problems such as core slippage and crushing,skin movement, and ply wrinkling can occur.Sometimes a two-step curing process can becost effective because of the significant im-
provements in quality and process repeatabil-ity.
.
.
.
.
Automated tow placement using no-bleed tows.
Pultrusion of constant section stiffen-
ers.
Resin injection molding (RIM) orresin transfer molding (RTM), usingstitched or woven preforms, for
complex shapes and small parts.
Use stack gas rather than nitrogen inautoclaves.
JOINTS AND ATTACHMENTS
As noted above, most companies favorcocuring or cobonding over secondary bond-ing. For example, cocuring was used veryeffectively on the F-15 speed brake wherecarbon/epoxy skins ranging in thickness from4 to 72 plies were cured and adhesivelybonded to aluminum honeycomb core in asingle autoclave cycle. Cocuring has also beenused effectively for the F-14 horizontal stabi-lizer and the F-18 wing skins where the com-
posite has been cured and bonded to steppedtitanium joint plates in a single operation.Cobonding has also been an effective method
of bonding precured composite stiffeners orframe sections during autoclave cure of anuncured skin.
Just as secondary bonding requires a near
perfect interface, mechanical joints alsorequire very close fit-up in order to preventany out-of-plane loads being induced byforcing adjoining surfaces into place duringassembly. Whenever possible, mating surfacesshould be tool surfaces to help maintaindimensional control. If this is not possible,either liquid shims or, if the gap is large, acombination of precured and liquid shims,should be used in all mechanically fastenedjoints. Another approach is to cocure thecomponent parts with a very thin steel sheetbetween the joint interfaces. The steel sheet isremoved after the parts have been cured, andassembly completed using mechanical fasten-ers.
LESSONS LEARNED
a.
b.
c.
Trying to conduct materials develop-ment in conjunction with a product
development program creates unduerisks.
Because of the high cost of compositematerials, designs and manufacturingmethods must attempt to minimize
scrap.
Experienced designers and processengineers must be readily available
N,0br°.ther B'0C/--A rP" Side mandrels --_/."___ F Separator
..................................Hat Mandrel J h Hat layup
NOTE: ASSEMBLIES -005, -007, -009, -011,
SIMULTANEOUSLY
LAYUP -500 FRAME
PLY NO. DASH NO. ORIENT MATERIAL
1 -029 0 DMS2224 W
-013 MAY BE WORKED
18
Figure 4. - Example of process description from shop manual.
d,
e.
f.
to help correct production problemsand assess and validate any proposed
changes in production processes.
Handbooks that pictorially describethe manufacturing process are easier
to interpret than engineering drawingsand result in fewer layup and process-
ing errors.
Cocuring and cobonding are preferredover secondary bonding. Secondarybonding requires near perfect inter-face fit-up.
Mechanically fastened joints requireclose tolerance fit-up. Liquid orstructural shimming is usually re-quired to assure a good fit and toavoid damage to the composite partsduring assembly.
MANUFACTURING AND
TOOLING
MANUFACTURING
General. - A recurrent theme throughoutall the discussions was the usefulness of a well
integrated design-build team. For productionprograms, the team should include mostly"generalists" with multidisciplinary experienceand training who understand the interactions
among the various disciplines and appreciatethe ramifications of decisions made at each
step of the development process. Specialists ineach field must be available to support theoverall effort, but there is no substitute for
experience.
Experience gained in an R&D program isnot easily transitioned into the productionenvironment unless the people who have the
experience are assigned to the productionprogram. Differences in personnel skills,component scale, "real-world" interfaces,schedules, and cost targets must be understoodand taken into consideration when production
program decisions are made based on priorR&D experience. Many R&D programs have
not addressed manufacturing technology,which is the key issue in affordability ofcomposites.
The building-block approach to develop-ment provides the best approach to solvingdesign and manufacturing problems andreducing program risk. The probability ofsuccess is greatly increased when a programincludes the time and resources for an inte-
grated engineering, process validation, testing,and manufacturing development program.Manufacturing development needs to take
place at a scale that will test and validate theconcepts proposed. The use of sub-scalecomponents, often as a cost-saving attempt,can prove risky and lead to erroneous conclu-sions.
Cost has become the principal concern in
the development, application and utilization ofcomposites, and rightfully so. All aspects ofthe manufacturing process must be considered
when determining costs. Nonrecurring costsinclude tool design and fabrication; software
for automated tape-laying machines (ATMs),tow placement machines, and nesting pro-grams for NC cutters; and manufacturingmethods and shop instructions. Recurring costsinclude composite and other material procure-ment and waste disposal; layup, autoclave cureand post cure; part trimming; installation ofdetail parts in assembly fixtures; specialfastener procurement; subassembly; and finalassembly. Generally, labor costs are slightlyhigher than material costs, but layup costs canbe reduced considerably through the use of
automation. Composite material costs are verymuch higher than the costs of aluminum alloysheet and plate and care must be taken tominimize composite material waste and scrap.
Simply building low-cost elements doesnot guarantee low cost structures if the ele-ments do not have the needed quality. Assem-
bly costs, cited by many as the major manufac-turing cost element, can soar if parts fit poorlyin assembly and an excessive effort is spent in
shimming, hole preparation, and fastenerselection. A seemingly simple, minimum part
count design will not be cost effective if thetooling becomes too complex and costly andpart layup and removal becomes difficult, time
19
consuming,andrisky. Some"build-to-print"programsintendedto fostercompetitionandreducemanufacturingcostsdid notachievetheir goalsbecausetheoriginal designdid notpayadequateattentionto producibilityandresultedin majormanufacturingproblems.Manufacturingpersonnelhaveoftenpridedthemselvesin their "can-do" attitudewheninsteadtheyshouldchallengedesignerstojustify adesignthatmakesmanufacturingdifficult. Producibility is akeyelementof costreductionandusuallycannotbeaddressedwith thefabricationof oneor two prototypes.A full-scaleproductionprogramis oftenneededto identify andcureproducibilityproblems.
Automation - As noted above, automa-tion of the layup process can produce costsavings. In general, however, automation isnot cost effective if production rates are low(1 aircraft per month, for example). Also,automation is most cost effective on largerparts and may not be cost effective at all onsmall parts. When automation is planned, thedesign should be optimized for the process tobe used. The process must meet the tolerancerequirements for ply placement, or else thedesign must be changed to accept the toler-ances achievable. Again, the need for integra-tion of all aspects of the design and the manu-facturing process is apparent.
Incorporation of automation into thefabrication process does not necessarily meanthat the part is automatically built to net shapeon a contoured tool. Automated methods have
been used effectively on a number of programsto lay down flat laminates that are thentrimmed to shape and placed onto the moldform. Good quality and repeatability has beenachieved using this process. For small parts,automated cutting of plies from broadgoodsand manual placement on the tool may be themost cost effective approach. Furthermore,automated processes, because of their repeat-ability, can often reduce quality control costs.
for hole preparation, measurement and inspec-tion of each hole, and the cost of specialfasteners. Also, fastener grip lengths must bebased on actual thicknesses (including shims)at the fastener location and not on nominalthicknesses. Fasteners made of materials suchas titanium or A286 stainless steel must be
used to avoid galvanic interaction with the
carbon fibers. In spite of the potential for highcosts, mechanical fasteners generally are usedin the assembly process and are preferred oversecondary bonding. If secondary bonding isused, enough fasteners should be used to carrylimit load without relying on the bond.
Large cocured parts reduce assemblycosts. The interfaces to other parts must becarefully controlled, however, to avoid the out-of-plane loads that can be induced if mis-matched parts are forced together. Also, gooddimensional control of thickness can helpreduce assembly time and costs. Becausetolerances are more critical for composites, thestructural designers and tool designers mustknow, early on, the capability of the proposedmanufacturing process to maintain dimen-sional tolerances. If necessary, design andtooling changes must be made to meet toler-ance requirements. Some extra effort spent ondesign andtooling can usually be recoupedthrough reduced assembly costs.
Shimming is commonly used to bringmating surfaces into alignment and somemanufacturers plan for 100% liquid shimmingat final assembly. If possible, the matingsurfaces should be tool surfaces in order to
maintain the best possible dimensional control.Most manufacturers believed that the cost of
quality tooling and dimensional control onmating surfaces was more than offset byreduced labor and inspection costs at assem-bly. The close involvement of manufacturing/assembly personnel in the design process isessential to strike the proper balance betweendesign requirements and the need for
producibility.
Assembly. - Some of the discussionpresented in the section on Joints and Attach-ments also applies to assembly. As statedabove, assembly costs can be high. Installationof mechanical fasteners can cost as much as
$100 per fastener because of the time required
2O
TOOLING
Tooling is a critical element of themanufacturing process. Nearly all manufactur-ers interviewed commented on tooling and had
experienced a tooling problem at one time oranother. Tool design, including tool materialselection, must be an integral part of the
overall design process, especially with cocuredstructures. Tools designed and built by anotherorganization without close coordination withthe design, materials and processes, andmanufacturing personnel that will ultimatelybe responsible for the product can lead toserious problems. The capability now exists tocouple the three-dimensional design of a partto the three-dimensional definition of toolingand include the capability to perform a finiteelement analysis of the tool with thermaleffects. At the present time, even with this
coupled design capability, most manufacturerswould expect some tooling changes to occurbefore or during the early phases of produc-tion.
Manufacturers have had success with a
number of different tooling materials. Alumi-num tools have been used successfully on
small parts, but are generally avoided for largeparts and female molds because of the thermalexpansion mismatch with CFRP. Invar is oftenused for production tooling because of itsdurability and low coefficient of thermalexpansion. Electroformed nickel also produces
a durable, high quality tool but is more expen-sive than some of the other materials. Steel or
Invar tools are needed for curing high tem-perature resin composites such as polyimidesand bismaleimides. Steel, although not asdimensionally stable during heating andcooling as Invar, has been used successfully ina number of applications. A summary of sometooling used on past and current programs is
given in Table 2.
The use of CFRP tools has been both
successful and disastrous. A key to the suc-cessful use of CFRP tooling is to build a tool
of very high quality. Most problems havestemmed from tools with poor surface qualityand internal porosity. Tools made from CFRPhave the obvious advantage of a coefficient of
thermal expansion that matches that of theCFRP part. Also, the heat-up of CFRP tools iscontrollable and uniform and the tools are
easily repaired. In general, CFRP tools arelighter than metal tools and therefore easier tohandle and transport. Although the life of aCFRP tool is probably less than that of a metaltool, some manufacturers have made up to 300parts on carbon/epoxy tools and as many as1000 on carbon/bismaleimide tools. Postcuring
APPLICATION COMPONENT TOOL TYPE TOOL MATERIALC-17
B-2
Boeing 777
V-22
A-6
Starship
Lear Fan
F-14
Main Landing Gear DoorsControl SurfacesInlet Ducts
Wing SkinsWing Substructure
OMLIMLOML
OMLOML/IML
SteelAluminumInvar
CFRPSteel and CFRP
Stabilizers
WingFuselage
Wing Skins
Wing SkinsFuselage SkinsDetail Parts
All Parts
Horizontal Stabilizer Skins
OML/IML
IMLOML
OML
OMLOMLIML
OMLFIML
IML
Invar
InvarCFRP
Steel
CFRPCFRPSteel
CFRP
Steel
Table 2. Examples of moldform tooling used on some past and current programs.
21
carbon/epoxy tools at 200 ° C has been foundto improve their stability and life.
The use of molded rubber and trappedrubber tools has had limited success and those
that have used it usually would not do so
again. Rubber can be used successfully in localareas as a pressure intensifier, such as insideradii on stiffeners of cocured structure.
All tools require periodic inspection toassure dimensional control. Although analysescan be done to predict the distortion or "springback" of a part after it is removed from a tool,
the problem is usually solved through trial anderror methods to define tool modifications.
The "spring back" problem is generally morepronounced on metal tools than on CFRPtools.
The decision on whether to use inner
(1ML) or outer mold-line (OML) toolingdepends on many factors. As stated earlier,tooling the interfaces that will later be joined
helps to maintain close dimensional tolerancesand usually simplifies assembly. Outer mold-line tools, on the other hand, generally providemore flexibility for design changes such asskin thickness changes. Often the decision touse OML tooling, particularly for parts with anair-passage surface, is based on the feeling thatthe mold surface will produce a better aerody-namic surface. The differences between IML
and OML surfaces may be inconsequential forthe application in question and should bechallenged if overall part quality and costcould be significantly affected.
Quality tools are essential to the produc-tion of quality parts. If possible, productionquality tooling should be used during thedevelopment program to validate the toolingconcepts and materials. The additional coststhat might be associated with quality toolingare more than offset by the benefits of produc-ing parts of consistently high quality. As onemanufacturer stated, "You can't make good
parts on bad tools."
22
LESSONS LEARNED
a.
b.
c.
d.
e,
f.
g.
h.
j.
Generalists with multidisciplinaryexperience are valuable assets to a
concurrent engineering team.
R&D experience can be best trans-ferred to a production program if thepeople with the experience are as-signed to the program.
The building block approach facili-tates process validation and manufac-
turing development.
Most of the costs of a composite partare associated with manufacturing.Effective use of automated processescan reduce both fabrication and
quality assurance costs.
Designing for producibility is essen-
tial. Assembly costs as well as partfabrication costs must be considered
when selecting a design and manufac-turing process.
Automation may not be cost effectiveif production rates are low or partsizes are small.
Dimensional tolerances are more
critical in composites than in metals.Dimensional control of mating sur-faces can reduce assembly costs andavoid damage to parts during assem-bly.
Selection of tool material is depen-dent on part size and configuration,production rate and quantity, andcompany experience.
Tools often require modifications.Tool designers should anticipate theneed to modify tools to adjust for partspringback, ease part removal, ormaintain dimensional control ofcritical interfaces.
Quality tools are essential to theproduction of quality parts and a costeffective element of low-cost produc-tion.
QUALITY CONTROL NONDESTRUCTIVE INSPECTION
GENERAL
Currently, costs associated with quality
control (QC) for composites are much higherthan for metal structures. The costs can range
from 15-40% of the manufacturing costs withthe higher rates associated with R&D or
development programs and the lower rateswith established production programs. Theeffort expended on review board actionsrelated to discrepancies is also higher for
composites. Much of the QC costs can beattributed to nondestructive inspection (NDI)
of completed parts. In many programs, thepolicy is to do 100% C-scan ultrasonic inspec-tion of all parts. In addition to in-process
inspections, post-assembly inspection isessential to verify assembly process and assure
the part has not been damaged in the assembly
process.
Techniques intended to reduce QC costshave been implemented by some manufactur-ers. These include:
1. Automated process controls to assurerepeatability.
2. Integrate QC with manufacture in realtime.
. Evaluate specifications to be certainthey are not unnecessarily restrictive.Relax on controls that have no direct
bearing on part quality.
. Review past Materials Review Board(MRB) actions to see if dispositioncould have made with out board
review. Alter criteria as appropriate.
5. Assure that QC issues are consideredas a part of design and producibility.
. Zone the structure based on structural
criticality and key the sensitivity ordegree of inspection to the structuralrequirements.
Ultrasonic C-scan is the most commonlyused NDI technique. Other techniques, includ-
ing X-ray, shearography, and thermography,are used to a lesser extent or in special caseswhere C-scan is not sufficient. The detection
of foreign materials that can find their wayinto a layup may require more than one NDItechnique. The techniques currently in userequire considerable manual interpretation ofthe results, usually by an experienced NDIengineer. The use of automation and expertsystems in the evaluation of NDI records couldhave some cost saving potential.
EFFECTS OF DEFECTS
Programs to establish the effects of
typical manufacturing defects should beinstituted early in the development program.The program should define QC and NDIaccept/reject criteria. Often the componentsmade during the early phases of the program,before all procedures, processes, and toolinghave been fine tuned, will contain defects.These articles should be tested to establish
boundaries on acceptable defects.
LESSONS LEARNED
a. Automated processes can help toreduce QC costs.
b. Focus inspection and controls onaspects of the process and part thathave a direct bearing on part qualityand performance.
c. Determine and understand the effects
of defects on part performance.
23
SUPPORTABILITY
GENERAL
During the course of this study, theindividuals visited and the discussions held
focused mainly on the design and manufactur-
ing aspects of composite structures. At thetime of the study, NASA personnel werevisiting airline operators in an effort to obtaintheir views and concerns about the applicationof composites on commercial aircraft. Resultsof those discussions are included in a paper byHarris (Harris, C., 1993).
As shown in the list of facilities and
individuals visited, two military logisticscenters were included in our visits. Just as the
comments of manufacturers and programoffice personnel were not specifically limitedto design and manufacture, comments fromlogistics personnel pertaining to design andmanufacturing have been included in the
. appropriate sections. The following discussion-- combines information from all sources, not
just logistics centers.
Civil and military aircraft must be re-painted at frequent intervals. Paint stripping isa special problem when composites are presentsince many commonly used solvents candamage epoxy matrices. For example, repair ofa Boeing 767 CFRP rudder that was damagedduring paint stripping was reported to cost$99,000 (Harris, C., 1993). Aircraft manufac-turers and repair depots are moving towardspainting with water-based paints and paintstripping by blasting with polyethylene beads.
IN-SERVICE DAMAGE & REPAIR
Service experience with compositeprimary and safety-of-flight structures hasbeen very positive: no aircraft have been lostdue to failure of composite structure. How-ever, secondary composite components are
being damaged repeatedly in airline andmilitary service. Much of the observed damageoccurs during aircraft servicing and mainte-nance that may be unrelated to the compositepart (Harris, C., 1993; Donnellan, 1991).Many parts can be repaired, but the cost and
24
time required for repair is much higher andlonger than needed to repair similar damagedmetal components. The extent and type ofdamage must be defined using NDT tech-niques. Damaged structure must then beremoved and prepared for repair. After therepair has been completed, it must again beinspected using NDI. Repair prepreg andadhesive materials are expensive. Manydifferent materials must be stocked that have
to stored in refrigerators or freezers and have ashort shelf life. Often, these materials cannot
be procured in small quantities and much ofthe repair material must be scrapped whenshelf life is exceeded. Special approved fasten-ers and spare parts are also very expensive andadd to the cost of repair.
In addition to composite control surfaces(rudders, elevators, spoilers, and flaps) beingused on Boeing's 757,767, and 737-300,Boeing's new model 777 will have compositehorizontal and vertical stabilizers and floor
beams. The 777 is expected to be ready forflight in 1994. On the other hand, Boeing plansto use fewer composite structures on its newderivative 737 (737-X) than it has on the 737-300. The new 737 is aimed at the small airline
market and typically the smaller operators willnot have the composite maintenance and repaircapabilities that the larger carriers have devel-oped.
Aluminum honeycomb (H/C) sandwich
construction has caused major problems on AirForce and Navy aircraft because of moistureingress, core corrosion, and in the case ofsupersonic aircraft, debonding of the skin-to-
core adhesive due to pressure build-up in thecore after moisture ingestion. Stabilizers andcontrol surfaces have had to be disassembled
and rebuilt after the core has been replacedwith corrosion resistarit aluminum core. The
Navy has prohibited use of any H/C sandwich
components on the V-22, F/A- 18E/F and allother new aircraft types. (Donnellan, 1991 )
Composite components on Navy aircraft
have performed welt in service, and most ofthe damage has been on secondary structuralcomponents as a result of handling or manu-facturing discrepancies. There have been somecritical shortfalls for repair technology R&Dand development of t20°C (250°F) repair
Operatorswould like to havemoreengineeringandmanufacturinginformationaboutspecificcomponentsaswell asfactoryrework,repairsandinspectionrecords.Manufacturer'sStandardRepairManuals(SRM)havebeencriticized asbeinginad-equateandmostrepairsfall outsideSRMguidelines.
LESSONSLEARNED
a. Supportability has not been ad-equately addressed during design.Composite structures must be de-signed to be inspectable, maintainableand repairable.
b° Most damage to composite structureoccurs during assembly or routinemaintenance of the aircraft.
c. Repair costs are much higher than formetal structures.
d. Improved SRMs, engineering infor-mation, and MRB records are neededfor in-service maintenance and repair.
e. Special long-life/low-temperaturecure repair materials are required.
f. Moisture ingestion and aluminumcore corrosion are recurring support-
ability problems for honeycombstructures.
CONCLUSIONS
AND RECOMMENDATIONS
Although composites technology hasmade great advances over the past 30 years,the effective application of composites toaircraft is still a complex problem that requires
experienced personnel with special knowl-edge. All disciplines involved in the develop-ment process must work together in real timeto minimize risk and assure total product
quality and performance at acceptable costs.The most successful development programshave made effective use of integrated, collo-cated, concurrent engineering teams. The
composition of the team should representdesign, analysis, materials and processes,tooling, manufacturing, quality assurance, costanalysis, and product support. Ideally, at leastone representative of each discipline wouldhave prior applicable experience in compositestructures. Program managers should under-stand and appreciate the need for a well-planned, systematic, development effortwherein the design and manufacturing pro-cesses are validated in a step-by-step or"building block" approach. Such an approachwill reduce program risk and is cost effective.
Producibility and supportability are thekey elements in the design of compositestructures. The design, tooling, and manufac-turing processes must function in concert toassure consistent high quality parts at anacceptable cost. Quality tooling is essential tothe production of quality parts.
Not all parts are suited to compositeconstruction. Their use should clearly warrantthe added care that must be taken at every step
of their development..
Some of the confusion and problemsrelated to certification stem from differences in
requirements set by the various certificatingorganizations. Much of the confusion could beresolved if the FAA and the military services
would jointly generate a uniform set of basiccertification requirements for compositestructures. Additional special requirementswould be needed for military aircraft compos-ite structures to meet uniquely military designconditions such as live fire, survivability,
Currentconcernsin the industryaboutthevalueof genericR&D programsneedto beaddressed.An industry-governmentforum,similar to ProjectRecastor theNationalResearchCouncil("Mar Committee")activi-ties,shouldbe institutedto assesstheefficacyof today'sresearchactivitiesandhelpsetthedirectionsfor futureefforts.
An interdisciplinaryteamof experienced,knowledgeablepeopleworking togethercanmakethetechnicalrisk of applyingcompositescomparableto thatof anyotheradvancedstructure.Weight savingsaloneareno longerconsideredsufficientjustification for usingcomposites.Compositestructuresmustbecosteffective.Theweightsavedandotherin-servicebenefits,suchasdurability or corro-sionresistance,musthaveenoughvaluetooffsetanyaddedcoststhatmayarisefrom theuseof composites.Optimizationfor produc-ibility andsupportabilitycansignificantlyreducethecostof compositestructureswithlittle, if any,weightpenaltyor lossof struc-tural performance.
26
APPENDIX
A CHRONOLOGY OF ADVANCED COMPOSITE APPLICATIONS*
Richard N. Hadcock
RNH Associates
Hundreds of different composite aircraftcomponents have been designed, developed, andproduced over the past thirty years. A few of thecomponents have been associated with technologydevelopment, were generic, and did not fly or gointo production. Other components were designedto assess in-service performance and were pro-duced in relatively small quantities. The majorityhave been production components, designed fromthe start in composites, that have provided up totwenty years' of reasonably trouble-free service.
The first structural composite aircraft compo-nents, made from glass fiber reinforced plastics(GFRP), were introduced in the 1950-60 timeframe. These components included the fins andrudders of the Grumman E-2A, helicopter canopyframes, radomes, rotor blades, etc. By 1967, theentire airframe of the small Windecker Eagle wasmade from GFRP. Since then, GFRP has become
one of the standard materials for light aircraft andlightly loaded structural components.
Many of the lessons learned in design andfabrication of the GFRP parts were used in theinitial development of advanced compositestechnology during the 1965-70 time period.
types of fiber reinforcement for access doors andfairings.
Epoxy thermoset-matrix composites havebeen used for most of the composite parts. A fewhigher temperature parts have been made frompolyimide or bismaleimide thermoset-matrixcomposites. Very few have been made with ther-moplastic matrices.
Details of some of these components andtheir associated successes, problems, and failuresare described in this Appendix and supplementsthe information obtained from the discussions with
industry and Government personnel.
MILITARY AIRCRAFT
Many composite structural components havebeen designed and produced during the past thirtyyears for US and foreign military aircraft. Most ofthese primary and secondary structures are listedby aircraft and component in Tables A-1 (a) and A-1(b). They include both development and produc-tion components. Tertiary structural components,such as landing gear doors, access doors, panels,and fairings are not included.
Boron filaments and carbon (or graphite)fibers first became available about 1965. Their
high compression strength and stiffness, in combi-nation with low density, enabled boron fiberreinforced plastics (BFRP) and carbon fiberreinforced plastics (CFRP) to be used instead of
aluminum for high performance airplane struc-tures. At about the same time, duPont introduced
Kevlar®, a low density aramid fiber. Aramid fiberreinforced plastics (AFRP), which have hightension strength but very low compressionstrength, have been used for some lightly loadedfairings and helicopter components and as hybridcomposites composed of two or more different
U.S. Military Aircraft. - Much of advancedcomposite structures research and development inthe U.S. during the past 25 years has been associ-ated with military aircraft applications. TheseR&D programs provided much of the technologybase for production of-composite aircraft struc-tures. During the 1965 to 1973 time period, boronfilaments were available at lower prices thancarbon (or graphite) fibers. Boron/epoxy also hadhigher specific strength and stiffness than the then-available carbon/epoxy materials. For these rea-sons, boron/epoxy (BFRP) was the advancedcomposite of choice in the late 1960s.
"This Appendix was prepared by Richard N. Hadcock, RNH Associates, from material that he had collected outside ofthe present contract effort. The material collected by Mr. Hadcock provides an excellent background for the current study andshows the growth of composites use on aircraft worldwide. The information is included here with Mr. Hadcock's permission.
27
Thefirst aircraft to fly with advancedcom-positecontrol servicesweresome50USAFMcDonnellF-4 aircraft that werefitted withboron/epoxyruddersin the late 1960s.About thesametime,a USNDouglasA-4 wasflown with aboron/epoxyflap.
Boron/epoxywasalsoselectedby GeneralDynamicsfor theF-111compositehorizontalstabilizerUSAF developmentprogram.Thecompositestabilizerwasflown in 1971.Boron/epoxyreinforcementwasalsobondedto theF-111D6-acsteelwing pivot fitting to reducestressandincreasefatiguelife of productionF-11ls. ThismodificationincorporatedaremovableCFRPfairing for inspectionof thepivot fitting.
In 1967,Grummanselectedboron/epoxyforanIRAD programto designandbuild awing boxextensionfor theFB-111.Thetechnologyandexperiencegainedon this andthe 1968USAFAdvancedCompositesWing Structure(ACWS)program,led to theGrummanF-14A boron/epoxyhorizontalstabilizer.Thestabilizerwasfullyqualifiedfor productionin 1969andflew for thefirst time in December1970.TheF-14horizontalstab!_izerwasthefirst advancedcompositesafety-of-flight componentto fly andgo intoproduction.(Lubin, 1971)
theMcDonnellF-15Ahorizontalandverticalstabilizersandrudders.TheF-15Aspeedbrakebecamethefirst carbon/epoxyproductionsafety-of-flight component.Thefirst F-15flew in 1972.Approximately1.2%of theF-15airframeiscomposite.McDonnell laterdesignedandbuilt aBFRPF-15wing underanAir Forcedevelopmentprogram,but it wasneverflown.
Both theGeneralDynamicsYF-16A proto-type,which flew in 1974,andit's competitorfortheUSAF Light WeightFighterprogram,theNorthropYF-17,hadCFRPstabilizers.GeneralDynamics,which hadcompletedaUSAF CFRPfuselagedevelopmentprogrambasedontheNorthropF-5A centerfuselage,wentonto designandbuild a CFRPYF-16A forward fuselagein1975.Thesecomponentswerestructurallytestedbut did not fly or go intoproduction.
Lockheeddesignedandbuilt two boron/epoxyreinforcedcenter-wingboxesfor theC-130.Thesehavebeenin servicesince1974.
Grummandesignedandbuilt very largecompositehorizontalstabilizersfor theB-1Abomber,which werestructurallytestedin 1976.These,andtheRockwell-designedverticalstabi-lizer,wereboth fully qualifiedfor theB-1Abomberprior to cancellationof theprogramin1977.WhentheB-1Bprogramwasrevivedin1981,Rockwell decidedto revertto theoriginalmetalstabilizers.
Elevenleft-handouterwingsof theVoughtA-7D weredesignedandbuilt underaUSAFprogramin 1973-76.Threewingswereusedforflight andgroundtests;theremainingeightwereput intoservicefor five yearsonAir NationalGuardairplanes.Thesewerethefirst compositewing componentsto entermilitary service.Aboutthe same time, Vought designed and made 28CFRP spoilers that were installed on Lockheed S-3aircraft for Navy in-service evaluation. (Manno,1977)
28
WING
BOX
C SO UV BE SR TS R
UCT
F
ORW
ARD
FUSELAGE
M R
I ED A
R
STAB.
I H V E
N O E LT R R EE I T V
R Z I AN O C TA N A OL T' L R
L
AIRCRAFT YEAR tGrummaan E-2 1960 PMD F-4E 1969
Douglas A-4 1969/70 D
Northrop F-5 1970 D DGD F-I 11/FB-I 11 1970/73 D D D DLockheed C-5A 1970NA F-100 1970 D D
Grumman F- 14A/D 1970/88 pl
HS Vulcan 1971MDF-15 1972 D 2 D 2 P P
Vought A-7D 1973 L LLockheed S-3 1973HS Harrier 1973 D 3
Northrop YF-17 1973 D DLockheed C- 130 1974 LGD YF/F- 16 1974/6 P P
Dassault Mirage III 1975Dassault Mirage FI 1976 DRockwell B-IA 1976 P D DMiG-29 1977 P PMD F/A- 18 1978 P P P
R A F S S AU I L L P ID L A A O RD E P T IE R L BR O E R
N R AS K
E
PL
D
D
DP
PL
D D D D
PD
P
P P P PP P
P P
DP P P
I I I I I I
tYear of first flight or completion of R&D program.
1 1970 P = Production
2 1975 L = Limited Production
3 Ferry wing tip D = Development
Table A-1. Composite components on military aircraft.
(a) 1960-1979
In 1978, the McDonnell Douglas F/A- 18Awas the first military aircraft to designed with an
advanced composite wing. The F/A-18 was aderivative of the Northrop YF- 17 with modifica-tions to meet Navy requirements. The weightincreases due to a new landing gear, arrester hook,
wing folding, etc., required a larger wing and
increased fuel capacity.
A composite wing was selected for the F/A-18 to save weight. The wing covers are CFRPmechanically attached to aluminum substructure.In addition to the wings, CFRP covers are used forthe horizontal and vertical stabilizers (produced by
29
Northrop), the center fuselage upper skin panels,
the speed brake, flaps, and various fuselage access
doors and panels. CFRP accounts for 9.5% of the
structure weight (Weinberger, 1977; Kandebo,1993a). The heavier McDonnell Douglas F/A-18E/
F, scheduled to fly in December 1995, has CFRP
wings that are 25% larger and stabilizers that are
36% larger than the F/A- 18A. Much of the center
AIRCRAFT YEAR'[
Alphajet 1980
Vought VSTOL A 1980Fuji MT-X 1981Grumman VSTOL A 1981MD AV-8B 1981Lockheed C- 141 1981
and rear fuselage is also composite. These changeswill increase the use of CFRP materials to 18% of
the structure weight (Kandebo, 1993a).
The McDonnell Douglas AV-8B had its
origins in the British Aerospace AV-8A, Harrier,two of which were modified into development
YAV-8Bs by McDonnell and were first flown in
FUSELAGE
M R'
I ED A
R
STABIL CONTROL SURFACESIZER
I H V E R A F S S AN O E L U I L L P IT R R E D L A A O RE I T V D E P T IR Z* I A E R L BN C T R O E RA A O N R AL L R S K
E
D 1 D 1 D2 D
P P
D D
P
DD D D DD DD D
p4 p4P PP P P PP P P P
D
D
D 3
P
D D
P P
D
D
D D
D D D D D DD
D D D D D D
P P P P PP P P P P P PP P P
P
D D D D D D DD D D D D D D D
D D D D DD D D D D
P P P PD D
P P P P PP P P P P
D D
P P P PP P P
D D D DP P
P PP P P P
I I I 1I I I I
D DP P P
P P
P D P P
I I I II I I I
PP P PP P P P
I I I 1 I I II I l I I I I
P
* Including canards? Year of first flight or completion of R&D program
I 1987, Dornier 2 1980, Dassault 3 Cargo doors 4 Boeing
P = Production
D = Development
Table A- 1. Concluded. (b) 1980-1995
3O
1978. The modifications including replacing thealuminum wing with a CFRP supercritical wingthat had 14% more wing area and 50% moreinternal fuel volume. The production AV-8Bs,which first flew in 1981, have a CFRP wing
The AV-8B inboard flaps, inboard fairings,and the strakes are high temperature carbon/bismaleimide. All the other composite componentsare carbon/epoxy. With the larger wing, the range-payload capability of the AV-8B is approximatelytwice that of the AV-8A. Advanced compositesaccount for 26.3% of the structure weight of theAV-8B.
About the same time, the US Navy became
interested in a multipurpose V/STOL aircraft toreplace the Grumman E-2 and the Lockheed S-3.Composite fuselage development contracts wereawarded to Vought (LTV) for development of rearfuselage structure and to Grumman for develop-ment of the center fuselage. Both these designswere capable of significant elastic post-buckling
capability and were also designed to be exception-ally damage tolerant.
Vought built a 6 ft full-scale section of therear fuselage using CFRP stiffeners, longerons,and bulkhead webs with AFRP skins. Grumman
built a 25 ft long, 10 ft deep, and 7 ft wide sectionof the center fuselage using CFRP skins thatincorporated GFRP crack-arrestment strips, rein-forced by integrally molded CFRP hat-sectionstiffeners and "J"-section frames. Tests at Naval
Air Development Center (NADC), Warminster,demonstrated both post-buckling capability andcapability to hold limit load with significant lowenergy impact damage as well as ballistic damage.
The first wing made from high temperaturecarbon/polyimide, the General Dynamics F-16XL,flew in 1982. The F-16XL has a gross weight of48,000 lb compared with the F-16C gross weightof 37,500 lb. The carbon/polyimide wing covers,
made using inner mold line (IML) moldforms, arebolted to aluminum substructure. The F-16XL
wing is twice the area of the standard F- 16 wingand carries 80% more internal fuel. The F-16XL
did not go into production; the two prototypeairplanes were acquired by NASA in 1989 and are
currently being used for flight test programs. TheJapanese FS-X, which is under development, is aderivative of the F-16 and will have a Mitsubishi-
designed composite wing.
The Grumman X-29A forward-swept-wingtechnology demonstrator aircraft first flew in1984. The X-29A wing had CFRP covers me-
chanically attached to a substructure composed oftitanium and aluminum. The X-29A wing is
divergence critical so the wing covers were de-signed using aeroelastic tailoring to precludedivergence by coupling wing bending and twist.This coupling was accomplished by orienting theouter wing cover laminate axis at a different anglethan the wing geometric axis. Two X-29A aircraftwere built and are now in storage at NASA
Dryden Flight Center after completing about eightyears of flight testing. (Hadcock, 1985)
The CFRP wing of the Navy Grumman A-6Ewas designed and built by Boeing to replace thealuminum wing, which had a relatively shortfatigue life. The A-6E wing is much more com-plex, larger, and more highly loaded than the wingof the AV-8B. Its geometry is identical to themetal A-6E wing, with the same wing-to fuselageattachments, fold joints, store stations, and controlsurfaces.
Design studies of the USAF AdvancedTechnology Bomber (ATB) were initiated in 1979.The ATB became the B-2A when the Northrop/
Boeing/Vought team won the development con-tract in 1981, and the first flight took place fromPalmdale, California, to Edwards AFB on July 17,1989. Almost all the skin and much of the sub-
structure are CFRP and other composite materials.The B-2A wing has a span of 172 feet and a wing
area of 5,140 square feet. The span and wing areaare only slightly less than the wing of a Boeing747 (Jane's All the World's Aircraft, 1993-94 and
prior ed.).
Northrop, the B-2A prime contractor, wasresponsible for overall design, and designed andmade the leading edges, the crew station assembly,wing tips, elevons, and fixed trailing edge assem-blies. Boeing designed and manufactured the outerwings and the wing-fuselage center section.Vought designed and manufactured the intermedi-ate wing section that included the engine bays,inlets, and the main landing gear bays and doors.Final assembly was performed at the Air Force
31
plantin Palmdale,CA. At thestartof theprogram,Northropsetup anadvancedCAD/CAM systemthatwassharedby BoeingandVought.
In 1986,theLockheedYF-22A andtheNorthropYF-23A wereselectedastheUSAFAdvancedTacticalFighterDemonstration/Valida-tion (Dem/Val)programwinners.Both aircrafthadmadeextensiveuseof CFRPcomposites.TheYF-22A wasdesignedandbuilt by aconsortiumcomposedof Lockheed,Boeing,andGeneralDynamics.Northropteamedwith McDonnell fortheYF-23A.
TheNorthropYF-23A first flew in June1990followedtwo monthslaterby theLockheedF-22A.Lockheedwasselectedto proceedwith theF-22AEMD programin April 1991.Lockheedhasoverallprogramresponsibilityaswell asdesignandfabricationof theforward fuselage,inlets,wing leadingedgeflaps,trailing edgeaileronsandflaperons,andtheverticalandhorizontalstabiliz-ers.GeneralDynamics(now partof Lockheed)isresponsiblefor thefuselagecentersectionandBoeingisresponsiblefor thewing (Jane's).
BoththeYF-22A andtheF-22A haveCFRPwing, fuselage,andempennagesurfaces.Ad-vancedcomposites,titanium,andaluminumaccountedfor 23%,23%and35%, respectively,ofthestructureweightof theYF-22A. Thecompos-itesportionwas13%thermoplastic(TP) matrixand10%thermoset(TS) materials.Following livefire testsandareevaluationof theTPmaterials,thedistributionon theF-22A will be26%compos-ites(22%TS,4%TP),30%titanium,and 14%aluminum.TheF-22Ais scheduledto makeitsfirst flight in 1995with full productionstartingin1998(Morrocco,1993a,1993b).
Muchof theairframeof theGeneralDynam-ics/McDonnellDouglasA- 12Navy attackbomberwasgoingto bemadefrom carbon/bismaleimidecomposites.This programwasterminatedby theNavy in 1991becauseof scheduledelays,costoverruns,andNavyconcernsaboutcorrosionoftheBMI matrix in ajet fuel/seawaterenviron-ment.
TheUSAF C-X programfor a heavyliftmilitary transportwaswon by McDonnellDouglasin July 1982andbecametheC-17A.Full scaledevelopmentwasapprovedin February1985andthefirst developmentaircraft first flew in Septem-ber 1991.Four airplaneswereonorderin FY 1992
32
andeightin FY 1993.Compositecomponentsconsistof controlsurfaces,fairings,andenginenacelles,whicharesimilar to thecomponentscurrentlyin serviceoncommercialairliners.About15,000Ib of compositematerialsareusedin theC-17A,accountingfor some8%of theairframeweight.Most of thecompositecomponentsaresuppliedby subcontractors.Thecomponentsandmanufacturersinclude(Jane's;Parker,1989a):
Foreign. Military. Aircraft. - Some of themajor European aircraft companies initiatedadvanced composite structures development
programs in the early 1970s. Foreign militaryaircraft and their associated composite componentsare included in Tables A-1 (a) and A-l(b).
The British government and aircraft industryhave been involved in development of polyacry-lonitrile (PAN) precursor CFRP structures since1964, when the Royal Aircraft Establishment filedthe patent for high-strength, high-modulus PAN-based carbon fibers. British Aircraft Company
(BAC) and Hawker Siddeley Aviation (both arenow part of British Aerospace PLC) were involvedin structural development of CFRP components. In1968, Hawker Siddeley began a CFRP structures
development program that resulted in design,fabrication, and flight test of a Harrier ferry wing
tip (which incorporated an additional fuel tank), anairbrake flap for the Vulcan bomber, and six CFRPrudder trim tabs that were installed on BAC Jet
Provost trainers to obtain service experience
(Sanders, 1971; Fray, 1991; Molyneaux, 1978).
About 1974, BAC initiated a cooperative
program with Grumman to develop CFRP enginebay doors for the Jaguar ground attack aircraft.This was followed by the BAC-MBB TornadoCFRP Taileron development program and theBAC CFRP wing development program, whichused the Jaguar wing as the baseline.
Around the same time, Hawker Siddeley wasworking with McDonnell Aircraft in the develop-ment of the AV-8B Harrier. Hawker Siddeley(now BAe) manufactures all the CFRP horizontalstabilizers for the U.S. AV-SB as well as theBritish Harrier GR Mk 5 and Mk 7 V/STOL close
support aircraft.
The BAe Experimental Aircraft Programme
(EAP) technology demonstrator program wasinitiated in 1982 as a result of joint British-Ger-man-Italian European Combat Aircraft studies.The German team members, MBB and Dornier,
dropped out of the program in 1983, but BAe andAeritalia went ahead with the program. Flight
testing began in 1986. The wing covers and sub-structure were all CFRP and were made using theBAe co-bonding process. One wing was built by
BAe in England, the other by Aeritalia in Italy.The canards had CFRP covers bonded to honey-comb/metal substructure. The remainder of theairframe and control surfaces were metal
(Braybrook, 1986).
In 1985, SAAB contracted BAe to designand manufacture the first 3 ship-sets of CFRP
wings for the JAS 39 Gripen multi-role lightfighter. SAAB is manufacturing the productionaircraft wings and designed and built the remain-der of the structure, which includes a GFRPvertical stabilizer, canards, control surfaces, and
landing gear doors. CFRP materials account for30% of the structural weight of the Gripen(Braybrook, 1986). The first SAAB Gripen flew inDecember 1988 but crashed in February 1989 justbefore landing due to loss of control. The test pilotsurvived the crash with very little injury, possiblybecause the composite wings broke off cleanly asthe aircraft hit the runway (Jane's).
SAAB started an advanced composite devel-
opment program with a British company in the1970s to design and build a CFRP elevator for thecanard of the Viggen. The elevator developmentwas followed by a cooperative program withGrumman to develop a CFRP vertical stabilizer for
33
theSAAB JA 37Viggen.Grummanmadetheprototypestabilizers,which incorporatedpartsdesignedandfabricatedby SAAB. Thesewereflight andgroundtestedby SAAB who thenproducedsome50stabilizersfor theSwedishAirForceto obtainCFRPstructuresin-serviceexperi-ence.
Avion MarcelDassault-Br6guetAviation(Dassault),with financialassistancefrom theFrenchgovernment,starteddevelopmentof CFRPstructuresfor theirmilitary aircraft in theearly1970s.Their first designwasthe rudder of theDassault Mirage III fighter. One flight test and oneground test article were built and tested in 1975.This was followed by the Mirage F1 CFRP hori-zontal stabilizer program in 1976. Four prototypestabilizers were built and flight and ground tested.Dassault then decided to design CFRP ailerons for
the F1. The prototypes were flight and groundtested in 1977. These were 26% lighter than thealuminum ailerons, so they were put into produc-tion at a rate of 7 ship sets per month. More than730 Mirage F1 had been produced by the begin-ning of 1990 (Chaumette, 1982).
:::-:_Dassault introduced the Mirage 2000 multi-role fighter in 1978 and the Super Mirage 4000 in1979. Both aircraft had CFRP vertical stabilizers
and rudders (the Mirage 4000 stabilizer was also afuel tank), inner and outer elevons, and landinggear doors. The Mirage 2000 avionics access doorand the Mirage 4000 canards were also CFRP.Only prototypes of the Mirage 4000 were built, but
approximately 500 Mirage 2000 fighters have beenproduced since 1984 (Chaumette, 1982).
Dassault followed with a program to developa CFRP horizontal for the Br6guet-Dornier AlphaJet. The stabilizer utilized CFRP covers bonded to
full-depth honeycomb core outboard, transitioningto a bolted multi-spar substructure inboard (similarto the SAAB Viggen Vertical Stabilizer design).The stabilizer was designed using an automatedoptimization computer program and was projectedto be lighter and less expensive than the metalbaseline (Chaumette, 1982).
About 1980, Dassault and Aerospatiale were
sponsored by the French government to design,build, test, certify and fly a composite wing for theDassault Falcon 10. The development was jointlyfunded by Dassault and Aerospatiale. The wingwas fully certified and was flown in the mid-1980s. The Falcon 10 wing program provided
Aerospatiale with the know-how and confidenceneeded for the ATR 72 CFRP wing, described
later in this section. It also gave Dassault theknow-how and confidence needed to baseline a
CFRP wing for the design of the Rafale(Chaumette, 1982).
The Dassault Rafale 'A' prototype first flewin July 1986, and the first production Rafale 'C'flew in April 1991. The Rafale is a land-based or
carrier-based multi-role fighter and is about thesame size and weight as the F/A-18. CFRP compo-nents account for about 35% of structural weightand include the wing, canards, vertical stabilizer,control surfaces, landing gear doors, and somefuselage panels. Significant use is also made ofaluminum-lithium alloys and superplastic-formeddiffusion-bonded (SPF/DB) titanium parts for theleading edge flaps and hot fuselage structure(Interavia, 1985).
Dornier used the Alpha Jet wing as thebaseline for their CFRP wing development pro-gram during the mid 1980s. The Alpha Jet groundsupport/trainer aircraft was used by both Dornierand Br6guet as a test bed for CFRP structures,
shown shaded in Figure 6-lB. The Alpha Jet winghad CFRP spars and integrally stiffened coversand with aluminum ribs and fuselage attachmentplates (Rose, 1986).
The latest Western European fighter toutilize significant quantities of CFRP structure isthe British-German-Italian-Spanish EurofighterEFA (European Fighter Aircraft), which first flewin 1992. CFRP structures include the wings,forward fuselage, vertical stabilizer, and controlsurfaces. The canards are metal (Jane' s).
The Israeli Aircraft Industries (IAI) Lavifighter full scale development program was startedin October 1982. Grumman was contracted to
design and build the GFRP wings and verticalstabilizer and IAI designed and built the CFRPcanards. The first prototype aircraft flew on De-cember 31, 1986.
Following the Lavi flight tests, the winggeometry was changed and the areas of the controlsurfaces were increased. This required a majorwing redesign. The original wing had CFRP skinsbolted to CFRP substructure. The redesigned wingretained CFRP for the covers, but the substructure
was changed to aluminum. There was a significantcost saving and only a small weight penalty chang-
34
ing to aluminumsubstructure.SincetheLaviproductionprogramwascanceled,IAI hasusedtheaircraftfor technicaldevelopmentandflighttesting(Jane's).
Mitsubishi proceeded with CFRP primarystructures development and is currently designingand building an all CFRP wing for the FS-X.Mitsubishi was appointed prime contractor of theFS-X program, which is a modified GeneralDynamics F-16C. The first of four prototypes isdue to fly in 1996. General Dynamics, Kawasaki,and Fuji are the major subcontractors.
COMMERCIAL TRANSPORT AIRCRAFT
Most of the composite structural componentsdesigned during the past twenty-three years for USand European commercial transports and business
and private aircraft are listed by component type inTable A-2. These include both development and
production primary and secondary structures.Tertiary structural components, such as wing andstabilizer fixed leading edges and trailing edge
cabin floors, arid engine nacelles and inlets are notincluded in the table.
U.S. Turbojet Transports. - Most of the USairliner components, designed and built during the1972-1986 time period, were developed under theNASA Langley Research Center Aircraft EnergyEfficiency (ACEE) program and a predecessorflight service evaluation program to establish along-term durability data base for compositematerials and structures. The NASA programsincluded limited production and airline serviceevaluation of various components. It also includeda program to determine the long-term effects ofexposure to moisture, ultraviolet radiation, fuels,and hydraulic fluids on the mechanical properties(NASA CP-2321, 1984).
The first airliner advanced composite compo-nent to fly was a Boeing 707 boron/epoxy fore-flap, which was flown in 1970.
Under the NASA flight service evaluationprogram, Boeing designed and built 108 carbon/epoxy spoilers that entered service on Boeing 737aircraft with six different airlines in 1973. Some
spoilers were later made from different carbon/thermoplastic materials, but had to be taken out ofservice because the matrix was degraded byhydraulic fluid (NASA CP-2321, 1984).
Under the NASA ACEE program, Boeingcontinued with ten CFRP elevators that entered
service in I980. These were followed by fourBoeing 737 CFRP horizontal stabilizers that wereinstalled on two aircraft in March 1984. Thesehorizontal stabilizers were the first commercial
transport CFRP primary structures certified forairline service. (NASA CP-2321, 1984)
Douglas Aircraft designed and built thirteenCFRP upper rudders and three boron/aluminum aftpylon skins for the DC-10 under a NASA pro-gram. Additional rudders were built under theACEE program. The rudders first entered servicein 1975. Some of the upper rudders, which are amulti-rib post-buckled design, are still in service.This design approach was later used for the CFRPMD-l 1 ailerons. The aft pylon skins were the firstmetal matrix composite components to enterairline service (NASA CP-2321, 1984).
Douglas followed by designing and building
a multi-spar vertical stabilizer for the McDonnell
35
I AIRCRAFT [ YEARt
WING [ FUSELAGE I STAB"BOX CONTROL SURFACES
C S F M R I |H V E R A F S AO U O I E N|O E L U I L P IV B R D A TIR R E D L A O RE S WR T AS R R
U DCT
R Eli T V D E P IR |Z I A E R L BN|O C T R O E RAIN A O N R ALIT' L R S K
tYear of first flight or completion of R&D program.
1 1984 P = Production
2 1973 L = Limited Production
3 Tail cone D = Development
Table A-2.- Composite components on commercial transport aircraft.
Douglas DC-10 in 1977. The stabilizer was still inservice with Finair in June 1993.
Lockheed participated in the NASA flightservice program with eighteen AFRP fairing
panels for the L-1011. These panels entered airline
service in 1973. In the ACEE program, eightCFRP ailerons were designed by Lockheed andbuilt by Avco and entered service in 1982 (NASACP-2321, 1984). Lockheed also designed and builta CFRP L-1011 vertical stabilizer. The vertical
stabilizer failed during static test partly because of
36
Boein_ 737 Boeinlz 757 Boeing 767 Boein_ 777CompositeComponentAileronsElevatorsRudder
CASA (Spain)HDH (Australia)ATA (Australia)GrummanGrumman
Alenia (Italy)Rockwell
JADC (Japan)Shorts (UK)
JADC (Japan)Various
Table A-3. Boeing composite component suppliers.
the method of load introduction. Extensive long
term environmental and cyclic load tests wereperformed on spar and skin panel components, butno flight articles were built.
Boeing started design of both the model 757and the 767 in the late 1970s. The 767 made its
first flight in 1981 followed by the 757 in 1982.Boeing decided to baseline CFRP composites forthe elevators, rudders, spoilers, landing gear doors,and engine cowlings for both these airplanes. The
flaps of the 757 are also CFRP. Composite compo-nents account for about 3,400 lb of structure onboth the 757 and 767.
When Boeing introduced the 737-300 in1985, CFRP composites were selected for aileronsl
elevators, the rudder, fairings, and engine cowldoors. Composites account for 1,500 lb of thestructure.
The Boeing 777, scheduled to fly in 1994,utilizes about 18,000 lb of composites, some 9000lb of which are CFRP. CFRP components includethe entire tail, control surfaces, floor beams, main
landing gear doors, and engine nacelles. Othercomposite components include wing-fuselagefairings, and wing fixed trailing edge panels. TheCFRP horizontal and vertical stabilizers are made
by Boeing. Many of the other composite compo-nents are being supplied to Boeing by U.S. andforeign subcontractors (Table A-3).
The structure of the McDonnell DouglasMD-11, which first flew in January 1990, includesalmost 9,500 Ib of composites. Componentsinclude CFRP elevators, winglets, ailerons, out-
board flaps, spoilers, wing fixed trailing edgepanels, tail cone, engine cowls, center engine inletduct, cabin floor beams, and AFRP/GFRP wing-body and aft body fairings. Almost all thesecomposite components are produced by subcon-tractors. Some of the suppliers include: Fuji(Japan): outer ailerons; Embraer (Brazil): outboardflap; Mitsubishi (Japan): tail cone; Westland (UK):flap vanes; Heath Techna (US): center engine inletduct (Therson, 1989, Colucci, 1991).
Foreign Turbojet Transports. - The Europeanconsortium, Airbus Industrie, uses about 4,000 lb
of composites on the A300, which first flew in1972. A300-600 components include CFRP/GFRPelevators and rudders, CFRP spoilers, nose landinggear doors, and main landing gear leg fairings,GFRP wing upper surface trailing edge panels, andAFRP wing-body fairings and flap track fairings.
Use of composites was extended in 1985 bychanging the Airbus A310 vertical stabilizermaterial from aluminum alloy to CFRP. The CFRPvertical stabilizer, built by MBB (DA) in Ger-
many, also contain a balance fuel tank on the long-range A310-600. About 7,400 lb of compositestructures are used on the A310.
Composites use was further extended in 1987on the Airbus A320. Both the horizontal andvertical stabilizers are CFRP as well as the eleva-
tors, rudder, ailerons, spoilers, flaps, wing leadingand trailing edge access and fixed panels, landinggear doors, and engine cowls and doors. Fairingsare GFRP and AFRP. Composites account forabout 9,000 lb or 15 % of the structure of the
A320. The larger Airbus A330/A340 uses compos-
37
Airbus A300-600 Airbus A310-300 Airbus Airbus 330/340320/319/321
CompositeComponentHoriz Stabilizer
Vert Stabilizer
AileronsElevatorsRudder
Inboard SpoilersOutboard SpoilersInboard FlapsOutboard FlapsWing LE PanelsWing TE Panels
Ldg Gear Doors:MainNose
Fairings:FWD Wing-Fuse
Rear Wing-FuseNacelles
CASADA*
AerospatialeDA
BAe**BAe
FokkerCASA
Aerospatiale
Rohr (US)
DA
CASADADADA
BAeBAe
CASACASA
AerospatialeBelairbus
Rohr (US)
BAeCASA
CASADA
AerospatialeCASA
DABAeBAeDADABAeBAe
CASA
Aerospatiale
AerospatialeAerospatialeRohr (US)
DABAeBAeDA
Textron (US)BAeBAe
ATA (Australia)Fokker
AerospatialeAerospatialeRohr/Grumman (US)
*Deutsche Airbus **British Aerospace ***Construcciones Aeromiuticas, S.A.
Table A-4.- Airbus composite component suppliers.
ites for similar components, but, although the totalweight of composite structure is much higher, thepercentage weight dropped to 12% (Parker,1989b).
The Airbus A320 and A330/340 CFRP
horizontal stabilizers are designed and are built inMadrid by CASA, Airbus Industrie's Spanishpartner (Barrio Cardaba, 1990, Marsh, 1991). TheA310, A320, and A330/340 CFRP vertical stabi-
lizers are designed and built by the DeutscheAirbus division of MBB (DA) at Stade in Ger-
many. Suppliers of composite components toAirbus are listed in Table A-4 (Jane's).
The Soviet nushin 11-86 wide-body airliner,which entered limited service with Aeroflot in
1980, had CFRP cabin floors. The derivative 11-96,
which first flew in 1988, has CFRP flaps and cabinfloors. The horizontal and vertical stabilizer
leading edges are also composite but are probablyGFRP or a CFRP/GFRP mix (Jane's).
The Tupolev Tu-204 medium-range airliner
structure is about 18% by weight composites(approximately 20,000 pounds). The Tu-204,which -first flew in 1989, is the Russian counterpart
of the Boeing 757. CFRP components includespoilers, airbrakes, flaps, elevators, and the rudder.
38
Other composite components include part of the
wing skins, stabilizer leading edges, and wing-fuselage fairings (Jane's).
Turboprop Transports. - Advanced compos-ites are used extensively for control surfaces(ailerons, elevators, rudders, and flaps) of turbo-
prop regional transport aircraft. Many of these aremade from AFRP or a mix of AFRP with local
CFRP reinforcing.
The de Havilland Canada Dash 8, which hasbeen in airline service worldwide since 1984, uses
AFRP for the wing and stabilizer and flap leadingedges, wing tips, flap trailing edges and shrouds,wing-to-fuselage fairings, and engine nacelles.AFRP components account for about 2,000 lb, or10% of the structure weight. Approximately 350Dash 8s are currently in service.
The first commercial transport airplane witha CFRP wing to enter airline service (in 1989) wasthe Avions de Transport R6gionale ATR 72.
Avions de Transport R6gionale is a French/Italianconsortium composed of A6rospatiale and Alenia.The ATR 72 is a derivative of the 42-passengerATR 42, which was certified and entered airline
service in 1985 (Pilling, 1988).
The complete outer wing of the ATR 72 ismade from CFRP instead of aluminum because the
wing weight was critical to operating performanceand costs. The wing, designed and manufactured
by A6rospatiale, incorporates fuel tanks and the
design is based on much of the experience which
A6rospatiale gained from their joint program withDassault from the Falcon 10 CFRP wing program.
A6rospatiale manufactures the wing compo-nents that are assembled at Toulouse, France.
Other ATR 42 and ATR 72 composite components
produced by A6rospatiale include the wing fixed
leading edge and trailing edge panels, ailerons,
wing-to-fuselage fairings, and engine nacelles. *
Alenia manufactures the fuselage and tail of
both the ATR 42 and ATR 72 in Italy. Composite
components include the elevators, rudder, landing
* Private communication on composites use on the ATR 72
supplied to R. Hadcock by Avions de Transport REgional,
Blagnac Cedex, France, November 1993.
gear doors and fairings, and tailcone. Final assem-
bly of the ATR 42 and ATR 72 takes place at
Toulouse (Pilling, 1988).
The first commercial transport to enter
airline service that has significant portions of the
fuselage made from CFRP was the German DA
Dornier 328. The rear fuselage, pressure bulkhead,
and nose cone are made from CFRP. The wing-
fuselage and the main landing gear fairings anddoors are made from mixed CFRP/AFRP. The
wing flaps, ailerons, and wing tips, as well as the
complete tail of the Dornier 328, are all CFRP.
CFRP/AFRP mixed composites are used for the
wing fixed trailing edges and the dorsal fin
(Jane's).
GENERAL AVIATION
Business Aircraft. - Business aircraft are
listed in Table A-5 together with their associated
composite structures components. The GermanClaudius Dornier Seastar light amphibious flying
WINGBOX
C SO UV BE SR TS R
FUSELAGE STABIL- CONTROL SURFACESIZERS
S I H V E R A FH N O E L U I LE T R R E D L A
L E I T V D E PL R Z* I A E R
N C T R OA A O S NL L R S
S
AIRCRAFT YEARt
LearFan 2100 1981 L L L L
Dassault Falcon I0 1983 D D
Dornier Seastar VT01 1984 D D
Dassault Falcon 50 1984
Avtek 400 1984 D D D D
Gulfstream IV 1985 P
Beech Starship 2000 1986 P P P P
Piaggio Avanti 1986
Cessna Citation 1987
Dornier Seastar CD2 1987 P P P P
P P P P
I 6 ] 6 I 6 I 7
Avtek 400A 1991
TOTALS
L L L L L L
D D D D
P
D D D D D D
P P P
P P P P P P
P P P P
P P P P
P P P P P P
P P P P P P
17 I 7191 9 I 81 6
* Including canardst Year of first flight
P = ProductionL = Limited Production
D = Development
Table A-5. Composite components on business aircraft.
39
boat is included in the list.
The only three business aircraft that havebeen almost entirely made of advanced compositesare all American-made. All three programs havehad major structural problems that caused weightgrowth, extension of design and developmentschedules, and cost increases.
The first of these composite aircraft was theLearFan 2100, which was conceived by Bill Learshortly before his death in 1978. Almost the entireairframe was made from CFRP and AFRP. All
primary structure was sheet-stiffened construction.
The prototype first flew in 1981. The air-frame was modified following wing and fuselagefailures during structural test. The modified air-frame did satisfy structural certification require-ments and was certified by the FAA. The struc-tural modifications required to meet certification incombination with premature failure of the gearboxdelayed the program, increased development costs,and caused most of the 200 orders to be with-
drawn. The first and only production aircraftfinally flew in t983. LearFan Corporation declaredbankruptcy in 1985 because of delays and difficul-ties in FAA and British CAA certification and
other financial problems (Jane's; AWST, Jan 12,1981; Whitaker, 1981 ; Wigotsky, 1983).
The second all-composite aircraft was the
Beech 2000 Starship. The Starship configurationwas originally conceived in 1982 by Burt Rutanand went into production in 1988. The Starship hasan airframe made almost entirely from CFRP-Nomex® honeycomb sandwich construction. An85% scale proof-of-concept vehicle was flown in1983. The first of three Starship prototypes flew inFebruary 1986. Major structural modifications hadto be made to the wing to satisfy FAA damagetolerance requirements, and to the fuselage follow-ing premature failure during structural test. Thefirst production Starship, which was flown in April1989, had a take-off gross weight of 14,400 lbcompared with an original target weight of 12,500lb. The weight increase in combination withaerodynamic efficiency that was lower than ex-pected based on the performance of the scalevehicle, reduced range and performance. Becauseof limited aircraft orders, Beech decided in 1993 to
terminate production at 50 aircraft (Jane's; Abbott,
1989; Aerosp Eng, Apr 1990).
The third composite airplane was the Avtek400 light corporate transport.* Avtek Corporationhas produced one proof-of-concept Model 400 thatfirst flew in 1984. The airframe, designed by Dr.Leo Windecker, is 72% Kevlar®/epoxy and 18%carbon/epoxy by weight. Following flight andwind tunnel tests, the aircraft was redesigned andincorporated so many major changes that thecurrent Model 400A is essentially a new design.CFRP is used for the wing spar caps and webs, andthe rudder. Two ground test aircraft are being builtfor FAA certification tests to FAR Part 23 require-ments. Current investors in Avtek include duPont,the State of Alabama, the government of Malaysia
and various foreign companies. Avtek is currentlylooking for about $70-million for the flight testsand structural tests needed to complete theFAR Part 23 certification program. They haveorders for 89 airplanes (about two years of produc-tion) (Jane's).
Other US business aircraft CFRP applica-
tions include the Gulfstream IV engine supportstructure and the pressure bulkhead as well as the
ailerons, rudder, and spoilers, which were de-signed and are made by Lockheed. The elevators,rudder, ailerons, and flaps of the Cessna Citation Vare also made from composites.
The Claudius Dornier Seastar was designedin Germany by a team led by the late Dr. ClaudiusDornier (who had no connection with Dornier
GmbH). The Seastar is a light twelve-passengerSTOL amphibian, designed to operate from grass,water, snow or ice. The airframe is made almost
entirely from GFRP with some CFRP reinforcing.Design was started in January 1982 and the proto-type Seastar was flown from July 1984 until it wasdamaged landing on Lake Constance in July 1985(Air Int, Oct 1988).
The damaged aircraft, which originally hadbeen flown with an aluminum wing from a DornierDo 28 and a composite fuselage, was repaired andrebuilt with an all-composite wing. DesignatedCD2, the rebuilt aircraft flew in April 1987. CD2Seastar production started in October 1989 but thecompany went into bankruptcy one month later.
*Private communication on composites use on the AvtekModel 400A supplied to R. Hadcock by Robert Adikes,Avtek Corp., September 1993.
4O
DomierCompositeAircraft, ownedby ConradoDomier,purchasedthecompanyinFebruary1990.TheCD2 wascertificatedby LBA in October1990andmeetstheFAR Part23commuteraircraftcertificationrequirements.Thefirst productionSeastarwasdeliveredat theendof 1991(Jane's;Air Int, Oct 1988).
DassaultandA6rospatialeweresponsoredbytheFrenchgovernmentin 1980to designandbuilda CFRPwing for theDassaultFalcon10businessjet. Theintegrally stiffenedwing coversandmostof thebeamsandribs in thesubstructurewereCFRP.Themainlandinggearribs, innerrearspars,andtherootwing-to-fuselagelapjoints werealuminum.
TheItalian PiaggioP.180Avanti lightcorporateturbopropwasproducedat arateof 24airplanespermonthin 1991.Thecompleterearfuselageandempennageassemblyof theAvanti isCFRPandwasdesignedandisbeingproducedbySikorsky.PiaggiomakestheCFRPcanard.TheAvanti prototypefirst flew in September1986andthefirst productionairplaneflew in January1990.Full Italian certificationwasobtainedin October1990.
Private, Trainer, and Competition Aircraft. -
A list of composite components on private, trainer,and competion aircraft is given in Table A-6. TheWindecker AC-7, Eagle I, was the first all-com-
posite private aircraft to receive FAA certification.The Eagle was designed by Leo Windecker whoused GFRP for the entire airframe. The Eagle flewin 1967 and received FAA certification in 1969
(Rosato, 1969; Taylor, 1989). The Eagle was ahigh-performance, single engine, four-seat mono-
plane but did not go into production. LeoWindecker later assisted in the design and devel-
opment of the Avtek 400.
*Private communication, Avions de Transport R6gional,
Blagnac Cedex, France, November 1993.
The Bellanca Model 19-25 Skyrocket II, a
six-seat light monoplane, was also made almostentirely from GFRP. The Skyrocket was designedand built by Bellanca Engineering Inc., a companyformed by August Bellanca. Design and construc-
tion of the prototype started in 1971 and theairplane first flew in March 1975. Powered by a435 hp Continental engine, the Skyrocket II had acruise speed of 331 mph and a range of 1,465miles and held five FAI speed records. Bellancawas working on FAA certification in 1984, but the
program was never completed. The airplane wasused by NASA Langley for flight and wind tunneltests in 1982 (Jane's; Taylor, 1989).
For the past 25 years, Scaled Composites Inc.led by Burt Rutan, has been involved in design andfabrication of many all-composite proof-of-con-
cept and competition aircraft. These aircraft, whichare made from CFRP/foam sandwich construction
are not included in this report. They include the
Voyager, which was the first airplane to fly aroundthe World without refueling, the Pond Racer, the
NASA AD-1 oblique wing research aircraft, thescale demonstration T-46, and the Starship.
The British Slingsby T67 Firefly aerobatic,training and sporting aircraft is an all-compositeversion of the French wooden Fomier RF6B.
Slingsby has made GFRP high performancesailplanes for many years and used their sailplaneexperience to design the Firefly. The T67B gainedCAA certification in September 1984 and about200 had been sold to customers world wide by1993.
The USAF ordered 113 Slingsby T67s,designated the T-3A, to fulfill the Enhanced FlightScreener (EFS) program. T-3As are being pro-duced at a rate of 5 aircraft per month and areassembled by Northrop. FAA certification is beingobtained to avoid the need for USA airworthiness
testing. The T-3A airframe is GFRP with localreinforcement of CFRP. The structure has been
modified to satisfy the USAF +6/-3g limit require-ments carrying two 260 Ib pilots (Jane's; Penney,1993).
Sukhoi first flew the Su-26M single-seat
aerobatic competition aircraft in June 1984 and
gained both the men's and women's team prizes inthe 1986 World Aerobatic Championships. The
wing and empennage have CFRP and CFRP/AFRPskins. Composite materials comprise more than to50% of the structure weight. The fuselage is made
41
AIRCRAFT ] YEARt
WING
BOX
C I SOI UV I BE I SR I TSIR
FUSELAGE STABIL-IZERS
s I H I VH N 0 I EE T R I RL E I I TL R Z* I I
N CA A
! L LI
CONTROL SURFACES
E I R [ A " FL I U I I LEl D I L AV I D I E PA E I RT I R I OO S I N
R SS
PRIVATE, TRAINER & COMPETION AIRCRAFF
.Wind.ecker Eagle
Bellanca Skyrocket II
FFT Speed Canard
Siingsby T-3A
Sukhoi Su-26M
Rutan VoyagerAvtek 400
Egrett D-500Rutan AT3
Rutan ARES
Rutan Pond Racer
Sukhoi Su-31
FFT Eurotrainer
Grob GF 200
RUDASA Fan Ranger
1967
1975
1980
1983
1984
1984
1984
1987
1987
1990
1991
1991
1991
1992
1993
TOTALS
L L L L
L L L L
P P P P
P P P P
P P
D D D D
L L L L
L L L L
D D D D
D D D D
D D D D
P P P
P P P P
P P P P
L L L L
1151151 14[
L L L L L L
L L L L L L
P P P P P
P P P P P P
P P P P P P
D D D D D D
L L L L L L
L L L L L L
D D D D D D
D D D D D D
D D D D D D
P P P P P P
P P P P P P
P P P P P P
L L L L L L
13 I 15115114 I 15 115 I 15
RPVs & DRONES
!Ryan BQM-34E
(NADC)
B oein_, YQM-94A
Rockwell HiMAT
Ryan BQM-34F DAST
Boeing Condor
1971
1976
1980
1983
D D
D D
1991 D
ToTALs I 3 I
D D
D D
D D
D
D D
15141
P = Production* Including canards? Year of first flight L = Limited Production
D = Development
D D D D D D
D D D D D D
D D D D D D D
3131313131313
Table A-6. - Composite components on private, trainer, competition, RPV, and drone aircraft.
42
from welded stainless steel tubing with removable
composite skin panels, and the landing gear istitanium. Operating g limits are + 12/-10 and theaircraft was designed to an ultimate load factor of+23. The Sukhoi Su-29 is a larger two-seat trainer
version of the Su-26M, designed to limit loadfactors of +11/-9 solo and +9/-7 dual. The Su-31
is an all-composites higher performance follow-onto the Su-26M and first flew in 1991 (Jane's;
Smith, 1993).
The German Grob company is producingvarious all-composite airplanes made primarily ofGFRP. Grob was also a partner in the E-Systems/Grob/Garrett Egrett-1 high altitude surveillance
aircraft project, which was terminated in 1993.The company is headed by Dr. Burkhart Grob,and, like Slingsby, has had many years of experi-
ence designing and manufacturing GFRP sail-planes.
The Grob G 115 two seat light aircraft has"conventional" GFRP structure and was certifi-
cated to FAR Part 23 standards by the GermanLBA in 1987 and by the British CAA in 1988. Itnow has FAA certification. Grob is currently
developing the GF 200 all-composite four seatlight aircraft.
FFF, another German company, has been
producing all-composite light aircraft for manyyears. Their FFT Speed Canard two-seat sportingaircraft first flew in 1980 and is certificated in
many countries including the US (Jane's).
The FFT Eurotrainer is a two seat trainer.
The airframe is primarily GFRP reinforced withCFRP. The first Eurotrainer flew in 1991 and
obtained certification in 1992 (Jane's).
The last trainer aircraft listed in Table A-5 is
the Rockwell International/Deutsche Aerospace
Ranger 2000. Rockwell teamed with DA to de-velop the Ranger 2000 for JPATS. The airplane isa derivative of the German RFB Fantrainer that
was first produced in 1984. The airframe is almostentirely made from GFRP with CFRP reinforce-ment (Jane's; AW&ST, Sep 13, 1993).
REMOTELY PILOTED RESEARCHVEHICLES AND DRONES
Various remotely piloted research vehicles
(RPRVs) have been made from advanced compos-ite materials to demonstrate performance. Some of
the US RPRVs and drones are listed in Table A-6.
The Boeing YQM-94A Compass Cope wasan Air Force long range, high-altitude, unmannedreconnaissance vehicle made almost entirely of
glass/epoxy with some Kevlar®/epoxy. Thevehicle, which had a wing span of 94 It, had anendurance of 30 hours at 50,000 to 70,000 ft. Twoaircraft were made. The first one flew on July 28,
1973, but it was destroyed in a crash nine dayslater. The second aircraft completed a successfulflight test program in 1974 and is now in the AirForce Museum. The YQM-94A never went into
production (Bowers, 1989).
An NADC program included design, fabrica-tion, and flight test of CFRP wings using a RyanBQM-34E supersonic drone as the baseline ve-hicle. Five ship sets of wings were fabricated andthe first was proof-tested to 120% design limitload (DLL) for one critical 5g maneuver conditionbefore flight. The other wings were proof tested to100% DLL before flight (Manno, 1977;McQuillen, 1971).
The wings were deployed at Pacific andAtlantic ranges and flown on operational BQM-34Es starting in 1976. After 10 flights or threeyears of service, the wings were returned toNADC for dissection and small specimen testing(Manno, 1977;
The NASA HiMAT (Highly ManeuverableAircraft Technology) RPRV was designed andbuilt by Rockwell International to demonstrateimproved transonic maneuver performance usingaeroelastic tailoring. The HiMAT was a 0.44 scalemodel of a 17,000 lb fighter and was designed to alimit load factor of 12g. Almost the whole air-flame was made out of CFRP and the anisotropic
properties of the CFRP wing covers were used toprovide aeroelastic tailoring (Monaghan, 1981;DeAngelis, 1982).
The NASA DAST (Drones for Aerodynamicand Structural Testing) program utilized a RyanFirebee II BQM-34T target drone aircraft. TheCFRP wing skins were mechanically attached tometal substructure and were purposely designed to
have fiber controlled bending strength and stiff-ness but matrix-controlled torsional stiffness and
shear strength. The wing was designed to encoun-ter flutter within the flight envelope so that anactive flutter control system could be investigated(Eckstrom, 1983).
43
During the third test flight of the DAST,divergent oscillations occurred with the fluttersuppression system on. The wing failed and theaircraft crashed. The primary wing componentswere recovered and flight testing was resumed in
1982. Although active flutter suppression waseffective, the study indicated that structural tailor-
ing using matrix-dominated properties should beavoided (DeAngelis, 1982).
Boeing Advanced Systems designed a largetwin-engine robotic aircraft in 1988 under aDARPA program. Nicknamed the "Condor", theHALE (High Altitude Long Endurance) aircraft seta high altitude record for piston engine aircraft at66,980 ft in 1989. The "Condor" has a wing spanof more than 200 ft and has an all-bonded airframe
made almost entirely out of carbon/aramid/epoxyhybrid materials _owers, 1989).
HELICOPTER APPLICATIONS
Composite materials have been used for manydifferent helicopter components including rotorblades, stabilizers, and fuselage structure. Many of
the helicopters and their associated compositestructural components are listed in Table A-7.
Rotor Blades. - In 1959, the VertoI Aircraft
Corporation (previously the Piasecki HelicopterCorporation and later Boeing Vertol and BoeingHelicopters) started development of an "OptimumPitch Blade" for the XCH-47 twin-rotor helicopter.These blades were made from E-glass/epoxy andsurvived a 150 hour whirl test. This success led to
fabrication of ten CH-47 GFRP blades for static,
fatigue, and flight tests in 1964.
The CH-47 blade test program was followedby the successful completion of a Navy-fundedGFRP production blade development program. Bythe mid-1970s, GFRP blades had essentially re-
placed all 4130 steel spar blades on Boeing heli-copters. The GFRP blades have a service life of atleast 10,000 hours compared with a life of about1,000 hours for the blades with steel spars. Boeinghad made more than 10,000 GFRP blades for the
CH-46 and CH-47 by the end of 1992.
About the same time that Vertol was develop-ing GFRP blades, Messerschmitt B61kow-Blohm(MBB) developed GFRP blades for the hingeless,
semi-rigid rotor system for the Bo-105 helicopter.Initial flight tests of the rotor system were made
using a Sud-Aviation Alouette helicopter and thefirst flight of a Bo-105 was made in 1967. The Bo-105 was still being produced in 1993 (Jane's).
Boeing Vertol and MBB reached a coopera-tive agreement for Boeing to utilize the MBBsystem and GFRP blade design for their Model179, YUH-61A helicopter, which flew in Novem-ber 1974. The tail rotor also had GFRP blades. The
YUH-61A was the Boeing Vertol contender forthe DoD UTTAS (Utility Tactical TransportAircraft System) competition, which was won bySikorsky with the UH-60A in 1976 (Air Int, Aug1975).
Boeing Helicopters is responsible for thefive-blade main rotor system of the Boeing/Sikorsky RAH-66A Comanche that is scheduled tofly in 1994. Boeing is using a version of the MBBall-composite bearingless system (Jane's).
Bell Helicopters and Kaman also began todevelop alI-GFRP blades in the late 1960s toreplace metal-spar blades. Bell blades had D-shaped aluminum spars with bonded aluminum
skins; Kaman blades had aluminum spars withGFRP skins. Kaman introduced all-compositeblades on the SH-2G in 1987. These blades have a
service life of 10,000 hours. (Jane's; Rosato, 1969)
By the mid-1970s, Bell started producingGFRP blades for the AH-1 Huey Cobra, Model214, and Model 222, but retained aluminum for the
tail rotor blades. Many Bell AH-I models wereproduced or retrofitted with Kaman-designed K-747 GFRP rotor blades between 1977 and 1988.These blades also have a service life of the order
of 10,000 hours. (Jane's; Peacock, 1988)
Belt won the OH-58D, US Army HelicopterImprovement Program (AHIP), in 1981. Theseaircraft have the Bell four-blade rotor systems with
CFRP yokes, GFRP blades, and elastomericbearings. The first OH-58D flew in 1983 anddeliveries started in December 1985. The Bell
Model 406, a lighter version of the OH-58D, firstflew in 1990. The Bell Model 680 4-blade rotor
system ("Rotor 90") is almost entirely composites.It is 15% lighter and has 50% fewer parts that thecurrent system (Jane's).
During the early 1970s, Sikorsky developedmain rotor blades composed of hollow titaniumspars with GFRP/honeycomb sandwich trailing
44
AIRCRAFT YEARt
Sikorsky S-61 1959
Sikorsky CH-53 1965
MBB Bo 105 1967
Sikorsky CH-54 1971
Westland Wasp 1971
Boeing CH-46/CH-47 1974
Bell All- IF 1974
Sikorsky YUH/UH-60 1974/75
Boeing YUH-61A 1974MD YAH-64/AH-64 1975/84
Sikorsky S-76 1977
Aerospatiale AS-365 1979
Westland Sea King 1979
BK-117 1979
Aerospatiale AS 332L 1980Bell 206L 1981
Bell 412 1981
Kaman SH-2G 1981
Kamov Ka-32 1981
Mil Mi-28 1982
Agusta 129 1983
Sikorsky S-75 (ACAP) 1985
Westland Lynx 1986
Bell 292 (ACAP) 1986
Boeing 360 1987
Aerospatiale AS 565 1987
EH Industries EH 101 1987
Mil Mi-34 1987
MBB Bo 108 1988
MD MD 520N/530N 1990
Eurocopter Tiger 1992
MD MD 900 1992
Kamov Ka-62 1994
BoeinflSikorsh_' AH-66 1995
TOTALS
ROTORBLADES
M TA A1 IN L
D 1
P
P
P
D D
P
P
P P
P P
P
P
P
P
P P
P P
FUSELAGE
MS T CFI E A ARDC l NA
T L OMI B PEO O YN O
M
D 2
P
P
P
P
D 3
D
P
D
P
P
P
P P
D
P P
D D
P P
D D D D
D D D
P P P P P
P P P P
P P P
P P P P P
P none P
P P P P P
P none P P P
P P P P P
P P P P
1251 ill 15113119
STABILIZERS OTHER
1 H V S DN O E P RT R R OE I T N SR Z I S HN C O AA A N FL L S T
P P
D
P P
P
P P
P
P
p p4
P P P
D D D
p5
D D D
D
P P P
P P P
P
P P P
P P P
P P P
P P P
P P P
[ 15 115 [ 14
D
P
P
D
D
D D
P
P P
11ol 3
t Year of first flight or completion of R&D pro_am
Westland/Agustapartnership.TheEH-101,a largemilitary andcommercialgeneralpurposehelicop-ter,wasfirst flown in 1987.Productionaircraftarecurrentlybeingdeliveredto British andItalianforces(Jane's).
In France,SudAviation (laterAerospatialeandnow partof Eurocopter)first introducedGFRPmainrotorbladeson theSA 341Gazelle,whichwasfirst flown in 1967.Tail rotor bladeswerealuminum.ThebladeshadaGFRPsparandskinssupportedby afoamcore.ThecurrentversionoftheGazelle,theSA 342,wasstill beingproducedin 1992(Jane's).
By 1990,AerospatialewasusingGFRPbladeswith a CFRPhubfor themainrotor systemandCFRPbladesfor theductedfan of theAS 365Dauphin2 helicopter(Jane's).
During thepastfew yearsMcDonnellDou-glashasbeendevelopingalow-noisefive-bladerotorsystemwith carbon/epoxyblades.Thecompany-fundedHARP (HelicopterAdvancedRotorProgram)flexbeamCFRProtorwasfirstflown onanMD 500Ehelicopter.Thisrotor isusedin combinationwith theNOTAR (NoTailRotor)systemfor theMD 520NandMD 900transport/utilityhelicoptersthatfirst flew in 1990and 1992,respectively.TheMcDonnellDouglasNOTAR systemwasfirst flown ontheMD 530Nin 1989andthefirst productionMD 520Nwasdeliveredin October 1991. With the NOTAR
system, the MD 520N and MD 900 are 50%quieter than comparable helicopters (Jane's;Proctor, 1993).
MBB teamed with Kawasaki to design andproduce the BK 117 multipurpose helicopter,which made its maiden flight in June 1979. TheBK 117 main rotor system is similar to the MBBBo 105 system with hingeless GFRP blades. The
BK 117A was certified to FAR Part 29 require-ments in Germany, Japan, and the US in December1982, and the BK 117B model was certified by theLBA, JCAB, and the FAA in 1987/1988. The
BK 117 is being produced in Germany, Japan,Canada, and Indonesia (Jane's; Air Int, Apt 1989).
Some of the other helicopters that utilizeGFRP or mixed CFRP/GFRP rotor blades are the
Franco-German Eurocopter (DA/MBB withAerospatiale) Bo 108 and the PAH-2 Tiger.
Westland started production of the Lynx inthe UK in 1972 under a cooperative agreementwith Sud Aviation. The Lynx had rotor blades withtitanium spars, based on the Sikorsky design(Westland was licensed to produce variousSikorsky helicopter models). The Sea King Mk 2,British derivative of the Sikorsky S-61D0 and the
Lynx AH Mk 9, an upgraded version of the Lynx,were the first Westland helicopters to use GFRPmain and tail rotor blades. The Lynx compositeblades, developed under the British ExperimentalRotor Programme (BERP), were first flown in1986, and the Lynx established a world helicopterspeed record of 249 mph (Jane's; Gething, 1990).
The blades of the EH Industries EH- 101 are
scaled-up versions of the BERP Lynx blades,designed and built by Westland. EH Industries is a
The Italian Agusta A 109 and A 129 helicop-ter blades have AFRP spars with GFRP skins,Nomex® honeycomb core, and stainless steelleading edge abrasion strips (Jane's).
All-composite blades were first produced inRussia for Mil Mi-28 combat helicopter, whichfirst flew in I982. The blades are made from a mix
of CFRP, GFRP, and AFRP composites with aNomex®-type core. All the latest Russian helicop-ters, which include the Kamov Ka-32 and Ka-62,
and the Mi-34 and Mi-38, have all-compositeblades (Jane' s; Fricker, 1990).
Helicopter Airframes. - During the past 35years, use of composite materials in helicopterstructures has grown from a few small accesspanels and canopy frames to almost all of the
46
airframe.Compositeshaveprovidedweightsavings,which is particularlyimportantto helicop-terperformance,aswell asimproveddurabilityandcorrosionresistance,andreduceddrag.
In 1971,SikorskycompletedaNASA pro-gramthatusedboron/epoxyto reinforcealuminumcomponentson therearfuselageof theCH-54BSkycrane.Staticandfatiguetestssatisfiedstrengthandlife requirementsandthereinforcedstructurewas130poundslighter thanthealuminumbaseline.A reinforcedrearfuselagewasput intoflight serviceonanArmy helicopterin April 1972.Thehelicopterwastakenout of servicein after itwasseverelydamagedin a wind storm.Thetailboomwasnotdamagedandthehelicopterwassubsequentlyrepairedandput backintoservicewith theNationalGuard.Theboron-reinforcedstructureapproachhasnotbeenusedfor anysubsequentSikorskyhelicopters(Rich, 1972).
Third generationcompositestructures,whichuseamix of CFRP,GFRP,andAFRPfor almostall of theairframe,werefirst demonstratedby the
US Army SikorskYS-75AdvancedCompositeAircraft ProgramI(ACAP)andtheBelt Model 292ACAP in 1985/1986.BoeingseparatelydevelopedtheModel360all-compositetwin rotorhelicopter,whichflew in 1987.Thesehelicoptersweretech-nologydemonstratorsanddid not go intoproduc-tion, but their developmentled to extensivepro-poseduseof compositesoncontendersfor theArmy Light AttackHelicopter(LHX). TheLHXcontractwaswon by theBoeing/Sikorskyteam'sRAH-66Comanchein 1991.The first RAH-66isscheduledto fly inAugust1995.TheRAH-66fuselageis beingmadeprimarily of carbon/epoxyandaramid/epoxyandhasabout350partscom-paredto about6,000partsfor theUH-60 fuselage(Jane's; Parker, 1993).
Other helicopters that have largely compos-ites airframes include McDonnell Douglas MD900 and the German MBB Bo 108, both of which
are in production, as well as the Eurocopter Tigerand the Russian Kamov Ka-62. The compositefuselage of the Bo 108 has almost 30% less dragthan the Bo 105 aluminum fuselage (Jane's).
The Tiger is an attack helicopter that is beingdeveloped by Eurocopter (the DA/Aerospatialeconsortium) for the German and French armies.Deliveries are scheduled to begin in 1997 (Jane's;Mordoff, 1988).
The Kamov Ka-62 is a multi-purpose heli-copter. Composites account for about 50% of theairframe weight and include the main cabin shell,floors, tailboom, stabilizers, fan duct, and fanblades (Jane' s).
The Italian Agusta A 129 Mangusta attackhelicopter airframe is a mix of composites andaluminum. The A 129 first flew in 1983 and
entered service with the Italian Army Aviation in
1990. Some 900 lb of composite materials are usedfor the nosecone, canopy frame, tailboom, tailrotor pylon, engine nacelles, and the stabilizers,accounting for about 45% of the airframe weight(Jane's).
EH Industries was formed in 1980 by
Westland Helicopters and Augusta. The EH 101multi-role helicopter first flew in Italy 1987.British and Italian civil certification was expected
in 1993. Military variants, scheduled to enterservice in 1995, include naval and land-based
helicopters for the British, Italian, and Canadianforces. Composites are used for the canopy frame,
Compositematerials,primarily GFRPorAFRPwith CFRPreinforcement,havebecomethestandardmaterialsfor helicopterstabilizers,enginedoors,cowlings,fairings,doors,landinggeardoors,floor panels,stubwings,sponsons,andfanducts.McDonnellDouglasmadesomeAH-64stabilizersfrom carbon/thermoplastics,but theywerenot flown or put intoproduction(Colucci,1991).
OBSERVATIONS AND CONCLUSIONS
The following observations and conclusionsare drawn from a review of the international
aircraft programs described in this Appendix aswell as from the interviews with industry andgovernment personnel.
Advanced composites are being used exten-sively for primary and secondary structures ofmany new US and foreign military and commer-cial aircraft. The technical risks involved with the
use of composites appear no greater than thoseassociated with metals.
Overall weight savings have been achievedby using composites instead of metals. Componentweight savings can be as high as 35%. Typically,composites make up between 22% and 35% of theairframe by weight for new US and foreign mili-tary tactical aircraft. The composite horizontalstabilizer on the Airbus A320 is 15% lighter than
its aluminum counterpart and the ATR 72 outerwing saves 20% (Barrio Cardaba, 1990; Pilling,1988).
Weight savings were the major considerationwhen Boeing decided to introduce advancedcomposites on the B757 and B767 in the late1970s. The price of jet fuel had increased from$0.12/gal in 1973 to $1.04/gal in 1981 (AerospaceFacts & Figures; Bowers 1989).By 1991, the priceof jet fuel had dropped to $0.69/gal, but the pricesof commercial transport aircraft have increased by
almost 500 percent since 1973. The cost of fueldropped from 30% of cash operating expenses in1981 to 14.8% in 1991. (Aerospace Facts &Figures) In today's business environment, weightsavings are not marketable unless they can beaccomplished at no additional cost.
Until the mid 1980s, Boeing was the only
company utilizing advanced composite materialsfor spoilers, elevators, rudders, and flaps of com-mercial transport aircraft (B757, B767, B737-300).
In 1985, Airbus moved ahead by adding the A310composite vertical stabilizer and, in 1987, theyadded both the horizontal and vertical stabilizers
of the A320 to their list of applications. The A330and A340 also have composite stabilizers. As ofDecember 1991,247 A320s were in airline ser-vice.
The European aircraft community movedfurther ahead when ATR introduced advanced
composite outer wings on the ATR 72 in 1988 andDeutsche Airbus selected CFRP for the rear
fuselage and pressure bulkhead of the Dornier Do328 in 1991. As of December 1991, 48 ATR 72swere in airline service.
Boeing is adding horizontal and verticalstabilizers and cabin floor beams to the list of
composite components on the B777, which is dueto fly next summer. By that time, Airbus will havean advantage of at least eight years of productionand service experience of advanced compositestabilizers. On the other hand, Boeing may use ofless advanced composites on their derivative 737(737-X) than the 737-300 because smaller airlines,
who are the potential customers for the 737-X, donot have the composite maintenance and repaircapabilities of larger carriers.
McDonnell Douglas is using advancedcomposites for most of the control surfaces of the
MD-11. The aileron is a post buckling designderived from the DC-10 rudders developed underthe NASA ACEE program.
The NASA Aircraft Energy EfficiencyProgram (ACEE) Primary Aircraft StructuresProgram, which ran for about 15 years from 1972until 1987, was very successful in demonstratingthe technology readiness and cost effectiveness ofcomposite structures for commercial transports(NAA CP-2321, 1984). In retrospect, the NASAflight service experience programs and the ACEEprogram had enormous influence on the accep-tance of advanced composite structures by industryand the aircraft operators in the US and abroad.Many of the components developed under theseprograms (e.g., the DC-10 vertical stabilizer andrudders and the Boeing 737 horizontal stabilizersand spoilers) are still in scheduled airline service
Theairlines,however,areveryconcernedaboutthin-skinhoneycombsandwichcompositesecondarystructures.Thesepartsoftengetdam-agedin service(generallyduringaircraftmainte-nance)andthecostsof repair,replacement,orleasingsparepartsareveryhigh. As notedabove,Boeingmaychangeanumberof secondarycom-positecomponentsbackto metal onthenewB737-X.
Costsareuniversallyrecognizedasthebiggestproblemassociatedwith compositesuseinplaceof conventionalmetalstructures.Manyofthemilitary aircraftprogramshavehadconsider-ablecostoverruns.Somecostincreasesaredi-rectlyassociated with composite structures andsome with inexperience in the design and manu-facture of composites. The costs of compositesmaintenance, repair, and replacement parts add tothe overall cost problem. Regardless of type ofmaterial, the prices of military aircraft structuresare much higher than those of comparable com-mercial and business aircraft structures. Based
upon current program cost estimates, the fly-awayprices of new military aircraft structures rangefrom $1,300/1b for the McDonnell Douglas C-17and $1,500/lb for the Lockheed F-22 to $4,000 for
the Northrop B-2B. Prices will increase further ifproduction rates and quantities are reduced. Incontrast, the current prices of commercial turbojetand turboprop transport aircraft range from $200/tb to $300/lb (Aerospace Facts & Figures;Hadcock, 1985, 1989; McCarty, 1991).
As was the case with military aircraft, costs
are also the biggest problem associated withcomposites use in place of conventional metalstructures on commercial transports. Costs includeproduction costs, as reflected in airplane prices, aswell as in-service costs associated with component
inspection, maintenance, repair, and replacement.
Nearly all the people interviewed thoughtthat the prices of composite components are higherthan their metal counterparts, and that, to bemarketable, their prices should be comparable. The
prices of current commercial transport airframes,
based on aircraft prices, have remained relativelyuniform at $200/lb to $300/lb (Whitehead, 1993;
McCarty, 1991).
Allowing for a 20% weight saving, prices ofinstalled composite structures should be competi-tive with metal structures in the $250/1b to $350/Ib
range. Component prices should be in the $200/1bto $300/lb range to allow for final assembly costs.
Many composite parts are supplied bycoproducers or subcontractors under fixed pricecontracts. Since the price is rarely broken downinto individual elements, the individual cost of a
unit or ship set of composite components is impos-sible to obtain.
As an example, the Japan Aircraft Develop-ment Corp. (JADC) is contributing 21% to theBoeing 777 project for design and production ofthe fuselage, center wing, and wing-body fairingsfor the life of the 777 program. Grumman has a10-year, $400 million contract to produce the 777composite inboard flaps and spoilers, andRockwell, CASA, HDH, and Alenia have con-
tracts to produce other composite components(O'Lone, 1991).
Since the end of the Cold War and the cut-
back in defense spending, there will be fewer newmilitary aircraft opportunities and the gestationtime period will probably be longer than ten years.US companies (McDonnell Douglas, Boeing,Northrop, Lockheed, Grumman, and Vought) arestill the world leaders in composites technologyand production experience for high performancemilitary aircraft. However, retention and transferof technical information and experience will be amajor problem in the future. Much more reliancewill probably have to be placed on use of compos-ite technology and materials developed for com-mercial aircraft.
Other than the Boeing 777 and 737-X,development of any new US commercial transportaircraft is unlikely during the rest of the century.The development of the Boeing 777 is almostcomplete, and it is improbable that any materialchanges will be made at this stage of the program.Although composites could be considered onderivative aircraft, Boeing has chosen to use fewercomposites on the 737-X than they presently haveon the 737-300 because of the problems thesmaller airlines have with repair and maintenanceof composites.
49
Realistically,thenextmajorUSopportunityfor extensivecompositesusewill be theHighSpeedCivil Transport(HSCT) aircraft,whichdesperatelyneedstheweight savingsprovidedbyhightemperaturecomposites.TheHSCTwillprobablynot fly before2005(Whitehead,1993;Blankenship,1991).
With afew notableexceptions,useof com-positesongeneralaviation(GA) airplaneshasbeenveryconservativeandhasbeenlimited tocontrolsurfaces,flaps, fairings, landinggeardoors,andenginenacelles.Certificationfor thesecomponentshasbeenrelativelystraightforwardandhasnotrequiredstructuraltestingof thecompleteairframe.Generally,in-serviceperfor-mancehasbeentroublefree.
Fivedifferentall-compositeGA airplaneshavebeendesignedandbuilt in theUSduring thepasttwentyyears.None has been an unqualifiedsuccess. Private aircraft include the Windecker
Eagle and the Bellanca Model 25 Skyrocket II.Business aircraft include the LearFan 2100, theBeech Model 2000 Starship, and the Avtek 400A.
FAA airframe certification was granted to
the Windecker Eagle, LearFan, and Starship afterlengthy and expensive test programs. The BellancaSkyrocket was never certificated and the certifica-tion program for the Avtek 400, which first flew in1984, has yet to be completed.
LearFan went out of business in 1985 follow-
ing a number of program delays and certification
problems. Two flight-test and one structural-testaircraft were built (Jane's). Windecker Researchceased operations in 1976 after completion of 8aircraft (Jane's; Simpson, 1991).
The all-composite (primarily GFRP) BritishSlingsby T.67 Firefly (USAF T-3A) and theGerman Grob G-115 civil/military trainers appearto have avoided the financial problems of theEagle and Skyrocket. Both aircraft obtained civilcertification in the mid-1980's and combined civil
and military sales have been about 320 for theFirefly and 100 for the G-115. Prior to the intro-duction of these trainer aircraft, Slingsby and Grobhad extensive design and manufacturing experi-ence producing high-performance GFRP sail-planes.
,_vtek Corporation is building two groundtest aircraft for FAA certification tests to FAR Part
23 requirements, but will require about $70-million to complete the flight tests and structuraltests needed for the FAR Part 23 certification
program.*
Beech Aircraft Corporation first flew theprototype Starship in 1986 following the flight-testprogram on an 85% scale proof-of-concept (POC)aircraft. The full-scale aircraft did not have the
performance projected from the POC program.Between 1984 and 1990, structure, empty, andtakeoff weights increased by 4%, 29%, and 19%,respectively. These weight increases reducedcruise speed, fuel efficiency, and range.
The airframe of the Starship, which is 67percent composites by weight, has been blamed forthe poor performance of the aircraft. It is approxi-mately the same price as the Piaggio Avanti andBeechjet ($4-million), but is slower and has ahigher approach speed than the other aircraft. Thepoor performance appears to have been caused bythe aerodynamics of the unconventional configura-tion of the aircraft and by systems weight growth,and not by the weight of the composite airframe,which is only 27 per cent of takeoff gross weight(Abbott, 1986; Aerospace Eng., 1990).
In retrospect, the late 1970s and early 1980swere inopportune times for Windecker, Bellanca,LearFan, and Avtek to enter the GA marketplace.Their composite aircraft were introduced during atime when sales of GA aircraft were on the de-
cline. In the case of US single-engine pistonaircraft, sales dropped from 14,400 aircraft in 1978to 1811 aircraft in 1983 and only 564 aircraft in1991. In the case of twin-engine turboprop aircraft,sales dropped from 918 airplanes in 1981 to 222aircraft in 1991 (Aerospace Facts & Figures). Thenewly formed companies were trying to enter thisdeclining market with new aircraft designs madefrom unconventional materials. Their competitors(Beech, Cessna, Piper) had extensive aluminumairplane design and production experience, world-
class reputations, and world-wide sales and sup-port organizations.
The huge costs of full-scale engineering
development and associated wind tunnel, struc-tural, and flight testing, as well as the delays andmodifications associated with certification of new
*Private communication from Robert Adikes, Avtek Corp.,September 1993.
5O
aircrafttypes,werenot fully anticipated.Thesedelayscausedlostordersandcompoundedfinan-cialproblems.Noneof thenew companieshadanestablishedproductline thatcouldcarrythemthroughthelengthydevelopmentprograms.Theirfuturedependedon timelydeliveryof their prom-isedproduct.As customerconfidencefaded,sodidtheirfinancialsupport.
to fulfill the composite promise of 20-30 per cent,surfaces must be mirror- smooth and yield laminarflow, attention to producibility and economy muststart when the CAD-CAM CRT is first switched on.Absence of any of these critical points of concen-tration .... will result in a product that will prob-ably be a market bummer even if it is fortunate to
get past the prototype stage.' (Blue, 1985).
The use of composite (primarily GFRP) rotorblades on helicopters has raised blade operationallife from between 1,000 and 3,000 hours to 10,000
or more flight hours. Composite blades can bedesigned to be "fail-soft" and do not require asfrequent inspection for cracks as do metal blades.In addition, blade and rotor system efficiencies
have been improved because of the tailorability ofcomposites. Composite blades are generally nolighter and are more expensive than their metal-spar counterparts. However, their longer life andreduced in-service inspection requirements makethem very attractive and cost-effective for bothmilitary and civil helicopters.
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The authors wish to acknowledge the coop-eration and assistance of individuals from the
following companies and organizations. Theirwillingness to share their experiences and criticalobservations was essential to the preparation of
this report.
Beech Aircraft Corporation
Northrop Aircraft (now Northrop Grumman)
Rockwell International, North American AircraftDivision
United Technologies Sikorsky Aircraft
US Army Aviation Research & TechnologyDirectorate
US AF Wright Aeronautical Laboratories
Vought Aircraft Company (now Northrop Grum-man)
Bell Helicopter Textron
Boeing Commercial Airplane Group
Warner Robins Air Logistics Center, Robins AFB,GA
54
Form Approved
REPORT DOCUMENTATION PAGE oMe No o7o4-o188
Puol_c reoo_,n_ Duroen for th_$coIlec'_on of _nformat_on ,%._%_mat_ ",3average _ hour per re$_r'se, JncludJn_ the time for reviewing instru_=on$. _arcmng exlstlmg aa_a source'5gatherlno and ma_nta_nlng the data needed, and co_letl_ and rev,ewrnc the _liec_lon of _rfor_atIon _?nd comments regardln 9 this burden estimate or any other as_ Of t_co#fecC_onof inforcoatlOn, including Sugge_tfOnr,fer redu¢lnc th_ burden _c '¢Vashln_On ,-ieacloua_ers Service, Dite_rate fcr EnfOr_atlOn O_e:a_lo_ and ReDcrl$. t215 Je_erso_Dav_s Highway. Su,te 1204. Arlington, V¢* 22202-_302. and t'o the O_ice o_ Management and 3uclge'.. Pal_erwork Rc_luCl_onProle_ (07C4-_ 1_) Wa_hkneton, DC 20503
1. AGENCY USE ONLY (Leave blank) 2. P_-_ORT DATE 3. REPORT TYPE AND DATES COVERED
November 1994 Contractor Re)ort
4. TITLE AND SUBTITLE S. FUNDING NUMBERS
Composite Chronicles: A Study of the Lessons Learned in
the Development, Production, and Service of CompositeStructures
6. AUTHOR(S)
Louis F. Vosteen and Richard N. Hadcock
!7. PERFORMING ORGANIZATION NAME{S) AND ADDRESS(ES)
Vosteen: Analytical Services & Materials, Inc., Hampton, VA, 23666.
Hadcock: RN-H Associates, 6 Sue Circle, Huntington, NY 11743.
Langley Technical Monitor: John G. Davis, Jr. Final Report
12a. DISTRIBUTION/AVAILABILITY STATEMENT
Unclassified Unlimited
Subject Category 24
12b. DISTRIBUTION CODE
13. ABSTRACT (Maximum 200 words)
A study of past composite aircraft structures programs was conducted to determine the lessons learnedduring the programs. The study focused on finding major underlying principles and practices that expe-rience showed have significant effects on the development process and should be recognized and under-stood by those responsible for using of composites. Published information on programs was reviewedand interviews were conducted with personnel associated with current and past major development pro-grams. In all, interviews were conducted with about 56 people representing 32 organizations. Most ofthe people interviewed have been involved in the engineering and manufacturing development of com-posites for the past 20 to 25 years. Although composites technology has made great advances over thepast 30 years, the effective application of composites to aircraft is still a complex problem that requiresexperienced personnel with special knowledge. All disciplines involved in the development processmust work together in real time to minimize risk and assure total product quality and performance at ac-ceptable costs. The most successful programs have made effective use of integrated, collocated, concur-rent engineering teams, and most often used well-planned, systematic, development efforts wherein thedesign and manufacturing processes are validated in a step-by-step or "building block" approach. Suchapproachs reduces program risk and are cost effective.
