December 2010 

NASA/TM-2010-216869 

Comparison of Requirements for Composite 

Structures for Aircraft and Space Applications 

Ivatury S. Raju/NESC 

Langley Research Center, Hampton, Virginia 

Kenny B. Elliot 

Langley Research Center, Hampton, Virginia 

Roy W. Hampton 

Owl Analytics, LLC, Cupertino, Clifornia 

Norman F. Knight, Jr. 

General Dynamics Information Technology, Chantilly, Virginia 

Pravin Aggarwal 

Marshall Space Flight Center, Huntsville, Alabama 

Stephen P. Engelstad 

Lockheed-Martin Aeronautics Company, Marietta, Georgia 

James B. Chang 

The Aerospace Corporation, El Segundo, California 
NASA STI Program . . . in Profile

 Since its founding, NASA has been dedicated to 

the advancement of aeronautics and space science. 

The NASA scientific and technical information (STI) 

program plays a key part in helping NASA maintain 

this important role. 

 The NASA STI program operates under the 

auspices of the Agency Chief Information Officer. It 

collects, organizes, provides for archiving, and 

disseminates NASA’s STI. The NASA STI program 

provides access to the NASA Aeronautics and Space 

Database and its public interface, the NASA Technical 

Report Server, thus providing one of the largest 

collections of aeronautical and space science STI in 

the world. Results are published in both non-NASA 

channels and by NASA in the NASA STI Report 

Series, which includes the following report types: 

 TECHNICAL PUBLICATION. Reports of 

completed research or a major significant phase 

of research that present the results of NASA 

programs and include extensive data or 

theoretical analysis. Includes compilations of 

significant scientific and technical data and 

information deemed to be of continuing 

reference value. NASA counterpart of peerreviewed formal professional papers, but having 

less stringent limitations on manuscript length 

and extent of graphic presentations. 

 TECHNICAL MEMORANDUM. Scientific 

and technical findings that are preliminary or of 

specialized interest, e.g., quick release reports, 

working papers, and bibliographies that contain 

minimal annotation. Does not contain extensive 

analysis. 

 CONTRACTOR REPORT. Scientific and 

technical findings by NASA-sponsored 

contractors and grantees. 

 CONFERENCE PUBLICATION. Collected 

papers from scientific and technical 

conferences, symposia, seminars, or other 

meetings sponsored or co-sponsored by NASA. 

 SPECIAL PUBLICATION. Scientific, 

technical, or historical information from NASA 

programs, projects, and missions, often 

concerned with subjects having substantial 

public interest. 

 TECHNICAL TRANSLATION. Englishlanguage translations of foreign scientific and 

technical material pertinent to NASA’s mission. 

 Specialized services also include creating custom 

thesauri, building customized databases, and 

organizing and publishing research results. 

 For more information about the NASA STI 

program, see the following: 

 Access the NASA STI program home page at 

http://www.sti.nasa.gov

 E-mail your question via the Internet to 

help@sti.nasa.gov

 Fax your question to the NASA STI Help Desk 

at 443-757-5803 

 Phone the NASA STI Help Desk at 

443-757-5802 

 Write to: 

 NASA STI Help Desk 

 NASA Center for AeroSpace Information 

 7115 Standard Drive 

 Hanover, MD 21076-1320 
National Aeronautics and 

Space Administration 

Langley Research Center 

Hampton, Virginia 23681-2199 

 

December 2010 

NASA/TM-2010-216869 

Comparison of Requirements for Composite 

Structures for Aircraft and Space Applications 

Ivatury S. Raju/NESC 

Langley Research Center, Hampton, Virginia 

Kenny B. Elliot 

Langley Research Center, Hampton, Virginia 

Roy W. Hampton 

Owl Analytics, LLC, Cupertino, Clifornia 

Norman F. Knight, Jr. 

General Dynamics Information Technology, Chantilly, Virginia 

Pravin Aggarwal 

Marshall Space Flight Center, Huntsville, Alabama 

Stephen P. Engelstad 

Lockheed-Martin Aeronautics Company, Marietta, Georgia 

James B. Chang 

The Aerospace Corporation, El Segundo, California 
Available from: 

NASA Center for AeroSpace Information 

7115 Standard Drive 

Hanover, MD 21076-1320 

443-757-5802 

The use of trademarks or names of manufacturers in the report is for accurate reporting and does not 

constitute an official endorsement, either expressed or implied, of such products or manufacturers by the 

National Aeronautics and Space Administration. 
iii 

Table of Contents 

1.0 Introduction .......................................................................................................................... 1

2.0 Special Features of Composites ........................................................................................... 1

3.0 Comparison of Mission Requirements ................................................................................. 3

4.0 Comparison of Specific Requirements ................................................................................ 4

4.1 Relevant Design Allowables: ........................................................................................... 4

4.2 Development of Relevant Design Allowables: ................................................................ 5

4.3 Factors of Safety: ............................................................................................................. 7

4.4 Structural Integrity: .......................................................................................................... 8

4.5 Damage Tolerance: ........................................................................................................ 10

5.0 Concluding Remarks .......................................................................................................... 11

6.0 References .......................................................................................................................... 12

List of Figures 

Figure 1. Sources of delaminations at geometric and material discontinuities. ........................... 2

Figure 2. Compression after impact strength for AS4/3501-6 composite laminate (NASA 

Langley test data, see Ref. 1). ....................................................................................... 3

Figure 3. Building block test approach. ....................................................................................... 6

Figure 4. Design development test specimens (numbers in parenthesis refer to the number of 

replicates for each different specimen, see Ref. 1). ...................................................... 6

Figure 5. Damage tolerance philosophy currently used by airplanes. (DUL = 1.5 * DLL. The 

structure has to sustain design ultimate loads in the presence of non-detectable 

damage) ....................................................................................................................... 10
Abstract 

In this report, the aircraft and space vehicle requirements for 

composite structures are compared. It is a valuable exercise to 

study composite structural design approaches used in the airframe 

industry and to adopt methodology that is applicable for space 

vehicles. The missions, environments, analysis methods, analysis 

validation approaches, testing programs, build quantities, 

inspection, and maintenance procedures used by the airframe 

industry, in general, are not transferable to spaceflight hardware. 

Therefore, while the application of composite design approaches 

from aircraft and other industries is appealing, many aspects 

cannot be directly utilized. Nevertheless, experiences and research 

for composite aircraft structures may be of use in unexpected 

arenas as space exploration technology develops, and so continued 

technology exchanges are encouraged. 

1.0 Introduction 

Aircraft and space vehicle structures are performance structures. Because of their light weight 

and high strength, composites are potential candidates for these structures [1]. They are widely 

used in secondary structures and are currently making their way into aircraft primary structures 

such as wings and fuselage components (Boeing 787 Dreamliner, Airbus A380 aircrafts, and 

many military aircraft (F-22, F-35)). While space vehicles are behind in the use of composites, 

composites are considered attractive alternatives to metallic structures. When new materials are 

considered for structural applications, one of the first issues to confront would be the 

requirements for these structures and structural components [2]. The purpose of this paper is to 

attempt a comparison of the requirements for composite structures for aircraft and space 

applications. 

The paper is organized as follows. First, a general discussion on special features of composites 

is presented. This is followed by a general comparison of features of aircraft and spacecraft 

missions. Then, the comparison of requirements for design allowables, development of these 

allowables, factors of safety, structural integrity, and damage tolerance is presented. For 

convenience in presentation, spacecraft and space vehicle structures are used interchangeably. 

2.0 Special Features of Composites 

Compared to widely used metallic materials, composites differ significantly in many 

characteristics. Some of these special features of composites are presented here. 

 Composites demonstrate high strength-to-weight ratios. 

 Composites exhibit higher dimensional stability, corrosion resistance, and thermal cycling 

capability. 

 Composites provide numerous options for structural tailoring for known loading conditions. 

 Constituent materials in the composite strongly influence composite structure’s stiffness, strength, 

and durability. 
2 

 Composites have practically no ductility. Most composites are brittle and have a nearly linear 

stress-strain law to failure. 

 Composites are non-isotropic and are inhomogeneous. 

 The mechanical properties of composites have a strong link to the manufacturing processes. 

Therefore, one needs to pay special attention to the manufacturing processes, realize that there 

can be wide variability in the allowables data, and realize the need for a careful building-block 

test program to develop the design allowable data. 

 Composites are highly sensitive to notches. Stress concentration factors higher than ten are 

possible for some notches, and the stress gradients near the notches are very high. 

 In composites, large in-plane stress gradients develop near geometric and material discontinuities. 

In-plane stress gradients lead to out-of-plane or interlaminar stresses. Because composites have 

low through-the-thickness strength, delaminations can initiate and grow. Delaminations degrade 

the strength of the composite laminate and can lead to failure. Figure 1 shows some typical 

configurations susceptible to interlaminar stresses where delaminations initiate because of 

geometric and material discontinuities. 

 Composites degrade due to moisture, humidity, and heat. The coefficient of thermal expansion 

along the fiber direction is near zero or a small negative number compared to the other two 

transverse directions. 

 Composites are sensitive to accidental damages. There is a significant loss of strength due to 

low-velocity impacts. Figure 2 shows the dramatic reduction in compression after impact 

strength as a function of impact energy for a commonly used composite system. In thick 

composites, the damage due to these impacts is in the interior of the laminate and is either not 

visible on the outer surfaces of the laminate or is barely visible. 

 Composites have low sensitivity to fatigue for in-plane loading. 

Curved free edge

Straight free edge

Internal ply drop

Solid-sandwich transition

External ply drop

Corner

Skin stiffener

interaction

Interlaminer

Stresses

z

xz

Figure 1. Sources of delaminations at geometric and material discontinuities. 
3 

Load Direction

48 ply stitched laminate

[+45/0/-45/90]6s

Comp.

Stress,

MPa

Impact Energy, J

700

600

500

400

300

200

100

0

0 20 40 60 80 100 120

Figure 2. Compression after impact strength for AS4/3501-6 composite laminate 

(NASA Langley test data, see Ref. 1). 

3.0 Comparison of Mission Requirements 

The commercial and military aircraft industries base their requirements on the performance 

required out of the appropriate certifying authority. Commercial transport aircraft usually have a 

design life of 75,000 cycles (take-offs and landings) over 30 years. Military transport aircraft 

may have similar operational life. Many copies, perhaps thousands of copies, of commercial 

aircraft are built, while tens to a few hundred copies of military aircraft are built. The 

manufacturing processes are highly specialized by suppliers through extensive testing of 

company designs of materials, coupons, joints, components, and structural elements. As a result 

of this comprehensive testing, an extensive database of material properties and structural 

performance data is created for each system. The aircrafts are certified to meet applicable 

requirements based on detailed reviews that ensure that the aircraft meet safety and performance 

criteria for the fleet mission environments and loadings, which are usually well defined [3-5] and 

can be incorporated in the certification test programs. In addition, the aircraft applications rely 

on extensive development programs that continue into the operational and maintenance regimes 

of the mission lifetime. These programs ensure continued airworthiness through periodic 

inspections that may include depot re-work using certified repair and/or structural upgrades as 

needed throughout the mission lifetime (see Refs. 3-5). 

In contrast, space vehicle missions typically have no or very limited opportunities to inspect, 

repair, or modify the structure once they lift off from the launch pad. This is a major and 

fundamental difference between the aircraft approach and the spacecraft approach to composite 

structural design methodology. While the aircraft in-service inspection/repair approach could be 

adopted for any space vehicle that returns periodically to Earth, it would require the space 

vehicle program to develop and certify capabilities similar to those in use for aircraft while 

addressing the more challenging aspects of missions that operate in space environments. There 

are some examples within the NASA Space Shuttle Program where refurbishment of vehicle 

components is performed, but this practice is not common because of the risks of damage to 

expensive space-qualified vehicles. Instead, spacecraft structures are typically designed for a 

single launch and are certified to perform their entire mission without interim inspections or 

repairs to reduce operational costs. 
4 

Space vehicles experience unique environments and loadings, and there will be few opportunities 

to validate the vehicle’s structural performance under the mission conditions. Therefore, the 

space vehicle structural design must be more robust than a design obtained using comparable 

aircraft structural design procedures. Furthermore, at most, only a few copies of the space 

vehicle structure will probably be built – perhaps more launch vehicles than crew modules and 

even fewer lunar and interplanetary systems. Thus while some aspects of the aircraft industry’s 

composite design best practices (such as those for ply drops, stacking sequences, fiber/resin 

combinations, etc.) may benefit composite spacecraft structural design, most of their best 

practices may not be directly applicable to spacecraft composite structures. Nevertheless, a 

careful comparison of requirements for aircraft and spacecraft composite structures is very 

useful. 

4.0 Comparison of Specific Requirements 

Five major requirements for both aircraft and spacecraft structures are discussed and compared. 

The requirements are focused on the standard practices as they are followed today for composite 

structures. 

4.1 Relevant Design Allowables: 

Strength allowables are consistent between the aircraft and spacecraft applications. Both aircraft 

and spacecraft standards require that strength allowables should properly account for the effects 

of environment (e.g., temperature, humidity, vacuum, chemicals), but beyond that they do not 

specify which properties (e.g., notched or unnotched strength) should be used for design 

allowables. 

Strength allowable values are generated typically using uniaxial tests performed at different 

environmental conditions (e.g., different values of temperature, humidity). These values are 

usually from load-to-failure tests with maximum load (‘stress’) and strain-at-failure recorded. 

For composite systems exhibiting a nonlinear response, repeated loading-unloading cycles with 

increasing load level may provide insight into the overall material response (i.e., nonlinear elastic 

or damage accumulation). 

The definition of statistical sampling and data reduction methods for what constitutes an A- or Bbasis for strength allowables is common among the aircraft and spacecraft design standards. 

Also common is the general requirement that B-basis design allowables will be used for a 

redundant or fail-safe structure and A-basis design allowables will be used for a single load path 

structure. A redundant or fail safe structure is defined as follows: A structural design criterion 

in which it must be shown that the structure remaining, after failure of any single structural 

member, can withstand the resulting redistributed internal limit loads without failure. The ability 

to sustain a failure and retain the capability to safely terminate or control the operation. Thus, 

redundant fail-safe structural designs are desirable. In fact, space vehicle fail-safe structural 

designs using B-basis strength allowables are used by the current NASA [6-9] and European 

Space Agency (ESA) requirements [10-12]. 
5 

Joint designs, whether bolted or bonded or both, are complex structural components. Load paths 

and reactions are dependent on many factors associated with the joint itself. Adding redundancy 

in the joint system does add complexity with respect to load path and other geometric limitation 

along with additional mass. Incorporating redundancy in a specific joint design to give fail-safe 

capability and migrate from an A-basis allowable requirement to a B-basis requirement may be 

as difficult or costly to achieve as establishing an A-basis allowable for the specific joint 

configuration. The interaction of other joint factors or fitting factors also needs to be reviewed in 

the designs. The airframe industry typically establishes statistically relevant performance data 

for each major joint configuration through comprehensive test programs. The industry also 

maintains heritage configurations and databases for each design. While it may be desired, it is 

unlikely that such an approach will be used for composite spacecraft structures. 

Probabilistic or non-deterministic design approaches are being proposed and pose an additional 

paradox. Without statistically relevant data, mean values, standard deviations, and distribution 

forms for design allowables will have to be assumed and this would result in uncertainties in the 

analytical predictions. These predictions can be used to determine the relative sensitivity of 

selected response variables within assumed ranges and distributions. This paradox will most 

likely lead to specific focused building-block test programs for key design-driving parameters. 

4.2 Development of Relevant Design Allowables: 

A building-block approach starts by studying simple basic configurations and then builds in 

complexities associated with the design in a systematic and progressive manner [1, 13, 14]. A 

basic building-block program can have different objectives, goals, and scope depending on the 

application and the program. In most cases, the objective of the building-block approach is to 

establish statistically relevant design allowables through an extensive test program moving from 

coupon-level specimens to component-level tests and even to full-scale tests. In other cases, the 

objective of the building-block approach is to define the composite structural strength and to 

identify and characterize failure modes using a limited number of specimens. In yet other cases, 

the building-block information must be sufficient to guide the definition of a meaningful fullscale/component test – including consideration for appropriate type/size/amount of damage [1, 

13]. Figure 3 and 4 demonstrate how this is practiced currently for aircraft structures. For 

composite structures design for aircraft and spacecraft, multiple types of building-block 

programs may be required to develop robust composite designs that fully exploit the potential of 

each composite material system. 
6 

Building Block Integration.

Verification of Design 

Data and Methodology

Development of Design 

Data

Number of Specimens

Chronological SequenceSpecimen Complexity

Certification Methodology (Mil-Hbk.-17)

Structural Levels of Testing & Analysis

Static/

Fatigue

Material Selection and Qualifications Coupons

Design Allowables Coupons

Structural Elements

Sub-components

Components

Full Scale 

Article

Analysis

Figure 3. Building block test approach. 

Buckling

Elements

Improved Concepts

Darted Spar

Caps

Spar Cap Longitudinal

Shear Panel

Splice Specimen

Upper Cover

Compression Panel

Flap Support

Skin Ply Drop-off

Bulkhead Shear

Clip Specimen

Substructure

Shear Specimen

Postbuckling

Specimen

Rib Clip Pull-off

Specimen

Stringer Runout

Panel

Compression

Tension

Repair Panel

Hi-Load Panels

(1) (7)

(7)

(8)

(5)

(5)

(13)

(1)

(1)

(4)

(2)

(2)

(2)

Figure 4. Design development test specimens (numbers in parenthesis refer to the number of 

replicates for each different specimen, see Ref. 1). 
7 

While the building-block approach has all the aforementioned advantages, the approach has 

some disadvantages. A “properly executed” building-block approach as used by the airframe 

industry may not be practical for space vehicle designs from a cost-schedule perspective. The 

airframe industry’s building-block approach usually involves testing of hundreds of specimens. 

For the airframe industry, significant R&D resources are put into understanding their specific 

structural systems and generating the necessary databases sufficient for design and certification. 

The resource cost of this effort can be amortized across the fleet of vehicles produced. 

Unfortunately, the space vehicle industry only produces a limited number of products making the 

institution of an extensive building-block test program impractical from a cost perspective. 

Often, spacecraft structural systems are unique or application dependent. Design-allowable data 

from a building-block approach may not be available early enough in the design cycle to affect 

the robustness or efficiency of the spacecraft design. A generalized building-block test program 

may produce design data and expose structural weaknesses. However, unless the test program 

performs testing of all material, configurations, and loading complexities, some structural 

weaknesses may not be identified before a flight mission or before a catastrophic failure at worst, 

or before large-scale damage testing results are obtained. Furthermore, this building-block 

approach could potentially lead to major program redesign/development cost increases and 

schedule delays. Finally, due to the extreme complexity or impossibility of a priori knowledge 

and simulation capability of the spacecraft operational environment on Earth, the development of 

a fully encompassing building-block program may not be achievable. 

Clearly, traditional building-block approaches used by the aircraft industry are not directly 

applicable for space vehicles. For successful composite structures and structural components, an 

appropriate building-block approach needs to be designed and executed to exploit all the 

benefits, avoid the pit-falls, and mitigate known risks. 

4.3 Factors of Safety: 

The historic factor of safety (FOS) for aircraft structures is 1.5 [15], while for space vehicle 

structures it is 1.4 [16]. These FOS are used generally for most of the acreage areas of the 

structures exhibiting near uniform stress states, and specific parts have higher factors. There is a 

significant difference in the NASA approach to areas of discontinuities and this will be discussed 

next. 

In the NASA-STD-5001A [16], “discontinuities are defined as an interruption in the physical 

structure or configuration of the part. Delaminations, debonds, and dropping of plies are all 

assumed to be discontinuities.” In addition, the term “discontinuity area” is defined in a recent 

structural design verification requirements document [6] as: “A local region of a composite or 

non-metallic structure consisting of built-up plies, chopped fiber or reinforced regions around 

fittings, joints or interfaces where the stress state and load distribution within the region may be 

difficult to characterize. A region is considered a discontinuity area until uniform section 

properties in the structure can be considered in the structural analysis. Bonded joints are 

considered discontinuities.” 

In the preliminary design phase of aircraft or spacecraft, predictive analysis tools and techniques 

for local discontinuities in composite structures are difficult to incorporate (and validate) in the 

design process. At the present time, these analysis tools and modeling practices are not 
8 

sufficiently robust and general to provide adequate fidelity of the stress state in areas of 

discontinuities. Through-the-thickness or out-of-plane stresses may not be correctly 

characterized and understood, leading to incorrect characterization of these discontinuities. In 

part, the use of a discontinuity area FOS is applied to address the modeling and analysis 

uncertainties early in the design. 

The discontinuity area FOS approach is suggested and defined for strength requirements in 

NASA-STD-5001A [16] and in the Federal Aviation Administration (FAA) Commercial Space 

Transportation guidance document for reusable launch and reentry vehicles [17]. FOS listed in 

NASA-STD-5001A [16] are stated as “minimum” values, and some structural members or 

systems may be required to meet other more stringent and restrictive performance requirements. 

Specific material systems and design configurations need to be tested to determine statistically 

relevant allowable data and characterization of failure modes. Such testing typically is 

performed as an integral part of a larger building-block test program for the overall design and is 

representative of current best practices. 

A possible misconception within the spacecraft technical community regarding FOS also needs 

to be addressed here. The misconception is that the FOS covers uncertainty in loads and/or 

material properties. The uncertainty in the loads is not covered by the FOS and needs to be 

accounted for by the use of appropriate load uncertainty factors (LUF). These LUFs depend on 

the environment in which the structure operates. In contrast, in the aircraft technical community, 

most references state that the FOS is for loads greater than those assumed (overshoots), and for 

uncertainties in design and fabrication (e.g., see W. G. Heath in Ref. 15, pp. 1526). In a similar 

manner, the uncertainty in material properties is not covered by the FOS. The uncertainties in 

the material properties need to be accounted for by the use of A-basis and B-basis properties for 

the operating environments. These FOS account for uncertainties and variations that cannot be 

analyzed or accounted for in a rational manner, including modeling uncertainties, analysis 

assumptions, defects in material and fabrication issues, and unknown/unexpected aspects of 

complex structures. 

In summary, a comprehensive building-block approach to developing statistically relevant data is 

the correct technical approach for space vehicle composite structures. However, the practical 

application of such an approach may not be cost effective or feasible. On the other hand, when 

the building-block approach for composite spacecraft structures is attempted but not “properly 

executed” due to cost and schedule constraints, there is a risk of structural failures. A 

methodology that establishes a balance between the deterministic FOS approach and a sufficient 

building-block program is needed. 

4.4 Structural Integrity: 

Generally, structural integrity is evaluated based on the following three factors: 

 The maximum estimated structural response. The estimated or anticipated structural response is 

usually determined through established analysis and/or test methods. 

 The uncertainty of the estimate. The uncertainty in the estimated structural response is 

established through best practice. NASA-STD-5001A [16] represents NASA’s view of the 

current “best practice” for establishing FOS for spaceflight hardware [18] and damage tolerance 

for composite structures is established using MSFC-RQMT-3479 [9]. However, these FOS 
9 

values may be tailored by Programs provided an adequate engineering risk assessment is 

performed and approval is obtained from the appropriate Technical Authority. 

 The capability of the structure. The structural capability is generally established by material 

capabilities. 

This paradigm for structural integrity is well accepted in the metallic structures community 

where analysis methods and material capabilities are well understood. However, in composite 

structural systems, the material capability is heavily influenced by the geometry, laminate 

stacking sequences, material architecture (uni-directional tape, fabric, textiles), manufacturing 

processes, and bulk material properties. These influencing factors lead to additional uncertainty 

in the structural integrity of composite systems. This uncertainty primarily pertains to structural 

regions, such as joints, variable thickness regions, stiffeners, or cut-outs, where the stress state or 

load path is difficult to assess using current analytical tools and methodologies and where the 

stress state is strongly dependent on the material definition, material properties, and fabrication. 

The airframe and space vehicle industries handle this additional uncertainty in different ways. 

The airframe industry maintains a consistent single FOS that covers general uncertainty (1.5) 

combined with a reliance on an extensive building-block test program to determine the structural 

capability. Such a program could involve hundreds of tests for a structural application using a 

specific material, fabrication processes, configurations, loadings, and damage in order to 

establish a statistical basis [14]. In contrast, the current NASA-STD-5001A [16] practice, as 

well as the FAA guidance document for reusable launch and reentry vehicles [17], relies on a 

dual FOS approach: one FOS is applicable to areas of the structure where the response is well 

understood (for example, 1.4 for uniform areas); and another FOS for areas of the structure that 

are not well understood (for example, 2.0 for discontinuity areas). However, the material 

properties are left to the individual programs to establish. NASA-STD-5001A states: 

“designs are assumed to use materials having mechanical properties that are well 

characterized for the intended service environments and all design conditions ... 

Material allowables shall be chosen to minimize the probability of structural failure 

due to material variability ... Allowables shall be based on sufficient material tests to 

establish values on a statistical basis” 

Again, these requirements are well understood and applied for metallic structural systems. 

However, they become problematic for composite systems because of the influencing factors 

previously mentioned. Hence, the airframe industry approach to developing material allowable 

data using a building-block approach appears attractive, but for spacecraft, the development of 

statistically relevant data is not always feasible. To further complicate the issue, in using the 

deterministic approach, there is evidence [4, 19] that a factor of 2.0 may not be conservative 

depending on the severity of manufacturing flaws, design features such as cutouts and holes, 

and/or impact damage. Higher FOS values may be warranted in some cases depending on the 

adequacy of the test database, the conservatism of the design allowable, inclusion of knockdown 

factors, and the method of design verification. 

In situations where structurally similar design and fabrication processes have demonstrated 

reliable strength in well-defined loading/environment conditions, Section 4.4 of NASA-STD5001A [16] has a provision to request an alternative approach by waiver when a substantial case 

can be made. This waiver would need to be approved by the responsible program authorities and 

technical authorities. Finally, in the event a structural weakness is discovered late in the 
10 

structural development program or launch planning and provided testing establishes that the 

weakness is an acceptable risk, then this waiver process can be pursued. Note that such an 

approach is feasible only when the component was originally designed with the specified FOS. 

4.5 Damage Tolerance: 

For commercial aircraft structures, damage tolerance is achieved through a methodology 

developed by the FAA [20] and is shown in Figure 5. Military aircraft structures use a similar 

method that is stated in the Department of Defense Joint Service Specification Guide [3]. In this 

figure, the design load level is plotted against increasing damage severity. On the abscissa, the 

allowable damage limit (ADL) and critical damage threshold (CDT) are plotted. The ADL 

corresponds to the limit of the Category-1 damage – “Barely Visible Impact Damage” (BVID) 

and “Allowed Manufacturing Damage”. The CDT corresponds to the damage limit when the 

residual strength of the structure falls below the design limit load (DLL). Note that aircraft 

composite structures are required to carry the design ultimate load (DUL) with any nondetectable damage. For reduced load capability with “Category 2 or 3” damage conditions, the 

structure is not permitted to have reduced strength below DLL. In addition, the flight mission is 

adjusted such that opportunities for certified inspection that detects the damage and certified 

repairs that restores the structure to ultimate load capability are provided. (In fact, in 

determining the CDT, one missed inspection cycle is built in.) Category-4 damage, such as 

liberation of rotor blade and subsequent impact with the aircraft structure, bird strikes, lightning 

damage, hail damage, etc., would cause a loss of limit load capability. However, in these 

situations, the pilot is aware of the damage and will execute a safe landing by adjusting the flight 

envelope to reduce flight loads. 

Ultimate

Limit

Continued safe

flight

Design Load Level

DUL

DLL

Allowable Damage 

Level

Critical Damage 

Threshold

Category 1 Damage: 

BVID, Allowable Mfg. damage

Category 2 Damage: 

VID, Damage requiring repair per

normal inspection process

Category 3 Damage: 

Obvious damage requiring repair 

after it is found within a few flights after occurrence

Category 4 Damage: Discrete source 

damage obvious to flight crew, requiring

repair after flight

Category 5 Damage: Anomalous 

Damage not covered in design 

but known operations

Requiring immediate repair

Increasing Damage Severity

Figure 5. Damage tolerance philosophy currently used by airplanes. (DUL = 1.5 * DLL. The 

structure has to sustain design ultimate loads in the presence of non-detectable damage) 

The Category-1 damage indicated in Figure 5 that corresponds to space applications would likely 

include damage due to fabrication, assembly, ground handling, transportation, and any other 
11 

sources of damage occurring prior to launch. According to the aircraft criteria, this damage must 

not decrease strength below ultimate strength, which is the same as the existing space vehicle 

requirement [9]. In addition, space missions must consider damage due to all other sources that 

may occur during a spacecraft mission that includes: liftoff, ascent, time spent in orbit from risk 

of impact by micrometeoroids or other space debris, the return to Earth, and subsequent 

missions. Additional unique damage risks may exist during long-duration lunar and 

interplanetary missions. These aspects are addressed in the damage threat assessments and the 

strength with damage requirements of Ref. 9. Damage from impact threats after launch and their 

consequences for space missions are uniquely different from those of the airframe industry. 

No space mission conditions are known at this time that permit the use of the aircraft Category 2, 

3, or 4 damage indicated in Figure 5 because inspection and repair capability are not feasible 

once the space vehicle is launched. 

The NASA fracture control damage tolerance requirements for composites specified in Section 

5.3.2.6 of Ref. 9 do not permit damage growth. This requirement was instituted to control risk of 

fracture in composites because technology for predicting damage initiation and accumulation for 

composites is not as reliable as it is for metals. This requirement is similar to what is currently 

practiced by the aircraft industry. A new fracture control standard, NASA-STD-5019A [21] that 

is applicable to both metallic and composite materials, is developed recently. It is also expected 

that this standards development activity will offer opportunity for discussion of different 

approaches to composite spacecraft structural design requirements. 

In summary, the key points related to damage tolerance requirements are: 

 Aircraft requirements are similar to space vehicle requirements to sustain ultimate loads with 

manufacturing and other defects/damage before launch. 

 Aircraft requirements permit some strength reduction below ultimate for Category 2 or 3 damage, 

but this is different from usual space applications because: 

o Aircraft practice specifies periodic depot inspections and repairs are performed to ensure 

continued airworthiness, and 

o Aircraft practice requires inspections and repairs to restore the structure to ultimate load 

capacity within a short fraction of the mission life between the required depot work. 

 Space vehicle programs usually do not include development of certified inspection and repairs of 

structure to restore it to ultimate strength capacity after launch. In addition, usually there are no 

opportunities for these activities once the spacecraft has been launched. 

5.0 Concluding Remarks 

Composites are being widely used in aircraft structures because of their high strength-to-weight 

ratios. Their use in space applications is increasing. The requirements for composite structures 

for aircraft and space applications are compared and discussed. The requirements for design 

allowables, strength allowables, factors of safety, structural integrity, and damage tolerance are 

compared. To develop design allowables, the aircraft industry employs a building-block 

approach involving testing of hundreds of specimens. As space vehicles are only produced in 

extremely limited numbers, traditional building-block approaches such as that used by the 

aircraft industry are impractical from a cost and schedule perspective. Requirements for strength 

allowables, on the other hand, are consistent between the aircraft and spacecraft applications. 
12 

The requirements on factors of safety used by the spacecraft structures are more stringent than 

those for the aircraft structures. To ensure structural integrity, a comprehensive building-block 

approach to developing statistically relevant data is the approach used by both aircraft and 

spacecraft composite structures. For space applications, a methodology that establishes a 

balance between factor of safety and building-block testing that is tailored for specific 

applications is needed. To ensure damage tolerance, aircraft and spacecraft follow similar 

requirements for composite structures. Aircraft requirements are similar to space vehicle 

requirements to sustain ultimate loads with manufacturing and other defects/damage before 

launch. Aircraft practice requires periodic inspections and restoration of the structure to design 

ultimate load capacity many times during its mission life. Most space vehicles do not have such 

opportunities after launch and lift-off. Thus, the damage tolerance requirements for aircraft and 

space vehicles are much different with only similarities up to the allowable damage limit. 

Because of mission and operations differences, the requirements for composite structures for 

aircraft applications are not directly applicable as requirements for space vehicle applications. 

Space vehicles experience unique environments and loadings and there will be fewer, if any, 

opportunities to inspect and repair the vehicle’s structure and validate its structural performance 

under mission conditions. Therefore, the space vehicle composite structural design must be 

more robust than the aircraft composite structural design. Nevertheless, the aircraft industry’s 

experience with composite structures provides a knowledge base for consideration and possible 

adaptability to spacecraft composite structural design. 

6.0 References 

1. Harris, C. E.; Starnes, J. H., Jr.; and Shuart, M. J.: “Design and manufacturing of aerospace 

composite structures, state-of-the-art assessment.” Journal of Aircaft, Vol. 39, No. 4, July–

August 2002, pp. 545560. 

2. Silverman, E. M. (Compiler): Composite Spacecraft Structures Design Guide, Parts 1 and 2. 

NASA CR-4708, March 1996. 

3. Department of Defense Joint Service Specification Guide Aircraft Structures. JSSG-2006, 30 

October 1998. 

4. MIL-HDBK-17-3F, Department of Defense Handbook, Composite Materials Handbook, 

Volume 3, Polymer Matrix Composites Materials Usage, Design, and Analysis, Chapter 7, 

Damage Resistance, Durability, and Damage Tolerance. 17 June 2002. 

5. CMH-17-3G, Working Draft Rev. G, Composite Materials Handbook, Volume 3, Polymer 

Matrix Composites Materials Usage, Design, and Analysis. Under review, publication 

anticipated in 2011. 

6. NASA CxP 70135, Constellation Program Structural Design and Verification Requirements. 

August 25, 2008. 

7. NASA-STD-5019, Fracture Control Requirements for Spaceflight Hardware. July 2008. 

8. NASA-HDBK-5010, Fracture Control Implementation Handbook for Payloads, Experiments, 

and Similar Hardware. May 2005. 

9. MSFC-RQMT-3479, EM 20, Multiprogram/project common-use document, Fracture Control 

Requirements for Composite and Bonded Vehicle and Payload Structures. May 2006. 
13 

10. ECSS-E-ST-32C, Rev. 1, Space Engineering – Structural General Requirements. November 

2008. 

11. ECSS-E-32-10C-Rev. 1, Space Engineering - Structural Factors of Safety for Space Flight 

Hardware. March 2009. 

12. ECSS-E-32-01C, Rev. 1, Space Engineering – Fracture Control. March 2009. 

13. Rouse, M.; Jegley, D. C.; McGowan, D. M.; Bush, H. G.; and Waters, W. A.: “Utilization of 

building-block approach in structural mechanics research.” AIAA Paper No. 2005-1874, 

April 2005. 

14. Mohaghegh, M.: “Validation and certification of aircraft structures.” AIAA Paper No. 2005-

2162, 2005. 

15. Meyers-Jens, R. J. (Compiler): Factors of Safety: Historical Development, State of the Art 

and Future Outlook. AGARD Report No. 661, 1977. 

16. NASA-STD-5001A, Structural Design and Test Factors of Safety for Spaceflight Hardware. 

May 2008. 

17. FAA Commercial Space Transportation, Guide to Verifying Safety-Critical Structures for 

Reusable Launch and Reentry Vehicles. November 2005. 

18. Raju, I. S.; Murthy, P. L. N.; Patel, N. R.; Bonacuse, P. J.; Elliott, K. B.; Gordon, S. A.; 

Gyekenyesi, J. P.; Daso, E. O.; Aggarwal, P.; and Tillman, R. F.: “Best practices for reliable 

and robust spacecraft structures.” AIAA Paper No. 2007-2270, 2007. 

19. Wagenen, R. V.: A Guide to Structural Factors for Advanced Composites Used on 

Spacecraft. NASA CR-186010, August 1989. 

20. FAA Advisory Circular No. 20-107B, Composite Aircraft Structure. September 8, 2009. 

21. NASA-STD-5019A, Fracture Control Requirements for Spaceflight Hardware. October 

2010. 
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

2. REPORT TYPE 

Technical Memorandum

 4. TITLE AND SUBTITLE

Comparison of Requirements for Composite Structures for Aircraft and 

Space Applications

5a. CONTRACT NUMBER

 6. AUTHOR(S)

Raju, Ivatury S.; Elliot, Kenny B.; Hampton, Roy W.; 

Knight, Norman F., Jr.; Aggarwal, Pravin; Engelstad, Stephen P.; 

Chang, James B.

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

NASA Langley Research Center

Hampton, VA 23681-2199

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

National Aeronautics and Space Administration

Washington, DC 20546-0001

 8. PERFORMING ORGANIZATION

 REPORT NUMBER

L-19948

10. SPONSOR/MONITOR'S ACRONYM(S)

NASA

13. SUPPLEMENTARY NOTES

Plenary Lecture presented at the International Conference on Composites for 21st Century: Current and Future Trends 

January 4-7, 2011, Bangalore, India

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Category 39-Structural Mechanics

Availability: NASA CASI (443) 757-5802

19a. NAME OF RESPONSIBLE PERSON

STI Help Desk (email: help@sti.nasa.gov)

14. ABSTRACT

In this report, the aircraft and space vehicle requirements for composite structures are compared. It is a valuable exercise to

study composite structural design approaches used in the airframe industry and to adopt methodology that is applicable for 

space vehicles. The missions, environments, analysis methods, analysis validation approaches, testing programs, build 

quantities, inspection, and maintenance procedures used by the airframe industry, in general, are not transferable to spaceflight 

hardware. Therefore, while the application of composite design approaches from aircraft and other industries is appealing, 

many aspects cannot be directly utilized. Nevertheless, experiences and research for composite aircraft structures may be of 

use in unexpected arenas as space exploration technology develops, and so continued technology exchanges are encouraged.

15. SUBJECT TERMS

Composite; Delamination; Design allowables; Material discontinuities; Test specimens

18. NUMBER

 OF 

 PAGES

19

19b. TELEPHONE NUMBER (Include area code)

(443) 757-5802

a. REPORT

U

c. THIS PAGE

U

b. ABSTRACT

U

17. LIMITATION OF 

 ABSTRACT

UU

Prescribed by ANSI Std. Z39.18

Standard Form 298 (Rev. 8-98)

3. DATES COVERED (From - To)

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

869021.01.07.01.01 

11. SPONSOR/MONITOR'S REPORT

 NUMBER(S)

NASA/TM-2010-216869

16. SECURITY CLASSIFICATION OF:

The public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, 

gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this 

collection of information, including suggestions for reducing this burden, to Department of Defense, Washington Headquarters Services, Directorate for Information Operations and 

Reports (0704-0188), 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302. Respondents should be aware that notwithstanding any other provision of law, no person 

shall be subject to any penalty for failing to comply with a collection of information if it does not display a currently valid OMB control number.

PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS.

1. REPORT DATE (DD-MM-YYYY)

01- 12 - 2010