April 2001

NASA/TM-2001-210844

An Assessment of the State-of-the-Art in

the Design and Manufacturing of Large

Composite Structures for Aerospace

Vehicles

Charles E. Harris, James H. Starnes, Jr., and Mark J. Shuart

Langley Research Center, Hampton, Virginia
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April 2001

NASA/TM-2001-210844

An Assessment of the State-of-the-Art in

the Design and Manufacturing of Large

Composite Structures for Aerospace

Vehicles

Charles E. Harris, James H. Starnes, Jr., and Mark J. Shuart

Langley Research Center, Hampton, Virginia
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1

An Assessment of the State-of-the-Art in the Design and

Manufacturing of Large Composite Structures for Aerospace

Vehicles

by

Charles E. Harris

James H. Starnes, Jr.

and

Mark J. Shuart

Structures and Materials Competency

Langley Research Center

National Aeronautics and Space Administration

Hampton, VA 23681 USA

Abstract

The results of an assessment of the state-of-the-art in the design and manufacturing of

large composite structures are described. The focus of the assessment is on the use of polymeric

matrix composite materials for large airframe structural components, such as those in

commercial and military aircraft and space transportation vehicles. Applications of composite

materials for large commercial transport aircraft, general aviation aircraft, rotorcraft, military

aircraft, and unmanned rocket launch vehicles are reviewed. The results of the assessment of the

state-of-the-art include a summary of lessons learned, examples of current practice, and an

assessment of advanced technologies under development. The results of the assessment

conclude with an evaluation of the future technology challenges associated with applications of

composite materials to the primary structures of commercial transport aircraft and advanced

space transportation vehicles.

Key Words

composite materials, graphite/epoxy, graphite, carbon, epoxy, thermosets, thermoplastics,

aircraft, rotorcraft, general aviation, rockets, wing, fuselage, empennage, NDE, building-block

approach, analysis.
2

Introduction

An assessment of the design and manufacturing practices for large composite structures

has been conducted to determine the current state-of-the-art for these technologies. The

background that motivated the assessment was a series of unexpected manufacturing and design

problems with the composite structures of several NASA experimental vehicles currently under

development. The focus of the assessment is on the use of polymeric matrix composite materials

for large airframe structural components such as those in commercial and military aircraft and

space transportation vehicles. The baseline for the assessment is the historical evolution of the

use of composite materials in actual aerospace vehicles. The assessment emphasizes the

application of composite structures in moderately to heavily loaded aerospace vehicles.

Applications of composite materials are reviewed for large commercial transport aircraft, general

aviation aircraft, rotorcraft, military fighter aircraft, and military transport aircraft. The baseline

also includes the application of composite materials for unmanned rockets and space

transportation vehicles. The assessment of the state-of-the-art includes a summary of lessons

learned, examples of current practice, and an assessment of advanced technologies under

development. The assessment concludes with an evaluation of the future technology challenges

associated with applications of composite materials to the primary structure of commercial

aircraft and advanced space transportation vehicles.

As a preamble to assessing the state-of-the-art in the design and manufacturing of

composite structures, the design requirements for aerospace vehicles are briefly reviewed.

Because of the universal design requirement to minimize the gross take-off weight of all

aerospace vehicles, aerospace structural components are designed at or near zero margin of

safety. While the margin of safety is not equal to zero for all the design criteria at each structural

location, there is typically one criterion for each structural element that governs the design

details of that element. The quest for the lowest weight structure then drives the design margin

to nearly zero for the design limit load condition. (The Code of Federal Regulations [1] for

Aeronautics and Space, Title 14, specifies that the structure shall undergo no permanent

deformation at the design limit load (DLL). In addition, the structure shall sustain the design

ultimate load (DUL) for at least 3 seconds before failing. The factor of safety between DLL and

DUL is 1.5. Since most structural materials exhibit plasticity for metallic materials or

microcracking for composite materials prior to structural failure, the factor of safety is mostly the

difference between repeatable, linear, elastic behavior and structural failure. In other words, the

1.5 factor of safety will not provide a positive margin against unanticipated permanent

deformation or damage to the structure. Therefore, aerospace structural designs will not

accommodate any deleterious structural behavior.

In some cases, names of products commonly used in the public domain may be used

herein. Any use of these company trademarks, trade names, or product names does not indicate

a NASA endorsement of those products.
3

Part I. Historical Development of Structural Composite Materials

Large Transport Commercial Aircraft

The first composite components on commercial transport aircraft were designed and built

as part of the NASA Aircraft Energy Efficiency (ACEE) Program and entered into flight service

during 1972-1986 [2]. The primary objectives of the ACEE Program were to obtain actual flight

experiences with composite components and to compare the long-term durability of flight

components to data obtained from an environmental-exposure ground-test program. Boeing

Commercial Airplane Company, Douglas Aircraft Company, and Lockheed Corporation agreed

to participate in the program. A common feature of all three programs was the use of the

Narmco T300/5208 graphite/epoxy material system. The T300 fiber is an intermediate modulus

and intermediate strain-to-failure graphite fiber and the 5208 matrix is a thermoset epoxy that

cures at 350 °F (177 °C) . In the early years of the ACEE Program, smaller components of

lightly loaded secondary structure were designed and entered into service. These components

included the Lockheed L-1011 fairing panels, the Boeing B-737 spoiler, the Douglas DC-10 aft

pylon skin and the Douglas DC-10 upper aft rudder. In the later years of the program, larger,

more heavily loaded control surfaces and empennage structures were designed and entered into

service. Some examples of these components included the Boeing B-727 elevator, Figure 1a, the

Boeing B-737 horizontal stabilizer, Figure 1b, the Douglas DC-10 vertical stabilizer, Figure 1c,

and the Lockheed L-1011 aileron, Figure 1d. A Lockheed L-1011 vertical stabilizer was also

developed during the ACEE Program. All three of the major flight components had premature

failures before they were re-designed and successfully tested. These premature failures were

related to an incomplete understanding of the differences in the failure characteristics of metallic

and composite structures at the time that these structures were designed. By January 1987, 350

composite components had entered into commercial airline flight service.

As of 1993, the 350 components originally placed in service had accumulated over 5.3

million flight hours. The service performance, maintenance characteristics, and residual strength

of numerous components were reported to NASA and compared to the data obtained from the

10-year, environmental-exposure ground-test program [3]. The data indicated an excellent inservice performance of the composite components during the 15-year evaluation period. The

airlines reported damage such as ground handling accidents, foreign object impact damage, and

lightning strikes. However, there was no degradation of the residual strength of the composite

components due to fatigue or in-service environmental exposure. Furthermore, there was good

correlation between the results of the ground-test program and the structural performance of the

actual aircraft components.

A comparison of the applications of composite materials as a percentage of structural

weight for large commercial transport aircraft is given in Figure 2. These data were obtained

from several issues of Jane’s All The World’s Aircraft [4]. The plotted data show an increasing

use of composite materials over the past three decades from lightly loaded secondary structure,

to control surfaces, to more heavily loaded primary structure in the empennage of the Airbus

aircraft (denoted by A3XX in the figure) and the Boeing B-777. The applications of composite

materials in these aircraft are described in more detail in the next paragraphs. The current

barriers to significant increases in the use of composite structures in primary structure are the
4

higher cost of composite structures relative to conventional aluminum structures, and the

unreliability in the estimates of the design and development costs of composite structures.

Airbus was the first manufacturer to make extensive use of composite structures [4] on

large transport commercial aircraft, see Figure 2. The Airbus A310 was the first production

aircraft to have a composite fin torque box. Composite components on the A310 include the

wing leading-edge lower access panels and outer deflector doors, nose wheel doors, main

landing gear fairing doors, engine cowling panels, elevators and fin torque box, fin leading and

trailing edges, flap track fairings, flap access doors, rear and forward wing/body fairings, pylon

fairings, nose radome, cooling air inlet fairings and tail leading edges, wing leading-edge top

panels, panel aft rear spar, upper surface skin panels above the main wheel bay, glide slope

antenna cover, and rudder. The A320 was the first aircraft to go into production with an allcomposite empennage. Also, about 13% of the weight of the wing on the A340 is composite

materials. The fabrication responsibilities of the Airbus Consortium partners are as follows:

Aerospatiale fabricates the cockpit, engine pylons and part of center-fuselage; British Aerospace

fabricates the wings; Daimler-Benz Aerospace Airbus fabricates the most of the fuselage, fin,

and interior; and CASA fabricates the empennage.

The Boeing B-777 makes extensive use of composite materials for primary structure in

the empennage, most control surfaces, engine cowlings, and the fuselage floor beams. These

components are shown schematically in Figure 3. About 10% of the structural weight is

composite materials [4]. As the schematic shows, several different composite material systems

were used. Graphite/epoxy composite materials were used for most secondary structure and

control surfaces. A toughened epoxy material system, Toray T800H/3900-2, was used for the

larger, more heavily loaded components including the vertical fin torque box and horizontal

stabilizer torque box components of the empennage.

Rotorcraft and General Aviation Aircraft

Rotorcraft and general aviation aircraft have made extensive use of composite materials

to achieve performance goals. The applications of composite materials as a percent of structural

weight are plotted in Figure 4 for selected rotorcraft and general aviation aircraft to contrast the

higher percent of composite materials in these aircraft relative to the large transport aircraft [4].

The V-22 tiltrotor aircraft designed by Bell and Boeing has a number of significant applications

of composite materials. Bell and Boeing used an integrated product team approach to designing

the V-22 airframe [4]. The approach is credited with saving about 13% of the structural weight,

reducing costs by 22%, and reducing part count by about 35%. Approximately 41% of the

airframe of the V-22, shown in Figure 5, is composite materials. The wing is IM-6/3501-6

graphite/epoxy material and the fuselage and tail are AS4/3501-6 graphite/epoxy material. The

nacelle cowlings and pylon supports are graphite/epoxy material. The main cabin has composite

floor panels and the crew seats are boron carbide/polyethylene material. The fuselage is a hybrid

structure with mainly aluminum frames and composite skins. The wing box is a high-strength,

high-stiffness torque box made from one-piece upper and lower skins with molded ribs and

bonded stringers, two-segment graphite/epoxy single-slotted flaperons with titanium fittings, and

a three-segment detachable leading-edge made of an aluminum alloy with Nomex honeycomb
5

core. The rotor also used significant amounts of graphite/epoxy (17%) and glass/epoxy (20%)

composite materials.

Military Aircraft

Military aircraft have been designed with significant applications of composite materials

in primary structure. While not all information on military aircraft is publicly available, the data

in Figure 6, obtained from reference 4, compare the application of composite materials as a

percent of structural weight for a number of fighter aircraft. For example, the Lockheed Martin

F-22 Raptor, shown in Figure 7, is approximately 39% titanium, 16% aluminum, 6% steel, 24%

thermoset composite materials, 1% thermoplastic composite materials, and 14% other material

systems [4]. The fuselage is a combination of titanium, aluminum, and composite materials.

The wing skins are made of monolithic graphite/bismaleimide materials. A view of the wings

being assembled is shown in Figure 8. The wing front spars are made of titanium and the

intermediate spars are made of a graphite/epoxy material. The horizontal stabilizer uses

graphite/bismaleimide skins with an aluminum honeycomb core. The vertical stabilizers use

graphite/bismaleimide skins over graphite/epoxy spars. The wing control surfaces are a

combination of co-cured composite skins and non-metallic honeycomb core.

The Northrop Grumman B-2, shown in Figure 9, is constructed of almost all composite

materials [4]. Development of the B-2 began in the late 1970’s. The first flight test of the B-2

was July 17, 1989. The wing is almost as large as the Boeing B-747 with a span of 172 ft (52.4

m) and surface area of 5,140 ft2

 (477 m2 ).The wing is mostly graphite/epoxy material with

honeycomb skins and internal structure. The fuselage also makes extensive use of composite

materials. The outer skin is constructed of materials and coatings that are designed to reduce

radar reflection and heat radiation. Boeing Military Airplanes produced the wings and aft

section of the fuselage. Northrop Grumman produced the forward center-sections including the

cockpit. Boeing completed the outboard wing section of the twenty-first and final aircraft on

May 3, 1994.

The original design of the McDonnell Douglas (now Boeing) C-17, shown in Figure 10,

uses about 8% composite materials, mostly in secondary structure and control surfaces. In 1994,

McDonnell Douglas proposed to re-design the horizontal tail using composite materials [4]. The

tail was redesigned using AS-4 fiber in an epoxy resin for a 20% weight savings, 90% part

reduction, 80% fastener reduction, and a projected 50% acquisition cost reduction. The

prototype composite horizontal tail was successfully tested in 1998 to 133% of the design

ultimate load. Orders have now been placed for 70 aircraft with the new composite horizontal

tail.

Unmanned Rocket Experimental Aircraft

The USAF DC-X and the subsequent NASA DC-XA experimental flight vehicles [5,6]

were developed to demonstrate vertical take-off and vertical landing (VTVL), aircraft-like

turnaround times between flights, and advanced technologies that will be required for a singlestage-to-orbit reusable launch vehicle (RLV). The DC-XA vehicle, shown in Figure 11, is about
6

a 1/4 of the scale of the size of an RLV. The DC-X demonstrated autonomous VTVL flight

operations by flying eight successful experimental flights, and also demonstrated operability and

supportability of a complex, liquid fuel RLV by a small crew. Under a cooperative agreement

with the USAF, NASA took over the DC-X program and created the DC-XA by implementing a

number of important advanced technologies that will be enabling for an RLV. The advanced

technologies implemented on the DC-XA included a composite shell intertank, a composite

liquid hydrogen tank, an aluminum-lithium liquid oxygen tank, a liquid-hydrogen to gaseoushydrogen conversion auxiliary propulsion system, enhanced avionics, and an in-situ health

monitoring system. (The composite components will be discussed in more detail in the

following paragraphs.) The DC-XA program demonstrated the use of rapid prototyping to

design and fabricate the advanced technology components. The DC-XA successfully flew four

flight tests and demonstrated the viability of significantly lighter-weight structural components,

the auxiliary propulsion system, the in-situ health monitoring system, and the VTVL autonomous

flight operation by a small operating crew.

The DC-XA composite intertank [5] resulted in a 44% weight savings over the DC-X

aluminum intertank. The intertank, shown in Figure 12, was constructed in two semi-cylindrical

halves and joined together by fore and aft aluminum attachment rings. The honeycomb

sandwich shell was fabricated using graphite/bismaleimide (T650/5250-4) face sheets and an

aluminum flex-core material. The face sheets were four plies of a fabric in a [0/+-45/90] lay-up.

The aluminum flex-core material was chosen because of ease of fabrication and ready

availability. Some fabrication problems occurred [6] in the form of ruptured core material as a

result of out-gassing of the foaming adhesive used at the high processing temperatures, 440 °F

(227 °C). The processing temperature was reduced to approximately 375 °F (191 °C), which is

below the out-gassing temperature of the foaming adhesive, and no further problems were

encountered. The intertank was ground-tested at NASA Marshall Space Flight Center (MSFC)

to 153% of the DC-XA maximum load, and subsequently ground-tested at White Sands Missile

Range (WSMR) before the flight test program. The intertank experienced no problems during

the flight test program.

The DC-XA composite liquid hydrogen tank [5] resulted in a 34% weight savings over

the DC-X aluminum tank. The liquid hydrogen tank, shown in Figure 13, was constructed in

two cylindrical pieces that were joined together by a “belly wrap” bonded splice joint. The shell

is a 24-ply IM7/8552 graphite/toughened-epoxy laminate. The shell thickness was somewhat

over-designed to avoid leakage. An internal three-dimensional reinforced urethane foam was

used to provide cryogenic insulation. Minor repairs were made to the tank due to a shop accident

and the resulting separation of the insulation from the shell wall. The tank was ground-tested at

MSFC for 29 pressure cycles filled with liquid nitrogen at 150% of the design limit load

pressure, and filled with liquid hydrogen at 100% of the design limit load pressure. The tank

was subsequently ground tested at WSMR for three engine firings and 16 pressure cycles. The

tank then performed flawlessly during the four DC-XA flight tests with no observed leaks. The

DC-XA is the first successful demonstration of a leak-free composite liquid hydrogen cryotank!

The objectives of the X-33 experimental rocket-powered vehicle is to demonstrate critical

technologies for a reusable launch vehicle at hypersonic flight approaching Mach 13. Among

these technology demonstration goals is a liquid hydrogen tank fabricated from composite
7

materials. The structural configuration, see Figure 14, is a complicated four-lobe (quadrant)

conical shell with a noncircular cross-section and a non-spherical two-lobe end cap [7]. The tank

shell is a sandwich construction with IM7/977-2 graphite/epoxy inner and outer face sheets with

a Korex honeycomb core. The internal stiffening substructure is fabricated from textile preform

graphite/epoxy composite materials. The all-composite, all-bonded tank is assembled using a

complex, nine-step 350 °F (177 °C) curing and bonding procedure. In addition to the internal

pressure required to maintain the liquid hydrogen in its liquid state, the vehicle is loaded by

thrust loads during launch that are transferred directly through the liquid hydrogen tank.

Unfortunately, the tank failed during the protoflight ground structural test [7]. The failure was

primarily due to an incomplete understanding of the permeability of liquid hydrogen through

composite materials and a lower than expected, as-manufactured bond strength between the

honeycomb core and the inner face sheet of the tank shell sandwich structure.

Part II. Assessment of the State-of-the-Art

NASA uses a technology readiness level (TRL) scale from 1 to 9 to indicate the level of

maturity of a technology. (The NASA TRL definitions are given in Table 1.) TRL values of 1

to 3 indicate research levels, with TRL 1 being fundamental research. TRL values from 4 to 6

indicate technology development levels. TRL values from 7 to 9 indicate advanced development

levels, with TRL 9 signifying mature technology that is ready for actual aerospace vehicles. As a

developer of advanced technology, NASA usually targets its technology development programs

to advance the technology to TRL 6, and then transitions the technology to the aerospace

industry. The description of TRL 6 is as follows: “system/subsystem validation model or

prototype demonstrated in relevant environment (ground or space).” The NASA TRL scale will

be frequently used in this section to indicate the level of maturity of technologies currently under

development.

Structural Design, Analysis, and Testing: Lessons Learned

Vosteen and Hadcock [8] conducted a study of past composite aircraft structures

programs to identify lessons learned and best practices. Interviews were conducted with 56

people from 32 organizations that were directly involved in design, fabrication, and

supportability of composite structures. The Vosteen and Hadcock [8] survey identified the

following lessons learned relative to structural design, analysis, and testing:

1. Design and certification requirements for composite structures are generally more complex

and conservative than for metallic structures.

2. Successful programs have used the building-block approach with a realistic schedule that

allows for a systematic development effort.

3. The use of basic laminates containing 0/90/+45/-45 plies with a minimum of 10% of the plies

in each direction is well suited to most applications.
8

4. Mechanical joints should be restricted to attachment of metal fittings and situations where

assembly or access is impractical using alternative approaches.

5. Large, co-cured assemblies reduce part count and assembly costs, but may require complex

tooling.

6. Structural designs and the associated tooling should be able to accommodate design changes

associated with the inevitable increases in design loads.

7. Understanding and properly characterizing impact damage would eliminate confusion in the

design process and permit direct comparison of test data.

The Building-Block Approach is the Industry Standard Practice

Successful programs have used the building-block approach to design development and

manufacturing scale-up, illustrated in Figure 15, to develop and verify the structural design

details and manufacturing processes necessary for large composite structure. The complexities

of light-weight, built-up structure led the industry to develop a building-block approach, which is

the standard practice for both metallic and composite structures. The building-block approach

relies on tests of elements and subcomponents to establish the effects of local details and internal

load paths on structural behavior. The building-block approach also must include development

tests to address manufacturing scale-up issues. This requirement is particularly critical in

processing polymeric matrix composite materials where it is particularly challenging to scale-up

accurately the curing kinetics to large-scale component fabrication. The lessons learned by the

industry provide strong motivation for practicing collaborative engineering to design composite

structures that can be reliably manufactured. Experienced materials and processing engineers

should be included in the design phase and must be readily available to correct problems in

production processes when they occur. The building-block approach must be used to avoid overdesigned structure and high-risk structural designs.

The building-block approach relies on tests of coupons, elements, and subcomponents to

establish the effects of local details and internal load paths on structural behavior. These tests

are illustrated in the schematic shown in Figure 16 for a wing structure. By testing at each

hierarchical level of detail, the interaction between the local elements are accurately represented

in the structural design. These development tests can only be omitted if a design-by-analysis

philosophy is supported by reliable, verified, high-fidelity design tools or by adopting a

conservative design philosophy with large factors of safety. Since over-designed (heavier than

necessary) structural components are not desirable and design tools are still under development,

the building-block approach must be used to avoid high-risk structural designs.

While significant improvements have occurred to structural analysis methodologies over

the past two decades, the current structural design and analysis methodologies used by the

aeronautics industry are still largely semi-empirical. Very accurate finite element methods and

sophisticated computer codes are used routinely to calculate the stress, strain, and displacement

fields in complex structural geometries. Superior graphical interfaces have significantly

improved pre- and post-processing of data files. Automated mesh generation, mesh refinement,
9

and automated adaptive remeshing have resulted in major efficiencies in model development

time, analysis time, and accuracy of the numerical solutions. Post-processing algorithms and

graphical interfaces have significantly improved the ability of the analyst to interpret the results

of the stress analysis. However, the prediction of structural failure modes, ultimate strength,

residual strength of damage-tolerant structure, and fatigue life has remained elusive for the

structural engineer. A rigorous structural analysis suitable for predicting structural failure

requires the generation of high-fidelity local stresses that can be used with failure criteria and

damage models. The global/local method, illustrated in Figure 17, is one method currently under

development to predict structural failure. At the present time, global/local analysis methods for

metallic structures [9] are more mature and rigorous than are the corresponding methods for

composite structures. This observation is primarily attributed to the fact that the failure modes

for metallic structures are less complicated and, therefore, more deterministic than is the case for

the failure modes for composite structures. In addition to damage and structural failure,

nonlinear structural response characteristics such as buckling, postbucking and pressurized

structural deformations are more difficult to predict for composite structures than they are for

metallic structures. This difficulty is attributed primarily to the fact that composite materials are

not isotropic, as are metals. Therefore, computational methods that rigorously account for

material orthotrophy and anisotropy should be used for composite structures, and designers

should understand the use of these methods.

Materials, Processes, and Manufacturing: Lessons Learned

The Vosteen and Hadcock [8] survey identified the following lessons learned relative to

materials, processes, and manufacturing:

1. Materials development in conjunction with product development creates undue risks.

2. Experienced materials and processing engineers should be included in the structural design

phase, and must be readily available to correct problems in production processes.

3. Manufacturing process scale-up development tests should be conducted to optimize the

production processes.

4. Co-curing and co-bonding are preferred over secondary bonding which requires near perfect

interface fit-up.

5. Mechanically fastened joints require close tolerance fit-up and shimming to assure a good fit,

and to avoid damage to the composite parts during assembly.

6. Dimensional tolerances are more critical for composite structures than for metallic structures

to avoid damage to parts during assembly. Quality tools are essential for the production of

quality parts.

7. Selection of the tool material depends on part size, configuration, production rate, quantity,

and company experience.
10

8. Tool designers should anticipate the need to modify tools to adjust for part springback, for

ease of part removal, or to maintain dimensional control of critical interfaces.

State-of-the-Art in Materials, Processes, and Manufacturing

Significant improvements in the properties [10] and processability of polymer matrix

composite materials have occurred over the past 30 years. New epoxies, as indicated in Figure

18, have been developed to improve significantly the toughness of composite materials. New

thermosets and thermoplastics, as indicated in Figure 19, have been developed to increase

significantly the use temperature of composite materials. Most epoxies cure at 350 °F (177 °C)

and require an autoclave to insure proper fiber wetting, remove excess resin, minimize porosity,

and promote the polymer cross-linking reaction. Material systems, such as T300/5208 with an

intermediate modulus and intermediate strain-to-failure graphite fibers, have been used to

manufacture structures such as the secondary structures and control surfaces for most

commercial transport aircraft and the primary structures of the B-2. Higher performing material

systems, such as the Toray T800H high-modulus, high-strain graphite fiber, and the toughened

epoxy 3900-2 was used to manufacture the empennage structural components on the Boeing B777. High performance military aircraft, such as the F-22, are manufactured out of materials

systems such as IM-7/5250-4, high-temperature bismaleimide (BMI thermoset) and highmodulus, high-strain graphite fibers.

In recent years, the maturity of composite curing processes, such as resin transfer

molding (RTM) and resin film infusion (RFI), have led to increased use of textile preforms such

as braided, woven, and knitted fiber preforms, and through-the-thickness stitching[11]. These

textile preforms have attractive features for low-cost manufacturing. Automated manufacturing

has been facilitated by the development of high-speed fiber placement and stitching machinery

which were adapted from those used in the textile industry. The development of advanced

processing methods, such as powder-coated fiber tows, also contributed to the automated

features of the processing of the material systems [12]. For example, fibers can be coated with a

dry epoxy powder, braided into the desired form, and then cured so that near net-shape

components can be readily fabricated. As an additional benefit, textile preform composite

materials significantly improve the toughness of the composite material by providing throughthe-thickness reinforcement. For example, the compression-after-impact (CAI) properties of

stitched laminates using low-cost brittle epoxy materials satisfy or exceed the CAI properties of

the significantly more expensive toughened epoxy systems. This technology is being used to

fabricate the braided components of the internal stiffening substructure of the X-33 liquid

hydrogen tank.

Quality Control, NDE/I, and Supportability: Lessons Learned

The Vosteen and Hadcock [8] survey identified the following lessons learned relative to

quality control, nondestructive examination/inspection (NDE/I), and supportability:

1. Automated processes can help to reduce quality control costs.
11

2. Inspection and quality control should focus on aspects of the process and part that have a

direct bearing on part performance.

3. Determine and understand the effects of defects on part performance.

4. Supportability should be addressed during design so that composite structures are inspectable,

maintainable and repairable during service.

5. Most damage to composite structures occurs during assembly or routine maintenance of the

aircraft.

6. Repair costs for composite structures are much higher than for metallic structures.

7. Improved Standard Repair Manuals are needed for in-service maintenance and repair.

8. Special long-life and low-temperature curing repair materials are required.

9. Moisture ingestion and aluminum core corrosion are recurring supportability problems for

honeycomb structures.

The State-of-the-Art of NDE/I Technology

While the visual inspection method remains the method of choice for most airlines,

nondestructive inspection (NDI) methods are also routinely used in both the manufacturing and

flight operations environments. These NDI methods include thermal, ultrasonic,

electromagnetic, radiography, and optical methods. Each method has strengths and weaknesses,

depending on the specific inspection requirement. These NDI methods are listed in Figure 20

and the technology readiness levels (TRL) of the various methods are compared for applicability

to metallic and composite structures with simple and complex configurations. (The comparative

summary given in Figure 20 was prepared by the NASA NDE Working Group.) Referring to

Figure 20, a TRL of 9 means that the technology is mature and is part of the industry standard

practices. The gray boxes without a number mean that the corresponding NDI methods are not

being developed for the specific application. The other colored boxes help to identify similar

TRL levels. The distinction between “conventional” and “advanced” systems refers to the use of

advanced computer-based numerical methods for signal processing. For example, conventional

thermography refers to techniques where the temperature distribution of a structure is mapped

using the image obtained from an infrared camera. The advanced thermography system relies on

sophisticated computational software to analyze time-phased images of the infrared radiation

given off by a structure and provides a map of the heat transfer or diffusivity of the structure. It

has been found that the diffusivity of a structure provides much greater fidelity for determining

the extent of damage than does the corresponding temperature distribution.

Boeing recently conducted an evaluation of current NDI methods for applicability to

inspecting composite fuselage structure [13]. Several methods for detecting defects in stringerstiffened structures and sandwich structures were compared. These methods include through-
12

transmission ultrasonic methods, lamb wave ultrasonic methods, pulse echo ultrasonic methods,

and a specialized C-scan ultrasonic method for disbond detection. The through- transmission

inspection system was found to be the most effective method and was able to resolve defects in

both the skins and the core. However, the technique requires access to both sides of the

component being inspected. Boeing concluded that improvements in current commercially

available systems would be necessary to inspect composite sandwich structures reliably in the

field.

Assessment of the Technology Requirements for an RLV

NASA has established a goal [14] of an order-of-magnitude reduction in the cost of

launching a pound of payload to low earth orbit from current costs of about $10,000 per pound.

A single-stage-to-orbit, reusable launch vehicle (RLV) is a leading concept for achieving this

dramatic reduction in the payload launch cost. An RLV will have to operate much more like an

airplane than the current Space Shuttle. The RLV must be robust, reliable, and require minimal

inspection and maintenance between flights. Recent systems studies have shown that

considerable reductions in the mass-fraction of the vehicle using conventional technologies must

be achieved to reduce the gross take-off weight to a level where the vehicle can achieve orbit.

(Mass-fraction is the ratio between the structural weight and the gross take-off weight.) Of all

the technologies that may reduce the mass-fraction [6], the application of advanced composite

materials for the primary structures and for the liquid hydrogen tank is projected to have the

greatest potential for achieving the current take-off weight goals.

Over the past several years, NASA planning teams have evaluated various technologies,

have estimated current technology readiness levels, and have prepared roadmaps for developing

the technologies that will be enabling for an RLV. A summary of the enabling structures and

materials technologies is shown in Figure 21 along with an estimate of the current TRL. (See

Table 1 for the TRL definitions.) For current purposes, Primary Structures are defined as all

load-bearing structures, exclusive of the integral cryogenic tanks. Cryotanks are defined as all

elements of the cryogenic tank system, including the tank pressure vessel structure and the

cryogenic insulation. Thermal protection systems (TPS) are defined as all elements of the

vehicle thermal protection systems including both external TPS surfaces and internal insulation.

Hot structural concepts applicable to vehicle leading-edges and control surfaces are also

considered to be TPS subsystem elements. Some of the technologies listed in Figure 21 are

much more complicated than others. The TRL is an estimate of the total technology, even

though parts of the technology may be at a substantially higher TRL than the number listed in the

figure. The technical evaluation summarized in Figure 21 indicates that extensive development

of structures and materials technologies will be required to enable an RLV that will replace the

Space Shuttle.

Lessons Learned from Technology Development Structural Tests

Over the past two decades, there have been a number of technology development

programs that have designed, manufactured, and tested large composite structural components.

Each of these tests is a source of lessons learned and provides valuable insight into further
13

developmental requirements. The following paragraphs summarize the results of six significant

structural tests that have provided lessons learned.

Under contract to NASA, Lockheed-Martin designed and manufactured a large

technology integration box beam [15]. The configuration, shown in Figure 22, resembled the C130 center wing box and was about 150 inches (3.81 m) long, 50 inches (1.27 m) wide, and 28

inches (0.71 m) deep. The stiffness requirements were established to meet the commercial flutter

requirements specified in FAR Part 25 [1]. The damage tolerance requirements satisfied both the

FAR Part 25 requirements and the corresponding military requirements. Two graphite/epoxy

material systems (AS4/1806 and AS4/974) were used. The ribs and spars were mechanically

fastened to the cover panels. The test plan included a test of the box beam to design limit load

with down-bending plus torsion loads, a test to design ultimate load with up-bending plus torsion

loads, and a residual strength test to failure with impact damage at several locations with upbending plus torsion loads. The box failed prematurely during the design ultimate load test at

only 125% of the design limit load condition (83% of design ultimate load). An extensive failure

investigation [15] determined that the failure initiated in the upper cover skin due to severe local

bending of the skin in the region of the hat stiffener termination. It was found that a very small

gap of unstiffened skin between the hat stiffener termination and the rib shear tie, as indicated in

Figure 22, led to a local short-wave-length shear-crimping failure in the skin that resulted in the

complete failure of the box. This unexpected failure highlighted the sensitivity (criticality) of

composite structures to local structural detail features and complex local stress gradients.

A composite wing stub box was designed and fabricated by McDonnell Douglas, and

tested by NASA Langley Research Center [16]. The wing stub box was the first of two major

technology demonstration milestones in the NASA Advanced Subsonic Technology Composite

Wing Program. The objective of the wing stub box was to demonstrate the viability of low-cost

manufacturing technology. The wing stub box was fabricated using graphite/epoxy textile

materials (AS4/3501-6 and IM7/3501-6) and stitched together using Kevlar thread. The IM7

graphite fibers were used only for the 0-degree fibers in the lower cover panel skin. The

composite skin and stiffeners were composed of layers of dry fiber preforms that were prekitted

in nine-ply-thick stacks with a quasi-isotropic stacking sequence. The resin film infusion (RFI)

process was used to impregnate the dry fiber preforms with resin and the subsequent composite

structures were cured in an autoclave. The composite test article, shown in Figure 23 (a), was

attached to a metallic extension box to provide a load transition section so that loads

representative of a transport wing structure could be applied to the stub box. The test plan

included design limit load and design ultimate load tests with impact damage. The comparison

of the test results to finite element model predictions, shown in Figure 23 (b), was excellent [16].

The model accurately predicted the onset of buckling in the cover panels of the box. The box

failed at only 93% of the design ultimate load, which was slightly less than expected. The posttest investigation determined that the failure was initiated by nonvisible impact damage in the

web and flange of a stringer that terminated near the front spar. This test highlighted the

sensitivity of composite structures to nonvisible impact damage in regions of load redistribution

such as the stringer termination. The cover panel was designed to account for compression-afterimpact conditions, but the damaged stringer added a transverse shear load component to the
14

locally damaged area of the cover panel. This local shear load component created a local

combined load condition in the damaged area of the cover panel.

June 1, 2000, marked the formal completion of the Advanced Subsonic Technology

Composite Wing Program with the successful test of a 42-ft-long (12.8 m) Stitched/Resin Film

Infused (S/RFI) Composite Wing Box [17]. The wing box, a manufacturing technology

verification article, was designed and fabricated by the Boeing Company under contract to

NASA to satisfy the requirements of a 220-passenger commercial transport aircraft. The S/RFI

manufacturing process stitches together layers of multi-axial warp-knit graphite/epoxy fabric

using Kevlar thread, and then impregnates and cures the resulting preform using the resin film

infusion process. The tests of the wing box was conducted at NASA Langley Research Center

and the wing is shown in Figure 24 for a design limit load condition. Prior to the failure test, the

wing box was subjected to several design limit load (DLL) tests to measure structural response

and to verify the accuracy of nonlinear finite element analysis procedures used to predict the

wing box response. These tests included a 100-percent DLL test representative of a braked rollout condition, a test with a 1-g down-bending condition, and a test with a 2.5-g up-bending

condition. In addition, the wing was subjected to 7 inch (17.8 cm) long saw cuts, which are

representative of discrete source damage, in the upper and lower cover panels. The wing box

successfully supported the 70-percent DLL requirement with the saw cuts as required by the

FAR Part 25 [1]. Prior to the design ultimate load (DUL) test, the saw cuts were repaired by an

airline maintenance contractor to restore the wing box to its design ultimate load capability.

Also, prior to the DUL test, the upper and lower cover panels were subjected to local impact

damage events with impact energies ranging from 83 to 100 ft-lbs (113 to 136 N•m) to simulate

foreign object damage. Sections of the wing structure were nearly 1-inch (2.54 cm) in thickness

and were subjected to average running loads greater than 24,000 lbs/in (4.20 MN/m). The wing

box failed at 97-percent of the DUL requirement with unrepaired nonvisible damage. This

failure load is within the failure prediction accuracy of the finite element analysis used for this

complex structure, and within the experimental scatter band for typical material properties. The

ability of the wing box to sustain discrete source damage and foreign object damage

demonstrates the robustness of the S/RFI composite manufacturing process and validates the

accuracy of state-of-the-art damage-tolerance analytical methods for primary composite aircraft

structures. (While few details were given, Aviation Week & Space Technology, March 20, 2000,

p. 61, reported that DASA Airbus successfully completed a similar test program on a full-scale

carbon fiber reinforced plastic wing and wing box.) These test programs clearly demonstrate

that composite structures and materials can be scaled-up effectively to realistic, heavily loaded

aircraft primary structures.

A full-scale segment of a reusable launch vehicle prototype wing was fabricated as a test

article to demonstrate the integration of the thermal protection system (TPS) with large

composite structural components and to validate the fabrication, design, and analysis methods for

this wing [6,18]. A honeycomb sandwich construction was selected to provide broader design

and fabrication experience. The upper and lower skin panels were fabricated using a

graphite/bismaleimide (IM7/5250-4) material system. This material system was selected

because it has good fracture toughness and good mechanical properties at elevated temperatures

up to 350 °F (177 °C). The honeycomb core was glass/polyimide HRH-327 with a 3/16-in.

(0.476 cm) cell size and a 4.5 lbs/ft3

 (72.1 Kg/m3) density. The wing box is approximately 10
15

feet (3.05 m) long, 5 feet (1.52 m) wide, and 43 inches (1.09 m) deep with three ribs and three

spars. While the wing box was not subjected to an elevated temperature test condition, three

different types of TPS were installed on the upper skin to demonstrate the load carrying

capability of the integrated structure. The test was conducted at NASA Langley Research Center

and the test set-up is shown in Figure 25 (a). The wing box was loaded to design limit load and

to design ultimate load with both up-bending and down-bending loading conditions. The box

was then loaded to failure with the up-bending loading condition. Selected measured strain

values recorded during the tests are shown in Figure 25 (b), and the results are in excellent

agreement with the values calculated by the finite element analysis. The predicted upper skin

buckling load was within 3% of the experimental value. The predicted shear failure load was

within 5% of the experimental value. While considerable work is still required to develop

manufacturing technology that can be scaled-up to an RLV size vehicle, the success of this test

clearly indicates the viability of composite structures technology for primary structures

applications to reusable launch vehicles.

A composite intertank design for the body of a reusable launch vehicle was developed,

and a full-scale segment was fabricated and tested [6, 19]. The intertank was designed to contain

the payload for the vehicle and, therefore, would have payload bay doors. The critical design

condition is the compressive load due to maximum ascent acceleration; and the load transfer

around the payload bay doors is a major design consideration. A design trade study resulted in

the selection of a stiffened-skin configuration with internal frames. The graphite/bismaleimide

(IM7/5250-4) material system was selected for the skin, stiffeners, and frames due to its good

fracture toughness and good mechanical properties at temperatures up to 350 °F (177 °C). A

curved section of the intertank design was selected as a structural test article. The test article,

shown installed in the test facility in Figure 26 (a), was approximately 10 feet (3.05 m) long and

22 feet (6.71 m) wide, and includes about a 90-degree section of the intertank. The test was

conducted in a structural test facility at NASA Langley Research Center. The test article failed

prematurely when subjected to a compression load due to the separation of the hat stiffeners

from the skin at approximately 70% of the predicted failure load. The failed test article is shown

in Figure 26 (b) with a buckled skin. The premature failure was attributed to a poorly

manufactured bond between the hat stiffeners and the skin. This test illustrates the critical need

to include manufacturing scale-up development tests in the building-block approach to the design

and fabrication of large-scale structural components.

The development of a lighter-weight liquid hydrogen tank is continuing with a series of

ground tests at NASA Marshall Space Flight Center, see Figure 27 [20]. The objectives of the

ground test program are to verify the structural integrity of an RLV flight weight tank, to verify

the structural design and analysis methods, to verify the impermeability of the tank skin to liquid

hydrogen, and to verify the lifetime performance of the insulation. The 8 ft (2.44 m) diameter

tank, 1/4- scale for an RLV, was fabricated using fiber-tow-placed IM7/977-2 graphite/epoxy

material and co-bonded stiffeners. A honeycomb core insulation was bonded to the outside of

the tank. The tank shell wall thickness is only 14 plies in contrast to the over-designed 24-ply

thickness of the DC-XA tank. To date, five pressure cycles in the test plan have been completed

with liquid nitrogen. While the tank is structurally sound, a few minor leaks have required some

repairs to the tank. The pressure testing with liquid nitrogen is continuing, and the test program

will eventually begin testing with liquid hydrogen. While considerable work is still required to
16

develop manufacturing technology that can be scaled-up to an RLV size vehicle, the success of

this test indicates the viability of composite structures technology for the cryotanks of an RLV.

Part III. Future Directions

Revolutionary Structural Concepts for Next Generation Aircraft

The quest for improved materials for aerospace vehicles is never ending. Design and

market drivers include lower weight, improved corrosion and fatigue resistance, and lower

acquisition and operation costs. It is interesting to contemplate the current use of composite

materials on commercial transport aircraft and to try to extrapolate to the next generation aircraft.

The most significant current barriers to increased use of composite materials are high

manufacturing costs, poor reliability in estimating the design and development costs, and the

inability to predict accurately structural failure. As illustrated in Figure 3, the advantage of

composite materials in secondary structures and lightly loaded primary structures has been moreor-less fully demonstrated. Given the current state of the technology, a consensus has emerged

within the community that the next step in the evolution of composite structures for commercial

transport aircraft applications is a composite wing. Beyond this developmental step, the

marketplace will decide the next opportunity for composite materials and structures. For

example, composite materials may prove to be an enabling technology for a new class of aircraft

that have superior performance characteristics compared to today’s commercial transport

aircraft. Several revolutionary aircraft configurations are illustrated in Figure 28. The

aerodynamic performance of these vehicles may prove to be quite superior to conventional

subsonic aircraft. The potential benefits that may be derived from these revolutionary aircraft

include significant increases in flight range or performance, significant reductions in fuel

consumption, significant reductions in engine emissions, and significant reductions in airframe

and engine noise. However, major improvements in the current state-of-the-art for composite

structures will be required to design and build these new aircraft reliably and economically. For

example, the noncircular cross-section and compound curvature features of the blended wing

body configuration will be a particularly significant challenge for structural designers.

Current research is expected to result in dramatic improvements in structural design and

analysis tools. Reliable, advanced analysis methods will significantly reduce current dependence

on the empirical design approach and provide better capability to optimize structural designs.

High-fidelity, physics-based structural analysis tools are under development using both

deterministic and non-deterministic computational methods. Rigorous, physics-based

computational methods to predict accurately damage initiation and growth, structural failure

modes, and the residual strength of damaged structure remains a grand challenge that is

motivating considerable research attention in the structures community. Next generation

structural design tools are under development that exploit the revolution in information

technology. The use of intelligent systems to improve graphical user interfaces and threedimensional immersive simulation of structural analysis results is illustrated in Figure 29. As

illustrated in Figure 29 (a), the next generation design tools will use libraries of smart

components to assemble finite element analysis models easily. Interface elements are under

development that will provide seamless transitions between regions of a finite element model

with different mesh refinements. These advanced methods not only automate model generation,
17

but also facilitate the implementation of global-local modeling strategies that are essential for the

prediction of progressive damage and structural failure. Finally, advanced three-dimensional

virtual reality capabilities, such as the system shown in Figure 29 (b), will greatly enhance our

ability to interpret the results of structural analyses.

Breakthroughs in Materials Synthesis and Processes

Current manufacturing technology requires an autoclave to cure polymer matrix

composite materials to provide high-quality, high-performance structural components.

Eliminating the autoclave will dramatically lower the cost and complexity of manufacturing

composite structures. Revolutionary new methods of curing composite materials are being

developed to eliminate the autoclave from the curing process. These new out-of-the-autoclave

processes, such as the electron-beam magnetic suspension process shown conceptually in Figure

30, may also facilitate the manufacturing of virtually final, near-net shape components. The

concept illustrated in Figure 30 has the added advantage of eliminating the expensive tools

currently required to make composite parts where precision controlled design tolerances are

required. This technology could eliminate the need for mechanical fasteners and adhesive bonds,

except for those associated with major airframe assembly splices.

The general field of nanotechnology offers the potential to be the next great industrial

revolution. In the field of materials science, a paradigm shift may occur away from the

traditional materials role of developing metallic, polymeric, ceramic, and composite materials to

a revolutionary role of developing nanostructured, functionalized, self-assembling, and selfhealing materials. Looking into the future, the theoretical potential of these revolutionary classes

of new materials will create breakthroughs that will enable technology developments that are

barely imaginable today. In the aerospace field, these new technologies may make space travel

routine and enable human exploration of space beyond our current practical limitation of low

Earth orbit. Imagine the possibilities if there was a material to replace aluminum that is an order

of magnitude stiffer and two orders of magnitude stronger!

Breakthroughs in the methods used to synthesize new polymer chemistries will lead to

highly tailored materials with significantly improved properties. For example, advanced

computer software is being developed to exploit new knowledge in nanostructure-property

relationships. Referring to Figure 31, “computational materials” is one of the emerging fields of

“computer-designed materials” that is attempting to build a bridge between our knowledge of

quantum physics and continuum-mechanics-based micromechanics. Computer models are

generated at the atomic and molecular levels that model the relationship between the atomic

structure of the material at the nano-scale and the physical properties exhibited at the macro-scale.

First principles of quantum mechanics, molecular dynamics, thermodynamics, and continuum

mechanics are being used to predict the properties of new material chemistries. These new

computational tools have an extraordinary potential to optimize chemistries for specific

performance goals and to conduct trade-off studies that quantify the effects of changes in

chemistry on various material properties.
18

Computer simulation results and limited experimental studies show that small diameter,

single-walled carbon nanotubes (CNT) may possess elastic moduli in excess of 145 Msi (1 TPa),

and strengths approaching 29 Msi (200 GPa) [21-34]. If small-diameter, single-walled tubes can

be produced in large quantities, and incorporated into a supporting matrix to form structural

materials, the resulting structures could be significantly lighter and stronger than those made

from current aluminum alloys and carbon fiber reinforced polymer (CFRP) composite materials

used in conventional aerospace structures. Properties of single-wall carbon nanotubes (SWNT)

and multi-wall carbon nanotubes (MWNT) reported in the literature exhibit quite a range in

values. Theoretical properties have been determined from computer simulations using quantum

mechanics [29, 30], atomistic simulation (molecular dynamics) [31, 32], and continuum

mechanics [33, 34]. Experimental measurements of properties have been reported using atomic

force microscopy and Raman Spectroscopy. The table below illustrates the variability in the data

reported in the literature [21-28]:

Tensile modulus 44 to 260 Msi (300 to 1800 GPa)

Tensile ultimate strength 0.9 to 26 Msi (6 to 180 GPa)

Bending strength 0.9 to 3 Msi (6 to 22 GPa)

Elongation (strain to failure) 6 to 15%

Thermal conductivity 1000 to 3300 Btu ft-1 hr-1 °F-1 (1750 to 5800 W m-1 K-1)

The specific modulus and specific strength of several aerospace materials currently used

in structural components of aerospace vehicles are compared in Figure 32. The properties of

sheet and plate forms of aluminum 2219 alloy were obtained from MIL-HDBK-5D [35]. The

CFRP composite material indicated in the figure is a high-modulus, high-strength fiber in a

toughened polymer matrix with a quasi-isotropic laminate stacking sequence and a 60% fiber

volume fraction [36]. Theoretical properties of the carbon nanotube fiber reinforced polymer

(SWNTFRP) composite were calculated using standard micromechanics equations. The

modulus of the SWNTFRP was assumed to be 174 Msi (1200 GPa). The SWNTFRP laminate

is assumed to be the same laminate as the CFRP laminate and the strength was limited to 0.9 Msi

(6 GPa) (1% strain) to reflect current structures design practices. The single crystal bulk material

(SWNT) plotted in Figure 32 represents the theoretical potential of nanostructured carbon that

will require several breakthroughs in nanotube production technology to achieve. This highly

perfect, single crystal, bulk material does not require a matrix binder material and is viewed as

theoretically possible. As is evident from Figure 32, the polymer composite reinforced with

nanotubes offers a significant advantage over conventional aluminum and carbon fiber

reinforced polymer composite materials.

The theoretical properties of the SWNTFRP were used in a simple, systems analysis

concept model of a reusable launch vehicle shown in Figure 33 (a). The computed vehicle dry

weight results are shown in the accompanying bar chart in Figure 33 (b). Dramatic reductions in

weight are achievable by replacing the aluminum components of the airframe structure and the

cryogenic propellant tanks with the CFRP composite material or the SWNTFRP composite

material, and then resizing the vehicle.
19

Evolution of Composite Structures NDE Technology

Nondestructive evaluation (NDE) techniques are currently used during component

manufacturing, design certification, maintenance, inspection, and repair. Current research is

focused on exploiting the role of computer simulations [37] to revolutionize the traditional NDE

role, see Figure 34. It is generally understood that NDE issues that are not addressed during the

component design stage must be addressed later in the manufacturing stage. This staging of the

use of NDE procedures can be, potentially, at a much higher cost as maintenance and repair

considerations increase with component age. If validated and robust NDE simulations are

available during the initial design stage, then component configurations may be adjusted in “realtime” to lower the overall life cycle NDE costs while maintaining optimized system level

benefits. Furthermore, these benefits are enhanced when manufacturing simulations make use of

NDE process control simulations. Validated simulations of NDE for process control, when

incorporated or embedded into the manufacturing process control, can reduce or eliminate

manufacturing process steps, including conventional inspections, while further optimizing the

yield of the manufacturing process.

For the foreseeable future, structural components will continue to incur operational

service-induced damage and degradation. The requirement to evaluate component integrity and

repair or replace damaged components will continue to challenge the NDE community. In the

future, NDE simulations may be optimized to the point that they may be used to generate the

plans for in-service maintainability and repair. Issues such as component design and functional

specifications, work space geometry and component access, and accept/reject criteria or

retirement-for-cause criteria will need to be incorporated into these NDE simulations. It is

anticipated that NDE technology will evolve to a state-of-the-art where virtual reality NDE

simulations in design, smart health monitoring systems, and telerobotic inspection and repair are

commonplace. The challenge for the NDE community is to develop and validate virtual reality

simulations that are robust and adaptable enough to function smoothly and autonomously.

Next Generation Design Tools and Collaborative Engineering

In order for NASA to meet its unique mission needs in space science, human exploration,

earth science, and aeronautics, NASA has a new initiative to develop an intelligent synthesis

environment (ISE). The ISE will utilize computational intelligence to synthesize existing,

developing, and future relevant technologies to create a new product and mission development

environment. In the ISE, synthesis will take place in three ways: synthesis of scientists,

engineers, technology developers, operational personnel and training personnel all working in

geographically distributed locations (collaborative engineering); synthesis of cutting-edge

technologies and diverse, life-cycle design tools seamlessly integrated together both horizontally

and vertically at all levels of fidelity; and synthesis of computers, intelligent hardware (robotics),

synthetic (virtual reality) simulated designs, and design languages. The intelligent nature of the

ISE will be derived from its concentrated use of non-traditional, intelligent computational

systems such as intelligent product objects, intelligent agents, and intelligent computational

methods. Computational intelligence will guide the utilization of vast resources of knowledge

and predictive capability that is built directly into the design environment. Effective

collaborative engineering, illustrated conceptually in Figure 35, will require that numerous
20

validated computer software modules, which represent a wide range of scientific disciplines, will

operate together in a robust fashion to yield credible optimized mission performance [38]. The

path to achieving credible optimized structural designs is dependent on the fidelity of the

individual software modules and the optimization processes.

Summary of Assessment

Aerospace structural components are usually designed to be very close to a zero design

margin. While the margin of safety is not equal to zero for all of the design criteria at each

structural location, there is typically one criterion for each structural element that governs the

design details for that element. The quest for the lowest weight structure then drives the design

margin to be nearly equal to zero for the design limit load condition. The factor of safety

between the design limit load and the design ultimate load conditions accounts for the difference

between linear, elastic behavior and complete structural failure, and for uncertainties in other

parameters such as loads and material properties. Therefore, aerospace structural designs do not

have a large factor of safety to accommodate any unanticipated deleterious structural behavior.

Composite structures fail differently than metallic structures. The 65 years of successful

experiences with the design of metallic structures cannot be directly transferred to the design of

composite structures. First, composite materials are not isotropic like most metallic alloys.

Second, the initiation and growth of material level damage and the failure modes of composite

structure are not well understood and cannot always be predicted accurately. Due to these

complications, the best design practices for composite structures are fully understood only by

those engineers who are experienced at designing composite structures.

Composite structural design and manufacturing technology is not yet fully mature for all

applications. There are three key factors that contribute to the lack of maturity of the design and

manufacturing technology for composite structures. These factors are the lack of a full

understanding of damage mechanisms and structural failure modes, the inability to predict

reliably the cost of developing composite structures, and the high costs of fabricating composite

structures relative to conventional aluminum structures. While the technology required to

overcome these uncertainties is under development, these factors are barriers to expanding the

application of composite materials to heavily loaded, primary structures. For those applications

where development and fabrication costs are not a factor or where risks to aircraft structural

integrity are low, there is extensive use of composite structures.

Successful programs have used the building-block approach to structural design and

manufacturing process development with a realistic schedule that allows for a systematic

development effort. The complexities of lightweight, built-up structure led the industry to

develop a building-block approach, which is the standard practice for both metallic and

composite materials. The building-block approach relies on tests of elements and

subcomponents to establish the effects of local design detail features and load paths on structural

behavior. The building block approach also must include development tests to address

manufacturing scale-up issues. This observation is particularly critical in processing polymeric

matrix composite materials where it is particularly challenging to scale-up the curing kinetics to

large-scale component fabrication. The lessons learned by the industry provide strong
21

motivation for practicing collaborative engineering to design composite structures that can be

reliably manufactured. Experienced materials and processing engineers should be included in

the structural design phase of a project and must be readily available to correct problems in

production processes when they occur. The building-block approach must be used to avoid overdesigned structure and high-risk structural designs.

Maintenance, inspection, and repair technologies for composite structures are not yet

fully mature for all applications. Technologies in everyday use today to support metallic

structures do not apply to composite structures. Furthermore, the long-term, field experiences

necessary to develop a support infrastructure does not yet widely exist for composite structures.

Therefore, support issues must be anticipated in the design phase for composite structures to help

facilitate effective maintenance, inspection, and repair procedures. Structures must be designed

so that they that can be inspected and repaired in the field. In addition, NDE experts should be

part of the collaborative engineering team so that inspectability is built into the structural design

from the outset of the design.

References

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Office of the Federal Register, National Archives and Records Administration, U. S.

Government Printing Office, Mail Stop SSOP, Washington, D.C. 20402, 1996.

2. Dow, M. B., "The ACEE Program and Basic Composites Research at Langley Research

Center (1975 to 1986), Summary and Bibliography,” NASA Reference Publication 1177,

National Aeronautics and Space Administration, Langley Research Center, Hampton, VA

23681, October 1987

3. Dexter, H. B., and Baker, D. J., "Flight Service Environmental Effects on Composite

Materials and Structures,” Advanced Performance Materials, Vol. 1, 1994, pp. 51-85.

4. Jackson, P., Editor-in-Chief, Jane’s All The World’s Aircraft, 1998-99 Edition, Jane’s

Information Group Limited, Coulsdon, Surrey CR5 2YH, UK, 1998.

5. Schweikle, D., “DC-XA Incident Investigation Summary,” briefing to the NASA DC-XA

Mishap Investigation Board, August 5, 1996.

6. Sawyer, J. W., “Graphite-Composite Primary Structure for Reusable Launch Vehicles,”

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9. Harris, C. E., Newman, J. C. Jr., Piascik, R. S., and Starnes, J. H., Jr., “Analytical

Methodology for Predicting Widespread Fatigue Damage Onset in Fuselage Structure,”

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12. Shuart, M. J., Johnston, N. J., Dexter, H. B., Marchello, J. M., and Grenoble, R. W.,

“Automated Fabrication Technologies for High Performance Polymer Composites,”

Composite Fabrication, Vol. 14, No. 8, August 1998, pp.24-30.

13. Flynn, B. W., “Advanced Technology Composite Fuselage-Repair and Damage Assessment

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Administration, Langley Research Center, Hampton, VA 23681, 1997.

14. Anselmo, J. C., “NASA Issues Wake-Up Call To Industry”, Aviation Week and Space

Technology, February 19, 1996, pp. 20-21.

15. Shuart, M. J., Ambur, D. R., Davis, D. D., Jr., Davis, R. C., Farley, G. L., Lotts, C. G., and

Wang, J. T., “Technology Integration Box Beam Failure Study,” NASA Conference

Publication 3178, Part 2, National Aeronautics and Space Administration, Washington, D.

C., December 31, 1994, pp. 951-965.

16. Wang, J. T., Jegley, D. C., Bush, H. G., and Hinrichs, S. C., “Correlation of Structural

Analysis and Test Results for the McDonnell Douglas Stitched/RFI All-Composite Wing

Stub Box,” NASA Technical Memorandum 110267, National Aeronautics and Space

Administration, Washington, D. C., July 1996.

17. Phillips, E. H., “Composite Wing Box Tested to Failure,” Aviation Week & Space

Technology, June 21, 2000.

18. Sawyer, J. W., Bush, H., and Sutter, T. R., “Experimental Investigation of a GraphiteComposite Wing-Box section for a Reusable Launch Vehicle,” Proceedings of The Space

Technology and Applications International Forum, CONF-970115, El-Genk, M. S., Ed.

American Institute of Physics, New York, AIP Conference Proceedings No. 387, Vol. 1, No.

3, January 26-30, 1997, pp. 1245-1258.
23

19. Sawyer, J. W., and Bush, H., “Experimental Investigation of a Graphite-Composite Intertank

Section for a Reusable Launch Vehicle,” Proceedings of the Space Technology and

Applications International Forum, 3rd Conference on Next Generation Launch Systems,

CONF-980103, El-Genk, M. S., Ed. American Institute of Physics, New York, AIP

Conference Proceedings January 25-29, 1998, pp. 1007-1019.

20. Anon., "Reusable Composite Hydrogen Tank System (RCHTS) TA-1 Final Tank Design

Report,” Boeing Internal Report, Huntington Beach, CA., June 20, 1997.

21. Halicioglu, T., “Stress Calculations for Carbon Nanotubes,” Thin Solid Films, Vol. 312,

1998, pp11-14.

22. Hernandez, E., and Goze, C., “Elastic Properties of Single-Walled Nanotubes,” Applied

Physics A, Vol. 68, 1998, pp. 287-292.

23. Lu, J. P., “Elastic Properties of Carbon Nanotubes and Nanoropes,” Physical Review Letters,

Vol. 79, No. 7, 1997, pp. 1297-1300.

24. Sinnott, S. B., Shenderova, O. A., White, C. T., and Brenner, D. W., “Mechanical Properties

of Nanotubule Fibers and Composites Determined From Theoretical Calculations and

Simulations,” Carbon, Vol. 36, Nos. 1-2, 1998, pp. 1-9.

25. Treacy, M. J., and Ebberse, W., “Exceptionally High Young’s Modulus Observed for

Individual Carbon Nanotubes,” Nature, Vol. 381, 1996, pp. 678.

26. Wong, E. W., and Sheehan, P. E., “Nanobeam Mechanics: Elasticity, Strength, and

Toughness of Nanorods and Nanotubes,” Science, Vol. 277, 1997, pp. 1971-1975.

27. Yao, N., and Lordi, V., “Young’s Modulus of Single Walled Carbon Nanotubes,” Journal of

Applied Physics, Vol. 84, No. 4, 1998, pp. 1939-1943.

28. Yu, M.-F., Lourie, O., Dyer, M. J., Moloni, K., Kelly, T. F., and Ruoff, R. S., “Strength and

Breaking Mechanism of Multiwalled Carbon Nanotubes under Tensile Load,” Science

Magazine, Vol. 287, No. 5453, 2000, pp. 637-640.

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Under Uniaxial Compression,” Physical Review Letters, Vol. 83, 1999, pp. 2973.

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25

Table 1. NASA Technology Readiness Levels (TRL).

TRL Description 

1 Basic principles observed and reported

2 Technology concepts and/or applications formulated

3 Analytical & experimental critical function and/or characteristic proof-ofconcept

4 Component and/or breadboard validation in laboratory environment

5 Component and/or breadboard validation in relevant environment

6 System/subsystem validation model or prototype demonstration in relevant

environment (ground or space)

7 System prototype demonstration in an air/space environment

8 Actual system completed & flight qualified through test and demonstration

(ground or flight)

9 Actual system flight proven through successful mission operations
Figure 1. Typical commercial applications initiated by the ACEE Program.

b. Boeing B-737 composite 

horizontal stabilizer

c. Douglas DC-10 composite

Rudder and vertical stabilizer

a. Boeing B-727 composite 

elevator

d. Lockheed L-1011 composite 

aileron

26
Figure 2. Composite structural applications in commercial transport aircraft.

1965 1970 1975 1980 1985 1990 1995

Composites 

as % of 

structural

weight

10

15

DC-9

B-747

L-1011

MD-80 B-737-300

B-747-400

B-757 MD-90

B-767

A300-600 MD-11

A310 B-777

A340 A330

A320 A321

1

2

3

4

5

A322

35

20

30

DC-10

In commercial 

transports, cost

has kept 

composite 

applications low

In commercial 

transports, cost

has kept 

composite 

applications low

Year of first flight

27
Figure 3. Structural composites on the Boeing B-777.

Figure 4. Composite structural applications in rotorcraft and 

general aviation.

28
Figure 5. Bell-Boeing V-22 Osprey.

Figure 6. Composite structural applications in military fighter aircraft.

29
Figure 7. Lockheed Martin F-22 Raptor.

Figure 8. Assembly of F-22 composite wings.

30
Figure 9. B-2 primary structure Is almost all composite materials.

Figure 10. C-17 horizontal tail redesigned using composite materials.

31
Figure 11. DC-XA experimental rocket aircraft.

Figure 12. DC-XA composite intertank.

32
Figure 13. DC-XA composite LH2 cryotank.

Figure 14. X-33 all-bonded, all-composite LH2tank.

33
Figure 13. Building-Block Approach is Industry Standard

Coupon testing for

material properties

Design concepts and

analysis development

Manufacturing process 

development and scale-up

Concept demonstration 

at component level

Full-scale 

structural

verification

Figure 15. Building-block approach is the industry standard practice.

34
Figure 16. Design development tests in building-block approach.

Figure 17. Global/local analysis for predicting structural behavior.

35
Figure 18. Evolution of composite resin development: Epoxies.

Figure 19. Evolution of composite resin development:

Intermediate and high temperature resins.

Advancements

in composite

technology

1960 1970 1980 1990 2000

Autoclave and vacuum hot press

curing (TRL=9)

Affordable processing (TRL 2-6)

E- beam cures

Non-nutoclave curing

RFI/stitched preforms

Textile

preforms

8551-7

3900-2

977

LTM45EL

}

Toughened

epoxies

Toughened

thermoplastics

T300 / 5208

Brittle

epoxies:

MY-720

ERL-0510

Carbon,

Boron,

S-Glass L1011, DC10

B 727, B 737

F14, F16

F18

AV8B

B757

B767

Lear Fan

A-6 

B-2

F 117

V22 

B777 

AS4 / 3501-6

Year

Year

36
NDE TECHNOLOGY

BASIC 

METAL 

STRUCTURES

BASIC 

COMPOSITE 

STRUCTURES

COMPLEX 

METAL 

STRUCTURES

COMPLEX 

COMPOSITE 

STRUCTURES

Planar, 

slight 

curvature 

Planar, 

slight 

curvature

B-2, B-777

Irregular, 

curved, 

hybrid, 

bonded, 

honeycomb,

built-up 

structure

Irregular, 

curved, 

hybrid, 

bonded, 

honeycomb,

built-up 

structure,

X-33, X-34

Conventional thermography 9

Advanced thermography 9 9 4 4

Conventional ultrasonics 9

Advanced ultrasonics 9 9 3 3

Conventional X-radiography 9 9

Reverse geometry X-ray 6 5 4 3

Computed tomography 9 9 9 6

Backscattered X-ray 5

Conventional eddy current 9

Advanced eddy current 9 4 3

Optical shearography 6 4 3 3

Microwave 3 4 3 3

Conventional acoustic emissions 9 6

Advanced acoustic emissions 9 6 3 3

Visual 9

Penetrants (surface defects) 9 9

Magnetic particle (surface defects) 9 9

Numbers in the table are TRL’s

In-Situ vehicle health monitoring 3 3

Figure 20. The state-of-the-art of NDE/I technology.

37
Figure 21. Assessment of technology needs for an RLV.

Leading edges / nose caps

• Refractory composites (TRL=9)

• Hot-structure control surfaces (TRL=5)

• High-temperature heat pipes (TRL=4)

• Active cooling (TRL=4)

Thermal protection system

• High-temperature metallics (TRL=5)

• Refractory composites (TRL=4)

• Rigid ceramics (TRL=9)

• Advanced flexible insulation (TRL=6)

• Validated thermal analysis/sizing (TRL=5)

Primary structure

• High-temperature polymeric composites (TRL=5)

• High-temperature metal composites (TRL=4)

• Noncircular composite shell structures (TRL=3)

• Joints and attachment techniques (TRL=4)

• Validated design and analysis methods (TRL=4)

• Nondestructive evaluation (TRL=4)

• Manufacturing technology (TRL=4)

Cryotanks

• Sandwich construction (TRL=4)

• Nonautoclave curing (TRL=3)

• Manufacturing scalability (TRL=4)

• Nondestructive evaluation (TRL=4)

• Vehicle health monitoring (TRL=3)

• Materials design allowables (TRL=4)

• Integrated TPS / cryo-insulation (TRL=2)

38
Figure 22. Technology integration box beam test.

Figure 23. Wing stub box test.

Load-introduction actuator

Stitched graphiteepoxy wing box

Metallic extension box

(a) Test set-up (b) Test results Normalized Tip Load

NASA prediction

Experiment

Test

Analysis

Normalized Tip Displacement

0.0

0.5

1.0

0.5 1.0

39
Figure 25. Test verifies RLV wing box technology.

Figure 24. Test verifies cost-effective composite wing technology.

(a) Test set-up (b) Test results

40
Figure 27. MSFC structural testing of RLV LH2 tank.

Figure 26. LaRC structural test of segment of an RLV intertank.

(a) Test set-up (b) Failed test article

41
Figure 28. Revolutionary airframe configurations.

Figure 29. Next generation structural design and analysis tools.

(a) Blended wing body

(b) Joined wing

(c) Strut braced wing

(a) Next generation model assembly 

and graphical user interface

(b) 3-D stereo viewing of analysis results

42
Figure 30. Low-cost composite materials processing of the future.

Low-cost, 

stitched 

textile

preforms

Magnetic field to 

shape/position material

E-beam 

sources for 

curing

Finished composite 

fuselage barrel 

section

Tooless, RTM processing 

with E-beam curing of 

nanophase resins and textile 

preforms

Tooless, RTM processing 

with E-beam curing of 

nanophase resins and textile 

preforms

Computer 

process 

control

Figure 31. Computationally designed materials and structures.

43
10

100

1000

0.1 1 10 100

Specific

Modulus

(normalized)

Specific strength (normalized)

Aluminum 2219

Baseline material

CFRP Composite

SWNTFRP composite

Single crystal

bulk material (SWNT)

Emerging material

Long-term 

potential

Vehicle

dry

weight

1.0

Aluminum

2219

CFRP

composite

SWNTFRP

composite

55 % reduction

82% reduction

Figure 32. Properties of carbon nanotubes and composite materials.

Figure 33 Systems analysis results for a reusable launch vehicle.

(a) Reusable launch vehicle (b) Systems analysis results

44
Figure 34. Evolution of composite materials NDE technology.

Figure 35. Collaborative engineering design environment.

ISE*

* Intelligent Synthesis Environment (ISE)

Life Cycle

Cost

Flight Loads

Uncertainty

Assessment

Component

Configurations

Structural/Life

Analysis

Emissions

Manufacturing

Fluid

Analysis

Maintenance Materials

NDE Simulations

BENEFITS

Emissions

Noise

Weights

Thrust

Range

Capacity

Speed

DOC

IOC

LCC

ROI

Risk

1970 1990 2010

• Birth of radiography

1900

• Ultrasonics

• Eddy current

• Fluorescent penetrant

• Magnetic particle

• Computer tomography

• Holography 

• Laser ultrasonics

• Shearography

Technology

advances

• Vehicle health monitoring

• Real-time process control

• Fatigue-life sensor

• Telerobotics

• Simulations in design

Year

• Thermal diffusivity

• Magnetoptic imaging

•Microwave

• Contamination monitor

• NDE/I computational simulations

45
REPORT DOCUMENTATION PAGE Form Approved

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1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE

April 2001

3. REPORT TYPE AND DATES COVERED

Technical Memorandum

4. TITLE AND SUBTITLE

An Assessment of the State-of-the-Art in the Design and Manufacturing of

Large Composite Structures for Aerospace Vehicles

5. FUNDING NUMBERS

WU 706-85-20-01

6. AUTHOR(S)

Charles E. Harris, James H. Starnes, Jr., and Mark J. Shuart 

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

NASA Langley Research Center

Hampton, VA 23681-2199

8. PERFORMING ORGANIZATION

REPORT NUMBER

 L-18070

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

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10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA/TM-2001-210844

11. SUPPLEMENTARY NOTES

 

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12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

The results of an assessment of the state-of-the-art in the design and manufacturing of large composite structures

are described. The focus of the assessment is on the use of polymeric matrix composite materials for large

airframe structural components, such as those in commercial and military aircraft and space transportation

vehicles. Applications of composite materials for large commercial transport aircraft, general aviation aircraft,

rotorcraft, military aircraft, and unmanned rocket launch vehicles are reviewed. The results of the assessment of

the state-of-the-art include a summary of lessons learned, examples of current practice, and an assessment of

advanced technologies under development. The results of the assessment conclude with an evaluation of the

future technology challenges associated with applications of composite materials to the primary structures of

commercial transport aircraft and advanced space transportation vehicles.

14. SUBJECT TERMS

composite materials, graphite/epoxy, graphite, carbon, epoxy, thermosets,

15. NUMBER OF PAGES

50

aircraft, rotorcraft, general aviation, rockets, wing, fuselage, empennage, NDE 16. PRICE CODE

A03

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