November 2011 

NASA/TM-2011-217185 

NESC-RP-06-019 

 

Composite Crew Module: Primary Structure 

Michael T. Kirsch/NESC 

Langley Research Center, Hampton, Virginia 
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November 2011 

NASA/TM-2011-217185 

NESC-RP-06-019 

 

Composite Crew Module: Primary Structure 

Michael T. Kirsch/NESC 

Langley Research Center, Hampton, Virginia 
Available from: 

NASA Center for AeroSpace Information 

7115 Standard Drive 

Hanover, MD 21076-1320 

443-757-5802 

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Composite Crew Module: Primary Structure 

May 5, 2011 
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Approval and Document Revision History 

NOTE: This document was approved at the May 5, 2011, NRB. This document was 

submitted to the NESC Director on August 22, 2011, for configuration control.

 Approved: Original Signature on File 8/23/11 

 NESC Director Date 

Version Description of Revision Office of Primary 

Responsibility 

Effective

Date

1.0 Initial Release Mr. Michael Kirsch, NESC 

Principal Engineer, LaRC 

5/5/11
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Table of Contents 

Technical Assessment Report 

1.0 Notification and Authorization ........................................................................................ 6

2.0 Signature Page ................................................................................................................... 7

3.0 Team List ........................................................................................................................... 8

3.1 Acknowledgements ............................................................................................................. 8 

4.0 Executive Summary .......................................................................................................... 9

5.0 Background ..................................................................................................................... 10

6.0 CCM Project .................................................................................................................... 10

6.1 Project Objectives and Success Criteria ........................................................................... 10 

6.2 Major Project Milestones .................................................................................................. 11 

6.3 Project Organization ......................................................................................................... 12

6.4 Team Communications and Decision Making .................................................................. 12 

6.5 Independent Review Process ............................................................................................ 14 

6.6 CCM Configuration Management .................................................................................... 15 

6.6.1 Data Management System ................................................................................................ 15

6.6.2 Numbering ........................................................................................................................ 15 

6.6.3 Revisioning ....................................................................................................................... 16 

6.6.4 Lifecycle States ................................................................................................................. 16 

6.6.5 Electronic Approval .......................................................................................................... 17 

6.6.6 Change Management ........................................................................................................ 17 

6.7 Cost and Schedule Management Approach ...................................................................... 17 

6.7.1 Schedule Management ...................................................................................................... 17 

6.7.2 Cost Management ............................................................................................................. 18 

6.8 Project Cost Analysis ........................................................................................................ 19 

6.8.1 Labor Analysis .................................................................................................................. 20 

6.8.2 Project Implementation Challenges .................................................................................. 22 

7.0 Design Overview.............................................................................................................. 22

8.0 Materials and Processes Overview ................................................................................ 28

9.0 Analysis Overview ........................................................................................................... 29

10.0 Manufacturing Overview ............................................................................................... 31

11.0 NDE Overview ................................................................................................................. 34

12.0 Test Overview .................................................................................................................. 38
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13.0 Orion Comparative Study .............................................................................................. 39

13.1 Mass Differences between CCM and Orion ..................................................................... 40 

13.2 Manufacturability Differences between CCM and Orion ................................................. 40 

13.3 Other Technology Gaps .................................................................................................... 41 

13.3.1 Damage Tolerance for Leak Integrity ............................................................................... 42 

13.3.2 Damage Tolerance as Related to Structural Integrity ....................................................... 42 

13.3.3 Inspection of CCM Pressure Vessel ................................................................................. 42 

13.3.4 Splice Joint Feasibility ...................................................................................................... 43 

13.3.5 Repair Techniques for Pressure Containment and Structural Integrity ............................ 43 

13.3.6 Verification and Certification ........................................................................................... 43 

13.3.7 Material System Solution .................................................................................................. 43 

13.3.8 Penetrations to the Pressure Shell – Maintaining Leak Integrity ...................................... 43 

14.0 Findings and NESC Recommendations ........................................................................ 44

14.1 Findings............................................................................................................................. 44 

14.2 NESC Recommendations.................................................................................................. 44 

15.0 Alternate Viewpoints ...................................................................................................... 45

16.0 Other Deliverables .......................................................................................................... 45

17.0 Lessons Learned .............................................................................................................. 45

18.0 Definition of Terms ......................................................................................................... 45

19.0 Acronyms List ................................................................................................................. 46

20.0 References ........................................................................................................................ 47

21.0 Appendices ....................................................................................................................... 47

List of Figures

Figure 6.8-1. CCM Total Costs: $29.8M (FY07–FY10) ............................................................ 19

Figure 6.8-2. CCM Costs by Phase: Total $29.8M .................................................................... 20

Figure 6.8-3. CCM Total Workforce – Full-Time Equivalent (FTE) and WYE ........................ 21

Figure 6.8-4. CCM Total Workforce – by Project Phase ........................................................... 21

Figure 7.0-1. NESC CCM .......................................................................................................... 23

Figure 7.0-2. CCM Upper Shell ................................................................................................. 24

Figure 7.0-3. CCM Lower Shell ................................................................................................. 24

Figure 7.0-4. Final CCM Assembly ........................................................................................... 25

Figure 7.0-5. Illustration of Backbone Structure within Lower Shell ........................................ 26

Figure 7.0-6. Backbone Structure in Lower Shell ...................................................................... 26

Figure 7.0-7. Lobed Floor on Lower Shell ................................................................................. 27

Figure 7.0-8. 3D Woven Preform Test Element ......................................................................... 27

Figure 9.0-1. CCM Loads “Freebody” ....................................................................................... 30
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Figure 10.0-1. CCM Test Article Manufacturing Flow Chart ...................................................... 33

Figure 11.0-1. CCM Primary NDE Methods by Component Type .............................................. 35

Figure 12.0-1. CCM in Upside Down Configuration within the Self Reacting Load Frame in 

LaRC’s COLTS .................................................................................................... 39

List of Tables

Table 6.2-1. Itemized List of Reviews ...................................................................................... 11

Table 6.6-1. CCM Document (Object) Types ........................................................................... 16

Table 6.6-2. Lifecycle State Access Rules ................................................................................ 17

Table 9.0-1. Margins of Safety Summary ................................................................................. 31

Table 11.0-1. NDE Methods for Common Composite Defect Types ......................................... 36

Table 11.0-2. NDE Methods Identified by Hardware / Component Type .................................. 37
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Technical Assessment Report 

1.0 Notification and Authorization 

In January 2007, the NASA Administrator and Associate Administrator for the Exploration 

Systems Mission Directorate (ESMD) chartered the NASA Engineering and Safety Center 

(NESC) to form an Agency team to design and build a Composite Crew Module (CCM) in 18 

months to gain hands-on design, build, and test experience in anticipation that future exploration 

systems may be comprised of composite materials. The NESC Review Board (NRB) approved 

an assessment plan on May 3, 2007. Status briefings were presented at NESC face-to-face 

meetings in June 2006, May 2007, December 2007, March 2008, July 2008, January 2009, June 

2009, and February 2010. The key stakeholders for this assessment are the ESMD and the 

Office of Chief Engineer. 

The CCM project started in January 2007 and the technical work was completed with the final 

test to destruction in March 2010. Each discipline of the project achieved great success and 

developed extensive written results. Accordingly, the final CCM project report has been 

separated into one summary report with executive summaries from each of the six disciplines, 

and then one stand-alone report for each discipline. This results in some duplication between 

reports, but allows readers to limit their review to their particular discipline of interest in 

searching for lessons learned for future composite structure activities. 
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2.0 Signature Page

Submitted by: 

Team Signature Page on File - 5/16/11 

Mr. Michael T. Kirsch Date 

Team Lead 

Team Leads: 

Mr. Jeffrey W. Stewart Date Mr. Paul W. Roberts Date 

Deputy Team Lead/Design Lead Deputy Team Lead 

Dr. Daniel L. Polis Date Mr. James W. Jeans Date 

Materials Lead Analysis Lead 

Mr. Larry Pelham Date Mr. Kenneth L. Hodges Date 

Manufacturing Lead Inspection Lead 

Dr. Sotiris Kellas Date 

Test Lead 

Signatories declare the findings and observations compiled in the report are factually based from 

data extracted from Program/Project documents, contractor reports, and open literature, and/or 

generated from independently conducted tests, analysis, and inspections. 
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3.0 Team List 

Name Discipline Organization/Location

Core Team 

Michael Kirsch Team Lead/NESC Principal Engineer LaRC 

Jeff Stewart Deputy Team Lead/Design Lead GSFC 

Paul Roberts Deputy Team Lead LaRC 

Jim Jeans Analysis Lead GSFC 

Dan Polis Materials Lead GSFC 

Larry Pelham Manufacturing Lead MSFC 

Ken Hodges Inspection Lead JSC 

Sotiris Kellas Test Lead LaRC 

Pamela Throckmorton Management and Technical Support 

Office LaRC 

Administrative Support Personnel 

Terri Derby Project Coordinator LaRC 

Christina Williams Technical Writer LaRC 

3.1 Acknowledgements 

Appendix B provides a full list of team members that contributed to the CCM project.
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4.0 Executive Summary 

In January 2007, the NASA Administrator and Associate Administrator (AA) for the Exploration 

Systems Mission Directorate (ESMD) chartered the NASA Engineering and Safety Center 

(NESC) to design, build, and test a full-scale crew module primary structure, using carbon fiber 

reinforced epoxy based composite materials. The overall goal of the Composite Crew Module 

(CCM) project was to develop a team from the NASA family with hands-on experience in 

composite design, manufacturing, and testing in anticipation of future space exploration systems 

being made of composite materials. The CCM project was planned to run concurrently with the 

Orion Project’s baseline metallic design within the Constellation Program (CxP) so that features 

could be compared and discussed without inducing risk to the overall Program. 

This report discusses the project management aspects of the project including team organization, 

decision making, independent technical reviews, and cost and schedule management approach. 

This report also contains an analysis of the actual expenditures by civil servants, contractors, and 

project phases. There are six final reports associated with the project being released: Design, 

Analysis, Materials and Processes, Manufacturing, Test, and Non-Destructive Evaluation (NDE).

Each of these reports contains a comprehensive discussion of the engineering behind the 

technical decisions within each discipline. Near the end of the manufacturing development 

phase, the team worked with experts from the Orion project to identify implementation issues 

associated with switching the Orion primary structure from their aluminum-lithium (Al-Li) 

baseline to the composite system. There were no mass or cost benefits identified to warrant the 

switch, and cost and schedule risk were increased for the Orion project, thus it was decided that 

Orion would remain with the Al-Li baseline. 
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5.0 Background 

In January 2006, the NASA Administrator, Dr. Michael Griffin, and NASA AA for ESMD,

Dr. Scott Horowitz, chartered the Crew Exploration Vehicle (CEV) “Smart Buyer Team (SBT)” 

Design project. The CEV SBT project’s goal was to design a CEV to meet the requirements 

established in the Call for Improvements Request for Proposals released January 10, 2006, to 

better understand trade space, design drivers, and technical solutions. 

During the CEV SBT activity, the structures team conducted a parametric analysis of polymeric 

composite materials for use on the crew module (CM) primary structure and pressure vessel.

Following the CEV SBT outbrief, an NESC design team was chartered to develop a CM design 

optimized around the advantages of composites and to characterize the design drivers such as 

geometry, mass, manufacturability, inspectability, repairability, damage tolerance, 

crashworthiness, micro meteoroid orbital debris, and radiation shielding. The outcome of the 

NESC design team was that a CCM was feasible, but full-scale fabrication was suggested to 

quantify mass and manufacturability benefits. The NESC feasibility study was published in the 

NESC (RP-07-028) Composite Crew Final Report, Volumes I and II. 

In January 2007, NASA chartered the NESC to form an Agency team to design and build a fullscale CCM to gain hands-on design, build, and test experience in anticipation that future 

exploration systems may be made of composite materials. 

6.0 CCM Project 

6.1 Project Objectives and Success Criteria 

The primary objective of the CCM project was to provide members of the NASA family handson design, build, and test experience on a complex composite structure so that NASA would 

have the skills for future Exploration hardware projects (e.g., a lunar lander, Earth departure 

stage, or lunar habitat module). The effort was to be completed on a full-scale Orion CM 

primary structure in parallel, but not in competition with the baseline Orion Al-Li design. The 

intent was to perform a detailed design and fabrication of both an Al-Li design by the main line 

Orion project, and a composite design by the CCM team, so that details could be compared and 

contrasted.

The project’s success criterion was defined as being able to predict both the as-built mass of the 

final assembly (using Preliminary Design Review (PDR) information) and the strain response of 

the structure as tested within 20 percent of the measured values. 

While not considered success criteria, one of the project’s goals was to deliver a full-scale 

structure for test, quickly; approximately 18 months from project kick-off to delivery. This goal 

emphasized rapid prototype management techniques, but from a geographically distributed team. 

This resulted in a project organization, communication plan, file management, and meeting 

approach that was not typical to most NASA projects. 
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6.2 Major Project Milestones 

The overall project from the first kickoff meeting to the final hydrostatic test to failure spanned 

36 months, as shown in Table 6.2-1. The first 12 months were spent in design and the second

5 months were spent developing manufacturing techniques. The project had a production stand 

down from May 2008 to October 2008 due to agency funding constraints and priorities. In 

October 2008, the project resumed and the test article was manufactured in approximately 

9 months. It took 2 months to make final repairs and prepare for and conduct shipping. The test 

article arrived at Langley Research Center (LaRC) in September 2009. Testing was completed 

over 5 months, from October 2009 to January 2010, with the final destructive test completed 

about a month later on March 2, 2010. See Appendix A for the project milestone schedule. 

Table 6.2-1. Itemized List of Reviews 

Phase 

Months

Duration Date Event

Design 12 

Jan-07

Project Kickoff: Kennedy Space Center (KSC) Beach 

House

Mar-07 Conceptual Design Review 

Jun-07 Preliminary Design Review 

Dec-07 Critical Design Review 

Development 5 Jan-08 Full Scale Tooling Delivered 

May-08 Manufacturing Plan Review 

Manufacturing 9 Oct-08 Full Scale Test Article Started 

Jul-09 Full Scale Test Article Completed 

Testing 5 

Sep-09 Delivery of Test Article to LaRC 

Sep-09 Test Plan Review 

Jan-10 Baseline Full Scale Tests Completed 

Failure 1 Mar-10 Hydrostatic Test to Failure 
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Although the 18-month goal was not achieved, the project still proceeded at a relatively rapid 

pace. 

6.3 Project Organization

The team was staffed the way a traditional NESC team is formed, in that expertise was recruited 

Agency-wide. The team had participation from nine of the ten NASA field Centers including 

Ames Research Center, Dryden Flight Research Center , Goddard Space Flight Center (GSFC), 

Glenn Research Center , Johnson Space Center (JSC), Jet Propulsion Laboratory, KSC, LaRC, 

and Marshall Space Flight Center (MSFC). There were also participants from industry including 

Alcore, Alliant Techsystems (ATK) Aerospace Structures, Bally Ribbon Mills, Collier 

Industries, Genesis Engineering, Janicki Industries, Lockheed Martin, Northrop Grumman and 

Tayco Engineering. The team was managed as a collaborative effort; every person on the team 

was performing an in-line function for the project rather than mixing in-line work with civil 

servant oversight. This approach enabled individual contributions as a function of expertise, 

rather than NASA or company affiliation and resulted in increased innovation and a higher 

motivated team. 

The team was organized into six groups; design, analysis, materials and processes, 

manufacturing, NDE, and test. A group lead provided focus for each discipline, but all project 

decisions were made in a consensus environment. Each discipline had multi-Center and industry 

participation, to ensure a broad decision space was considered. 

6.4 Team Communications and Decision Making 

Since the team included membership spanning all four time zones of the continental United 

States, actions were taken early on to build relationships amongst the different participants. The 

project began with 5 weeks of back-to-back collocations, starting with a kick-off meeting at 

KSC. The collocations were meetings where the team was physically present in one location.

The objective of the first week of the first collocation was to define the project goals and 

objectives, expectations, and to understand concerns from individual team members. Following 

the week at KSC, the team travelled to Janicki Industries in Sedro Woolley, Washington, to 

participate in a hands-on build exercise of a complex curvature part. Again, the primary goal of 

the week was to begin forming the relationships between participants, but also to see firsthand 

how tooling complexity directly affects the manufacturability of a composite part. 

The following 3 weeks were spent at GSFC in a high intensity extended duration conceptual 

design phase. During this time, many hours late into the evening and through the weekend 

resulted in trades between different structural concepts. The following 2 weeks, the team 

returned to their home-base work place, and operated virtually. During weeks 8 and 9 of the 

project, the team returned to GSFC to prepare for and conduct the first independent technical 

review of the conceptual trades considered and the proposed solution and path forward. 

The first 9 weeks of the project were important to the spirit of the project. It provided an 

opportunity to assess the strengths and weaknesses of the different participants; it illustrated, for 

the participants, the pace and methodology for decision making on the project; and “drew a line 
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in the sand” for the assumptions on loads and interfaces that would remain for the duration of the 

project.

The collocations were beneficial for the achievement of great results in a timely manner. 

However, given the likely duration of the total project, and the fact that team membership 

spanned the country, a secondary goal for the project was created to develop project management 

and system engineering techniques that would leverage the state-of-the-art in communication 

tools and to facilitate project management in a virtual environment. Emphasis was increased on 

decision making, conflict resolution, and problem solving, as if the participants were in the same 

room. In other words, when conflicts were encountered, the question was asked, “how would we 

solve this if we were in the same room?” This mindset led the team to more collaborative 

problem solving sessions and significantly increased the multi-location desktop sharing. 

Traditional teleconferencing was used extensively, coupled with the use of WebexTM meetings. 

Overall project communication meetings occurred Mondays, Wednesdays, and Fridays late in 

the morning to allow east and west coast time zones to participate with minimal inconvenience. 

Additionally, design sessions were scheduled for Tuesday and Thursday afternoons, but occurred 

nearly daily to reinforce the need for participatory design decisions. Teleconferences were 

augmented with messaging services, such instant messaging. This instant messaging tool had the 

largest impact in terms of creating the environment simulating face-to-face participant 

interaction. An extension of the instant messenger tool is a desktop sharing function in a simple 

two-step process that connects two desktops virtually to allow a team member in one location to 

view and, if necessary, control an application at a different location as long as both team 

members are connected on the internet. This was the most powerful tool for simulating the 

virtual problem solving “as if in the same room.” 

Although, state-of-the-art tools enabled communication, they were still not as effective as a 

collocated team. Additionally, when team members were dispersed too long, old habits of 

“singular” decision making and other distractions would creep in. To keep the relationships and 

focus intact, 2 week collocations continued on the project about every 6 weeks for the first year, 

encompassing the full design phase. 

During the project planning phase, the challenges of distributed teams were considered, 

particularly the challenge of team members being on travel for extended periods of time. There 

were two aspects of collocations that could be difficult for team members. The first, 

acknowledged early on, was the burden of team members having to travel to collocations and 

being away from home for extended periods of time. This was managed by establishing 

aggressive collocation goals that resulted in longer hours than normal. The second burden was 

the challenge of having the collocations at team member’s respective home Center. While there 

may have been an expectation of regular work hours, the collocations required long daily hours 

and weekend work from every team member. This schedule often competed with the 

responsibilities at home and made the burden of collocations difficult for the local participants.

This factor was not completely appreciated at the inception of the project. 
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After a year of design, the project transitioned into the manufacturing development phase. 

Routine project communication meetings continued to occur Mondays, Wednesdays, and 

Fridays, along with the Tuesday and Thursday afternoon design sessions. Additionally, a 

manufacturing tagup occurred on a daily basis each morning. The purpose of the manufacturing 

tagup was to discuss the plans for the day and the results from the previous day. These meetings 

were the primary communication tool between manufacturing and design, facilitating the

“hand-off” from design to manufacturing, and discussing any issues that were identified. The 

daily virtual sessions were coupled with frequent visits to the manufacturing facility at ATK 

Aerospace Structures in Iuka, Mississippi. Since one of the project’s primary objectives was to 

gain the hands-on experience, learning opportunities were accessible with design, analysis, 

materials, test, and project management personnel spending “time on the shop floor” 

participating in the handling, layup, bagging, curing, etc. This gave immediate feedback to team 

participants on the design and manufacturing details. The frequent trips to Iuka, and the daily 

manufacturing tagups continued throughout the 18-month manufacturing development, and 

manufacturing phase. 

As the project evolved from manufacturing to test, the daily tagups continued, but moved to 

daily test goals and results discussions accordingly. During the test phase, the Monday, 

Wednesday, and Friday project management meetings were reduced to Mondays only. 

6.5 Independent Review Process 

Throughout the different phases of the project, numerous independent technical reviews were 

conducted to challenge assumptions and methodologies. The reviews were typically conducted 

during the second week of a 2 week collocation. The first week was dedicated to completing 

engineering tasks and preparation of review materials. 

The review committee was formed by recruiting senior experts from NASA and Industry with 

the intent of maintaining participant continuity throughout the review process. The review team 

was generally about 15 to 20 multi-discipline experts, representing materials, analysis, 

manufacturing, systems engineering, or test expertise. The team did not use a closed loop 

required item disposition process; rather, the reviewers were asked to state or record their 

questions, comments, and concerns during the review. Their notes were then collected at the end 

of the day, copied, and returned. The team then reviewed the notes and developed internal team 

actions. These actions were delegated to the appropriate discipline on the team and tracked 

internally. 

An independent review was conducted approximately 2 months into the project to review the 

conceptual designs and the down selects that occurred. Approximately 6 months into the project 

an independent review was conducted of the preliminary design and sizing results. The PDR 

was also the point that locked in the project interfaces, loads, environments, and final tool 

surface. Approximately 12 months into the project, an independent review was conducted of the 

critical design. This included review of the different computer models: analytical and computeraided design (CAD). At 17 months into the project, an independent review was conducted on the 

manufacturing plan. This review discussed the manufacturing approach for the entire 
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manufacturing phase including a discussion on the development work that had been done to date. 

Finally, at 33 months into the project, the team conducted the final independent review of the 

planned testing. At this review, the team presented the rationale comparing the design conditions 

to the planned test conditions and presented the analytical predictions based on the test 

conditions. Finally, the methods were presented for inducing the planned loads during test. An 

itemized list of reviews is contained in Table 6.2-1 above. 

6.6 CCM Configuration Management 

6.6.1 Data Management System 

All data for the CCM project were maintained in the Integrated Collaborative Environment (ICE) 

Windchill data management system. Windchill is a product lifecycle management system 

commercially available from Parametric Technology Corporation® (PTC®), The Product 

Development Company®. Team members had access to the system by using their registered 

username and password through the URL http://ice.exploration.nasa.gov/windchill. Within the 

Windchill system, two repositories were used: a Windchill “project” and a Windchill “product.” 

Windchill Project 

The CCM project repository was used for the creation and storage of design and manufacturing 

data not subject to revision control. Documents submitted to the project repository were 

maintained with a minor version history—a new minor version was created each time the 

document was “checked-in” to the project. Data was segregated into project sub-folders by 

discipline—“Design,” “Materials,” “Structural Analysis,” and “Project Administration.” 

Windchill Product 

The CCM product repository was used for the creation and management of design 

documentation subject to revision control and lifecycle state based access control, which formed 

the basis of the release and change management process for product data. Documents (reports, 

test plans, specifications, etc.) and all CAD data (mainly Pro/ENGINEER® (Pro-E®)) were 

controlled within the CCM product. “Part items” and bill of material data was also generated, 

though this capability was not exploited. 

6.6.2 Numbering 

To avoid future potential numbering conflicts within the ICE Windchill data management 

system, all data items checked into the CCM product were prefixed with a unique project 

designator. For all documents, the designator was “CCM-”; for CAD data, the designator was 

“CCM_”.
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Document Numbering 

Document numbers were generated sequentially by document type. Table 6.6-1 lists the 

document types and corresponding document nomenclature. 

Table 6.6-1. CCM Document (Object) Types 

Document Type Number Prefix 

Analysis CCM-ANYSDocument (General) CCM-DOCPlan CCM-PLANProcedure CCM-PROCReport CCM-RPTStatement of Work CCM-SOWSpecification CCM-SPECStandard CCM-STDCAD/Drawing Numbering 

Blocks of numbers were reserved from NASA LaRC. These CCM number ranges were 

1247531 through 1247917 and 1247951 through 1248573. Assignment of numbers was 

maintained on a spreadsheet that was stored in the CCM project. All object numbers were 

prefixed with CCM_ but the prefixes were removed on the drawing title block and in the onsheet parts lists. 

6.6.3 Revisioning 

Object revisions were designated using Military Standard (MIL-STD) convention. Pre-release 

CAD data was assigned numerical revisions (1,2,3..). All data was assigned Rev “-” (no 

revision) at initial release. MIL-STD alpha revisions were assigned to revised items. 

6.6.4 Lifecycle States 

Access controls and revisions of all data in the Windchill product were controlled by applying 

rules at different data maturity levels (lifecycle states). These rules are listed in Table 6.6-2. 
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Table 6.6-2. Lifecycle State Access Rules 

Object

Type

Lifecycle State Revision 

Sequence

Access 

Document Development “-“ Create / Modify 

Pending Approval / Under Review “-“ Read Only 

Released “-“, A, B, 

C... 

Read Only 

Production Change A, B, C... Create / Modify 

CAD Development 1, 2, 3 Create / Modify 

Pending Approval / Under Review “-“ Read Only 

Released “-“, A, B, 

C... 

Read Only 

Approved “-“, A, B, 

C... 

Read Only 

Production Change A, B, C... Create / Modify 

6.6.5 Electronic Approval 

All data was released using an electronic approval process. All CAD data that were submitted 

for release were viewable using ProductView, the Windchill visualization client. ProductView 

enabled non-CAD users to view CAD data without requiring the native CAD system. Successful 

completion of the electronic approval advanced the lifecycle state-of-the-objects to be released.

For CAD data, revision and lifecycle state values were reflected on the face of the drawing. 

6.6.6 Change Management 

Changes to released data were initiated using an engineering order (EO) for CAD data or a 

document change order (DCO) for documents. The EO or DCO was released as a document. 

Upon release of the EO or DCO, the prescribed changes were implemented as a revision to the 

affected items. 

6.7 Cost and Schedule Management Approach 

6.7.1 Schedule Management 

The cost and schedule management approach for this project was critical path focused. The 

critical path of the project was reviewed and updated no less than weekly by most team leads. At 

the onset of the project, the critical path was fairly high level including milestones such as 

concept review, PDR, and critical design review. As the project progressed, detail was added to 

the schedule based on an understanding of the design. As the design was completed, the 

manufacturing schedule emerged in conjunction with the release of the drawing tree. Thus the 

manufacturing schedule had considerable more fidelity for managing major manufacturing 

operations. As manufacturing tasks were underway, a day-by-day, or even hour-by-hour 

schedule was developed for the following 2–4 weeks. Major manufacturing tasks beyond the 

immediate 4-week window were left vague with estimated durations at the parent task. As the 

detailed manufacturing tasks were completed and work progressed through the schedule, detail 
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was added to tasks that entered the most current 4-week sliding window. This provided a lower 

maintenance mechanism for incorporating early lessons learned in later manufacturing 

operations. It also highlighted the effects of threats and opportunities since most views of the 

schedule including the 4-week detail could be viewed on a one page printout. To illustrate the 

concept, during the manufacturing phase, the test phase was one line in the schedule. This 

rolling 2–4 week list of critical path elements was used throughout the manufacturing phase into 

the test phase, until test completion. This made the schedule a valuable tool for managing 

priorities, exploiting opportunities, and minimizing the effects of identified risks. It also served 

as an effective communication tool for communicating information between team leads. 

6.7.2 Cost Management 

The CCM project’s original cost estimate was considered an analogous estimate based on expert 

judgment, meaning labor costs were derived by multiplying a rough estimate of the size of the 

workforce by a generic hourly rate and then multiplied by the total duration of the project. Raw 

materials and other procurement costs were then added and estimated travel and other 

miscellaneous expenses. 

Once the project was underway, the estimate at completion (EAC) was the primary tool for 

managing overall costs. EAC was calculated by looking at actual costs (AC) and adding an 

estimate to complete (ETC). The ETC was calculated by multiplying the estimating labor burn 

rate for the remainder of the work times the remaining project duration based on the critical path.

Travel, major procurements, and onetime expenses were then added to the labor total to 

determine the total EAC. The burn rate was adjusted periodically using actual data, engineering 

judgment, and the project phase. 

EAC = AC + (estimated burn rate x remaining schedule duration) + onetime expenses 

While the project was managed considering technical, schedule, and cost, strict earned value 

management (EVM) formulas, including the use of planned value (PV) and the earned value 

(EV) were not used on this project due to the high amount of uncertainty associated with rapid 

prototype projects. In that regard, the basis of estimate for individual tasks beyond 4 weeks in 

the future was too uncertain to accurately estimate traditional EVM. To roll the estimate up at a 

higher level would have resulted in maintaining two different schedule and cost estimates: one 

for managing earned value and one for prioritizing work. Alternatively, one schedule could have 

been used for tracking both EV and critical path management, but the project would have been 

rebaselining the PV schedule curve as the detailed schedule for future tasks became evident. 

Given the short term (2–4 weeks) rolling schedule detail, the project would also have to 

rebaseline the EV cost curves every time the 2-week rolling window shifted. This schedule 

uncertainty made the use of traditional EV tools too labor intensive and of limited value since the 

project would have rebaselined far too frequently to determine any proactive cost performance 

index or schedule performance index. In that regard, the project was managed using EVM 

constructs, considering schedule, costs, and technical content. However, it proved best to 
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manage the costs based on current and future burn rates and compare anticipated versus actual 

cost and schedule deviations. 

6.8 Project Cost Analysis 

The CCM project total cost was $29.8M. The most significant contract contributors, as defined 

by resources expended, were ATK Aerospace Structures (design and test article fabrication), 

Northrop Grumman (test article design and manufacturing engineering), Janicki Industries (test 

article tooling fabrication), Genesis Engineering (design and analysis), and Collier Research 

(structural analysis). Contracted services and hardware represented some 79 percent ($23.5M) of 

total project cost, while civil service labor accounted for 19 percent ($5.6M) of the total project 

cost. Despite the significant use of collocations, and travel associated with much of the team 

participating in the manufacturing phase and the test phase, the overall civil servant travel 

expenditures were relatively low at 2 percent ($0.7M), well within the bounds of cost uncertainty 

on this development project. As a development project with extensive engineering, the project 

costs were large with respect to the hardware costs. This was driven by the need to maintain all 

engineering skill sets throughout the project lifecycle. The breakdown of total project costs 

among the full-cost categories of procurement (which includes work year equivalent (WYE) 

costs), civil service labor, and civil service travel is illustrated in Figure 6.8-1.

Figure 6.8-1. CCM Total Costs: $29.8M (FY07–FY10) 

A breakdown of costs by project phase is depicted in Figure 6.8-2.
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Figure 6.8-2. CCM Costs by Phase: Total $29.8M 

6.8.1 Labor Analysis 

Figure 6.8-3 depicts the total work effort expended in the execution of the CCM project. Over 

the lifecycle of the project, some 94 equivalent years of work effort were expended by the CCM 

civil service/contractor team—52 percent by civil servants and 48 percent by contractors. Figure 

6.8-4 depicts the work effort between the civil servants and the contractor team by project phase. 

The design phase work was fairly equal between the groups, but the contractor group was 

slightly higher by percentage in manufacturing as expected, given the touch labor associated with 

manufacturing. The test phase had a high percentage of civil servants. This was driven by the 

civil servant labor pool at the LaRC combined loads test facility. 
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Figure 6.8-3. CCM Total Workforce – Full-Time Equivalent (FTE) and WYE 

Figure 6.8-4. CCM Total Workforce – by Project Phase 
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6.8.2 Project Implementation Challenges 

During the lifecycle of the CCM project, a number of finance-related circumstances presented 

significant challenges for the management of the project. First, a temporary limitation of 

available funding required a production slow-down in the pace of technical progress during the 

time period from June 2008 through September 2008. This circumstance, and other schedule 

extensions driven by funding limitations, resulted in increased levels of total project cost, due to 

requirements to maintain staffing levels over longer periods of time than initially planned. The 

project also experienced some unanticipated cost growth which resulted from a number of 

technical circumstances, such as: 

• Metallic components were not included in original estimate. 

• Support tooling and shop aids such as drill fixtures, router templates, etc., were not 

included in the original estimate. 

• Transition from Pro-E®/NASTRAN® models to released drawing took 2–3 months longer 

than expected. 

• Just in time engineering to manufacturing caused manufacturing inefficiencies and false 

starts. 

• Overall layup was more complex than originally expected, specifically core application, 

including core forming, impression testing, and final bonding. 

• Manufacturing estimate did not accommodate for multi-cure process that included 

inspection of the skin, and installation and cure of the core, and lastly installation and 

cure of the outer skin. 

• Manufacturing estimate did not account for significant post cure assembly (e.g., 

500 precision reamed holes for Service Module (SM) and Alternate Launch Abort 

System (ALAS) fittings). 

• Overall project schedule growth caused design burn rate to be sustained through 

completion of testing. 

7.0 Design Overview 

The CCM baselined a set of loads and interfaces based on the Orion design in March of 2007. 

Beginning at the docking tunnel, the CCM used an aluminum ring to represent the docking 

system interface. This provided a generic interface that could easily be tailored to any of a 

number of docking system options. If the CCM had been a flight project, the ring would have 

been fabricated from titanium to alleviate thermal expansion mismatch between the metallic ring 

and the composite tunnel interface. A titanium ring could not be procured in time to meet the 

CCM critical path. The aluminum ring was sufficient to verify the design. The drogue chute 

interface from Orion is on one unique side of the docking system interface. For Orion, this 

interface was moved to coincide with a main chute interface. CCM did not incorporate this 

interface change. 
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The CCM has six gussets that align with six internal longerons. These include three metallic 

main parachute interface fittings on every other longeron. There are six metallic SM and launch 

abort system (LAS) interface fittings along each longeron. Although Orion had baselined an 

LAS interface that connected to the pressure shell primary structure near the main parachute 

fittings, a trade was underway in Orion that considered attachment at the major diameter near the 

SM interface hard points. The CCM project chose to adopt the interface at the major diameter, 

assuming Orion would also choose this path. However, Orion did not choose this interface and 

this represented one of the first points of divergence between CCM and Orion. The longerons 

are internal to the facesheets in the sandwich system and do not represent an additional part for 

assembly. See Figure 7.0-1 for locations of key features. 

Figure 7.0-1. NESC CCM 

The CCM included the major cutouts consistent with the Orion design: two side windows, two 

docking windows, and a main hatch. The cut outs are machined into the side of the pressure 

shell and fit with a matching aluminum ring. The composite structure was sized to carry all of 

the design loads, and the aluminum frames were installed to provide a sealing surface for the 

closeouts during test. Similar to the tunnel, a flight like design would incorporate either titanium 
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frames, or would have the window frames incorporated directly into the layup. For CCM, both 

of these options were not incorporated due to schedule constraints. 

The CCM was built as two halves, an upper and lower pressure shell as shown in Figures 7.0-2 

and 7.0-3. It was manufactured on two different male tools: one for the upper and one for the 

lower. The two shells were joined using a double lap shear joint, with twelve different internal 

and external doublers; one for each longeron, and one for the acreage between longerons. The 

internal and external doublers were cured coincident with each other during 12 separate cures.

The cures were done under vacuum bag pressure with a purpose built Kapton® heater. The 

motive for the design was that the spacecraft assembly and integration timeline could be 

shortened by allowing for parallel integration of labor intensive subsystems such as wiring, 

plumbing, etc., on each shell without the confined space constraints, and then the two halves 

could be joined outside the autoclave. 

Figure 7.0-2. CCM Upper Shell 

Figure 7.0-3. CCM Lower Shell 
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The majority of the construction was honeycomb sandwich with unvented aluminum honeycomb 

core. The unvented core was selected to provide double redundancy in pressure containment 

using the two skins of the sandwich. Given the mild temperature environment of the crewed 

pressure shell, unvented core was an acceptable selection since large thermal excursions would 

not be a concern. Some areas of solid laminate were used in the design, specifically the tunnel, 

the backbone cap, and the pan down regions where metallic fittings such as parachute fittings 

and SM interface fittings were bolted through the shell. 

The CCM incorporated the backbone feature from the Orion design. The backbone was a

grid-like structure in the lower shell used to manage packaging in the Orion solutions, see

Figure 7.0-4. For Orion, it was originally a non structural feature but, for CCM, the design team 

chose to integrate the backbone into the primary load paths. Specifically, the backbone was 

extended and secured to the aft dome, see Figure 7.0-5. This allowed for the reduction of 

structure around the main diameter of the aft dome since the backbone could carry some of the 

bending loads in the aft dome. Connecting the aft dome to the backbone saved approximately 

100 lbs of mass in the CCM baseline. In addition, by extending the backbone to the floor, a hard 

point pattern that the heat shield could bare up against on the exterior of the floor was provided.

By load sharing between the SM interface and the backbone interface, the heat shield stiffness 

could be reduced, thus saving mass in the heat shield—an overall system advantage. 

Furthermore, the aft dome was allowed to take the desired membrane shape between the major 

beams of the backbone system. This created the “lobed” shape, a hallmark in the CCM design, 

and also reduced the mass of the aft dome by another 50 lbs, over the CCM baseline, see Figures 

7.0-6 and 7.0-7. The exterior foot print for sharing heat shield loads was also preserved. 

Figure 7.0-4. Final CCM Assembly

Design Summary

Service 

Module/Alternate 

Launch Abort 

System (SM/ALAS) 

Fittings (6 pl )

Tunnel

Ceiling

Barrel

Conic

LIDS Ring 

Gussets (6)

Drogue Chute I/F (1)

Main Chute I/F (3)

Docking Windows (2)

Hatch Opening (1)

Side Windows (2)

Assembly Splice
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Figure 7.0-5. Illustration of Backbone Structure within Lower Shell

Figure 7.0-6. Backbone Structure in Lower Shell
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Figure 7.0-7. Lobed Floor on Lower Shell 

One of the challenges in composite system design was the orthogonal joints. The team, 

collaborating with industry partners, adopted three dimensional (3D) woven Pi-preforms to 

facilitate orthogonal structural joints, particularly between panels of the backbone, between the 

backbone and the CCM floor and between the gussets and CCM tunnel, see Figure 7.0-8. These 

3D woven preforms were key enablers in the composite design and represent a promising future 

for composites joints due to their robustness as possible design solutions. 

Figure 7.0-8. 3D Woven Preform Test Element 
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8.0 Materials and Processes Overview 

The materials and process activities on the CCM project were prioritized based on a rapid 

prototype developmental model of the project. As such, the materials selection activity was 

driven as much by project constraints as performance optimization. Specifically, one of the 

primary selection criteria was that a sufficient amount of data was pre-existing, such that the 

design activity could proceed without an extensive lamina/laminate characterization program. 

Other examples include the fact that a single prepreg system (fiber and resin) was chosen for 

both autoclave shell construction and out-of-autoclave splice construction. Finally, an 

aluminum-docking ring was chosen over a titanium version, precluding the ability to co-cure the 

ring with the shell and driving out-of-autoclave and out-of-oven closeout bonds. While the team 

was able to successfully implement these solutions, they were responsible for several 

developmental anomalies. In addition, further optimization of materials selection could 

potentially reduce weight and/or improve performance. 

Allowables and associated failure criteria were assembled from a compilation of literature 

sources with the intent of providing a flexible design tool that accounted for damage tolerance. 

This was accomplished by combining a lamina based failure criterion (Hoffman) with laminate 

damage tolerance considerations (open hole strain allowables) [ref. 1]. For the allowable defect 

sizes that were chosen, this approach is shown to be conservative, and may provide a fruitful 

avenue for weight savings at the expense of laminate-based testing. 

By eliminating the traditional lamina and laminate building block level testing, the materials and 

process activities could focus on design details with greater risk and uncertainty, such as woven 

preform joints. The 3D woven preforms, specifically Pi-preforms, were used to join critical 

subassemblies together on the CCM. Because this technology was relatively immature 

compared to the rest of the materials technology on the CCM, considerable resources were 

focused on establishing process baselines, performance baselines, and design envelopes.

The Pi-preform joining technology proved to be an efficient design solution (~2 times stronger 

than composite L-clip design) and provided significant manufacturing advantages over such 

traditional joining technologies. One of these advantages was the ability to do all the Pi-preform 

joining using out-of-autoclave curing. In addition, the 3D woven technology was extended to a 

woven cruciform, capable of handling bi-directional joint tension to levels >3 times that of a 

Pi-preform (7,000 lbs/in versus 2,000 lbs/in pull-off). 

Much of the process validation occurred during the full-scale manufacturing phase of the project. 

Witness testing was used to validate many of the engineering assumptions made during the 

design phase of the project, including elastic properties, strength distributions, thickness 

variation, and moisture effects. 

Several process anomalies were encountered during the development of the CCM. They include 

a disbond of the inner skin-to-core, a splice re-design, and Pi-preform consolidation defects. 

While process development and associated building block testing were performed in each of 

these areas, the anomalies manifested themselves at the full-scale level due to interactions 
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between process robustness and manufacturing scale-up. This highlights the importance of fullscale developmental work early in the development phase of a composite program. 

A full-scale damage tolerance program was implemented to demonstrate that the CCM could 

meet the intent of current NASA requirements with minimal design changes. This included 

impacting the vehicle at 18 unique locations at 6 ft-lb, defined as the allowable threat level.

Following these impacts, the test article was taken through two critical design ultimate load cases 

followed by life cycling. In addition, five design details were taken to critical threat levels 

(reliable detection or threshold energy of 26 ft-lb) and subsequently taken through life-cycling 

without detrimental defect growth. The full-scale impact testing did highlight the variability 

associated with the paste bond of the docking ring, which was a potential area for design 

improvement despite meeting requirements. 

Leakage (consumable gas loss rate) remains one of the tipping points between accepting a 

composite replacement over a metallic solution. In this area, the team demonstrated how a 

simple polymeric film, loosely coupled to the facesheets using a viscoelastic adhesive, 

dramatically improved the leakage after impact behavior of the composite system. In addition, 

the team demonstrated that even the minimum gage skins were not appreciably permeable 

(below detection limits) under biaxial strains of greater than 4,000 μin/in.

A much more detailed discussion of the materials and process engineering for the CCM project 

is provided in the NESC 06-019 Composite Crew Module: Materials and Processes Report. 

9.0 Analysis Overview 

The CCM structure was designed and analyzed to support 115 separate load conditions, 

including launch, abort, pressure, trans-lunar insertion and landing cases. A “freebody” of the 

load magnitudes is shown in Figure 9.0-1. 
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Figure 9.0-1.CCM Loads “Freebody” 

CCM analysis was accomplished through the use of computer simulation using finite element 

modeling techniques backed up with hand analysis and engineering judgment. 

Two NASTRAN® finite-element models were constructed of the CCM composite structure. The 

first was a fairly coarse grid model consisting of about 30,000 nodes and elements. Many of the 

design drawings were released for manufacturing based on this coarse grid model. This model 

was updated along with the design of the CCM as the design evolved. Once the design reached a 

stable level of maturity, a fine grid model was constructed. The fine grid model consisted of 

215,000 nodes and elements. The fine grid model confirmed the sizing accomplished with the 

coarse model and was used to write final margins and predict strains. 

A goal of the CCM project was to accurately predict the as-built mass at the PDR. To 

accomplish this, the coarse grid model was sized to achieve all positive margins using 

Hypersizer®, along with internally imposed constraints such as layup symmetry and similarity. 

This sizing was used to accurately predict the final mass of the CCM. 

Hypersizer® was used throughout the CCM project. Hypersizer® was used in initial sizing to 

determine suitable layups for a given area. It was also used to track margins as the design 

evolved, ensuring that each area maintained sufficient margin when something changed. 

Additionally, Hypersizer® was used to perform trade studies to study the impact of a suggested 

change to the design. 
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The final sizing of the CCM yielded numerous low margins that are summarized in Table 9.0-1. 

Section reference in Table 9.0-1 refers to the section of the Analysis final report where the 

margins are described. 

Table 9.0-1. Margins of Safety Summary 

Another goal of the CCM project was to accurately predict strains during test. Using the fine 

grid model, strains for all critical channels were predicted well within the goal of 20 percent. In 

fact, most critical channels were predicted within 5 percent for each test condition. 

A more detailed discussion of the CCM analysis activities is provided in the NESC 06-019 

Composite Crew Module: Analysis Report. 

10.0 Manufacturing Overview 

The CCM manufacture, including pressure shell fabrication, composite component fabrication, 

and CCM assembly, employed a combination of existing manufacturing technologies and best 

industry practices in conjunction with processes designed around the data produced from the 

building block approach used in the rapid prototype model for the project. Additionally, a fullscale pathfinder upper pressure shell assembly was manufactured to qualify and, in many 

instances, revised key manufacturing processes and requirements.

Location Description Material Section Load Case Failure Mode FOS MS

Longeron 1 Gusset Web IM7/977-2 10.3.1.1 707 Hoffman 1.40 +0.05

Longeron 1 below hip radius Al Core 5052 10.3.1.2 601 Core Shear 2.00 -0.07

Longeron 4 parachute fitting pandown IM7/977-2 10.3.1.3 921 Hoffman 1.40 -0.07

Docking Window Edge IM7/977-2 10.3.1.4 601 Hoffman 2.00 +0.09

Ceiling Core Al Core 5052 10.3.1.5 601 Core Shear 2.00 +0.35

Area above SM/ALAS fitting IM7/977-2 10.3.2.2 302 Hoffman 1.40 -0.16

Drogue Bracket to LIDS A286 10.4.1.2 9917 Bolt Yield 1.25 +0.14

Gusset Cap to LIDS bolt IM7/977-2 10.4.1.3 901 Bearing 1.40 +0.06

Gusset Cap to Parachute Fitting IM7/977-2 10.4.1.4 703 Bearing 1.40 +0.05

Gusset Web to LIDS A286 10.4.1.5 914 Bolt Yield 1.25 +0.31

Gusset Web to Parachute Fittings A286 10.4.1.6 932 Bolt Yield 1.25 +0.33

Parachute Fitting to Ceiling A286 10.4.1.7 706 Bolt Yield 1.25 +0.20

Parachute Fitting to Upper Conic IM7/977-2 10.4.1.8 921 Bearing 1.40 -0.03

SM/ALAS to lower backbone fittings A286 10.4.2.1 302 Bolt Yield 1.25 -0.16

Lower Backbone fitting to shell A286 10.4.2.2 302 Bolt Yield 1.25 +0.23

SM/ALAS to Upper Backbone Fitting A286 10.4.2.3 302 Bolt Yield 1.25 +0.03

Lower Backbone Fitting to Lower Backbone Fittings A286 10.4.2.4 301 Bolt Yield 1.25 +0.09

Upper Backbone Fitting to Lower Backbone Fitting A286 10.4.2.5 301 Bolt Yield 1.25 +0.08

Gusset Web Pi to Upper PV Pi 10.5.1 601 Pi pull-off & shear 2.00 -0.34

Gusset Web Pi to Gusset Cap Pi 10.5.1 723 Pi pull-off & shear 1.40 +0.30

Backbone Web to Backbone Cap Pi Pi 10.5.2 302 Pi pull-off & shear 1.40 +0.01

Backbone Stub Keel Web to Lower Shell Pi Pi 10.5.2 601 Pi pull-off & shear 2.00 +0.30

Backbone Cruciform IM7/MTM-45 10.6 302 Cruciform Tension 1.40 +0.15

Acreage Splice IM7/977-2 10.7 601 Hoffman 2.00 +0.46

SM/ALAS Fitting 7050 10.9.2 302 tension yield 1.25 +0.04

Backbone Fittings 7050 10.9.2 302 tension yield 1.25 +0.03
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Existing manufacturing technologies employed for composite fabrication included composite 

lay-up molds, hand lay-up using laser projection for ply placement, in and out-of-autoclave cure 

processes, machining processes, contact and through-transmission ultrasonic inspection and flash 

thermography. Special tooling was developed to accomplish pressure shell machining and 

assembly for both bonded composite and fastened metallic joints. Existing technologies were 

adapted for this unique application including the use of Pi-preforms from the aircraft industry to 

bond cured composite components to internal and external surfaces of the pressure shell. A new 

application of preform developed for the CCM was the “cruciform” used to assemble four cured 

composite panels in a single cured joint as part of the backbone assembly. The use of existing 

material and process technology was successfully applied to the splicing of two large composite 

structures to create a pressure shell, which was not previously applied to this application on the 

scale of the CCM. The precision manufacture of the lay-up molds combined with simple and 

effective assembly tooling enabled the manufacturing process to overcome alignment concerns 

that could have significantly impacted the test approach and overall project results. Numerous 

lessons learned were identified throughout the course of the manufacturing process, including the 

development and execution of several significant repair processes. These lessons learned are 

addressed in more detail in the NESC 06-019 Composite Crew Module: Manufacturing Report.

The CCM Manufacturing Quality plan employed basic AS9100 processes and systems with 

reasonable exceptions needed to support the rapid prototype project. Success in real time 

changes to design and planning was made possible by the concurrent and on-site presence of key 

design, analysis, materials and processes, manufacturing and quality personnel. The resulting 

manufacturing flow for the CCM test article is shown in Figure 10.0-1. 
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Figure 10.0-1. CCM Test Article Manufacturing Flow Chart 

The CCM Manufacturing Development Team (MDT) worked side-by-side with Design, 

Materials and Processes, and NDE Engineering to support producibility and inspectibility of the 

CCM. Additionally, the MDT developed the planning and processes for the full-scale article 

fabrication, assembly and inspection as part of an overall concurrent engineering activity to drive 

the project in the shortest time possible. The MDT was led by NASA and chartered to 

concurrently develop baseline manufacturing planning in coordination with Design and Analysis, 

Upper Shell Lower Shell

Lay-up and Cure 

IML Skin

Install external 

SM/ALAS brackets

Cut Openings and 

Bond Frames

Install and Cure 

Gusset Caps

Install LIDS Ring

Install Parachute 

Brackets

Install and Cure 

Gusset Panels

Trim and Extract 

from Mold

Lay-up and Cure 

OML Skin

Install Core and 

Cure Adhesive

Match drill all 

SM/ALAS Brackets 

and install inner

Install and Cure 

Backbone Cap

Install and Cure 

Backbone Assy

Trim and Extract 

from Mold

Lay-up and Cure 

OML Skin

Install Core and 

Cure Adhesive

Lay-up and Cure 

IML Skin

Lay-up, Cure and 

machine Backbone 

Panels & Cap

Assemble and Cure 

Backbone

Lay-up, Cure and 

Machine Gusset 

Panels & Caps

Pattern Cut OML

Form, Machine 

and Prep Core

Pattern Cut IML

Pack and Ship

Lay-up and Cure 

Splice

Bond Upper and 

Lower Shells
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NDE, Quality, Test and CCM team leadership. The MDT was comprised of the key industry 

experts, each selected to ensure success of critical and challenging aspects of the project. 

• Northrop Grumman – Composite manufacturing, assembly and inspection 

• Lockheed Martin – Pi-preform process technology 

• ATK Aerospace Structures – Composite and metal manufacturing, assembly, inspection 

and special tooling 

• Janicki Industries – CCM special tooling and composite lay-up molds 

• Alcore – Aluminum Honeycomb Core processing 

• Tayco Engineering – Out-of-Autoclave cure processing 

Manufacture of the CCM pathfinder and test article was conducted at the ATK Aerospace 

Structures manufacturing facility in Iuka. Building block, test panel and component 

manufacturing was supported by the National Center for Advanced Manufacturing at MSFC in 

Huntsville, Alabama. 

11.0 NDE Overview 

NDE is a noninvasive measurement science and, for CCM, was selected using several factors 

such as: measurement system maturity, hardware-specific flaw criteria (critical flaw size 

detectability), local part geometry, thickness, surface condition, and access to the inspection 

surface. Compound curvature and unique geometry part inspection required some tailored NDE 

techniques and specialized NDE/inspection tools, which are discussed further in the CCM NDE 

final report. 

The following four NDE methods were approved for use on CCM. Figure 11.0-1 shows the 

primary NDE methods by component-type on the CCM structure: 

1. Visual

2. Ultrasound

3. Thermography 

4. Radiography
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Figure 11.0-1. CCM Primary NDE Methods by Component Type 

Composite manufacturing-induced defects such as voids and disbonds, and damage induced 

during manufacturing, processing and handling can lead to premature and/or unexpected failure 

of the hardware. Therefore it was essential to the CCM project that the appropriate NDE 

methods were carefully selected. Inspection methodologies such as ultrasound, infrared (IR) 

thermography, and radiography were chosen as the primary NDE methods because of their 

maturity, portability and field proven reliability. 

Inspectability and damage tolerance were two of several key-driving requirements identified 

during the design process that influenced some of the project decisions during the early design 

trades. A comprehensive inspection goal was established during the early phase of the project 

that included the inspection of all piece-parts, primary and secondary structure. However, NDE 

was limited during certain phases of the build process by access restrictions after populating the 

cabin, geometric constraints, etc. The inspection and flaw detection capability was dependent on 

several factors such as the accessibility of the region of interest, the NDE system setup logistics 

and acquisition requirements, and the CCM part geometry, thickness, surface condition, etc. 

The inspection feasibility for common defect types identified for the CCM project is shown in 

Table 11.0-1. Composite defects encountered on the CCM project included:

• Delaminations 

• Debonds/unbonds (e.g., core-to-facesheet unbond, missing adhesive, fiber-to-matrix 

bond)
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• Core discontinuities (e.g., tear, buckle, crush)

• Matrix porosity / cracking 

• Excess/insufficient resin 

• Fiber breaks

• Impact damage 

• Inclusions (adhesive backing peel ply, moisture) 

• Missing material 

Table 11.0-1. NDE Methods for Common Composite Defect Types 

The NDE team recommended the use of multiple NDE methods (where applicable) for the first 

article production unit and subsequent test articles. All candidate commercial-off-the-shelf NDE 

systems met or exceeded the established minimum detectable flaw criteria that were established 

for the CCM project. The primary NDE methods are identified by component type in 

Table 11.0-2, and the advantages and disadvantages for the use of each NDE method are detailed 

in Section 5.2. 
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Table 11.0-2. NDE Methods Identified by Hardware / Component Type 

All inspections were performed per CCM-SPEC-006, Rev A (which defines defect acceptance 

criteria) for the following configurations: 

a. Facesheets / Flat Laminates (Laminate caps, Doublers) 

• Maximum defect width / diameter was < 0.250-inch 

b. Sandwich Facesheet-to-Core Bonds

• Maximum defect width / diameter was < 0.50-inch (2 x cell size minimum) 

c. Pi-Preform-to-Laminate Bonds 

• The maximum permissible bondline defect width / diameter was < 0.125 inch 

d. Core-to-Core Bonds

• Foaming adhesive gaps not to exceed 0.06-inch in length (direction parallel to the 

splice) 

e. Core-to-Insert Bonds 

• Defects or adhesive voids cannot span the entire thickness of the panel (i.e., inner 

mold line (IML) skin to outer mold line (OML) skin of a sandwich panel) 

NDE was performed during three key phases in the project timeline: 1) test article fabrication 

and developmental testing, 2) full-scale manufacturing, and 3) full-scale testing. The final 

inspection series occurred after completion of full-scale testing to assess the final condition of 

the CCM test article. 

Inspection data was presented to the team during regular team meetings or during special root 

cause/failure investigation meetings. The NDE team was a multi-center effort with specific 

expertise (i.e., ultrasound, thermography, etc.) usually residing at a particular center. Therefore, 
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a standardized NDE reporting format was not adopted by the project. Each NDE organization 

was permitted to acquire post-process and present data and conclusions based on their centerspecific best practice. However, NDE data was consistently presented in a manner that 

emphasized any critical conditions or features, such as defect-like anomalies, locations, depths, 

and sizes of pre-existing / known damages, and other pertinent data that would be required for an 

engineering review board disposition of a defect. Inspections performed by the CCM NDE team 

are archived in the ICE / Windchill CCM project folder. 

A detailed discussion of the NDE for the CCM project is provided in the NESC 06-019 

Composite Crew Module: NDE Report. 

12.0 Test Overview 

For the design and manufacturing of the CCM, the team adopted the building block approach 

where design and manufacturing risks were mitigated through manufacturing trials and structural 

testing at various levels of complexity. Because of time constraints, the team was required to 

execute the building block approach in a quasi-parallel. 

The primary objectives of the test project were to: 

(a) Reduce project risk through the building block approach. 

(b) Provide data for analysis verification in areas of structural complexity (areas around 

fittings, core ramps, joints, and inserts). 

(c) Demonstrate structural integrity under design limit and ultimate load conditions. 

The CCM Test report contains information on all element, component, and full-scale level 

testing. For element and component level tests the analysis / test correlation is presented in the 

NESC 06-19 Composite Crew Module: Test Report. However, for the full-scale tests, the 

analysis and analysis / test correlation are presented in the Analysis report. Coupon level testing 

and element tests, which were associated with material development are presented in the NESC 

06-019 Composite Crew Module: Materials and Processes Report. 

Structural details that were selected for evaluation at the element level included joints 

(composite-to-composite and composite-to-metal), aluminum core taper, and metal inserts. At 

the component level a simplified version of a highly loaded fitting SM / ALAS was designed and 

tested. Element tests were useful in verifying both the fabrication processes and the analytical 

techniques. It was found that detailed pre-test analysis was crucial to the success of the element 

tests. 

Full-scale tests included internal pressure, combination of pressure and mechanical load at 

specific hard points such as the SM / ALAS, main parachute fitting, and mechanical load 

application at the drogue chute fitting, and inserts. 

Full-scale testing was accomplished through the design of a multifunction self-reacting test 

frame, which served as the backbone for all full-scale load applications. CCM testing was done 

in an upside down configuration so that installation of loading hardware could be done at the 
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working level, and the risk of damage from dropped hardware was reduced. In addition, the test 

frame served as the skeleton of the shipping container. The test frame was built in Iuka and was 

delivered to NASA LaRC along with the full-scale CM for testing. Full-scale testing took place 

at LaRC’s Combined Loads Test System (COLTS) facility. Figure 12.0-1 is an image of the 

CCM within the shipping and testing frame at the LaRC COLTS facility. 

Figure 12.0-1. CCM in Upside Down Configuration within the Self Reacting Load Frame in LaRC’s 

COLTS 

A much more detailed discussion of the test activity for the CCM project is provided in the 

NESC 06-019 Composite Crew Module: Test Report. 

13.0 Orion Comparative Study 

In September 2008, the Orion Project Manager received an action from the NASA Administrator 

to evaluate the implications of incorporating the CCM primary structure into the Orion main line 

project as part of the CxP. A technical study ensued from September to November 2008 to 

characterize the implications. 
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13.1 Mass Differences between CCM and Orion 

Since the CCM project had baselined loads and interfaces in March 2007, one of the biggest 

challenges in evaluating the implications of a CCM for Orion was the gap between the CCM 

configuration and the then current Orion configuration. CCM was at least two load cycles 

behind Orion; some load cases had increased, such as water landing loads, and some had 

decreased, such as parachute and drogue loads. In addition, CCM had baselined the LAS 

interface low on the cylindrical section of the pressure shell at the SM interface. Orion had 

baselined the LAS interface at the main parachute interface. A goal of the project was to obtain 

data as quickly as possible and did not allow for a design cycle on the CCM. Therefore, an 

accurate direct mass comparison was not possible. 

However, first order materials comparisons were made that would suggest that mass between the 

CCM and the Al-Li Orion would be similar. The CCM was strength driven rather than stiffness 

driven primary structure. The majority of the structure was sized by multiaxial loads caused by 

internal pressure. The CCM will also be exposed to possible damage during spacecraft 

integration and during the mission, so damage tolerance knockdowns were used to accommodate 

for possible impact damage. These factors combined strength rather than stiffness, multiaxial 

loading, damage tolerant material properties, and limits the performance benefit a composite 

material system might traditionally offer to a design since strength per unit mass of damage 

tolerant, quasi isotropic composite materials systems are similar to Al-Li properties. Therefore, 

there were no expected significant mass differences between the CCM and the Orion Al-Li CM. 

13.2 Manufacturability Differences between CCM and Orion 

The CCM was built in two halves, an upper and lower, with system assembly performed on both 

halves before joining the upper to the lower. The major manufacturing operations of the CCM 

are:

1. CCM Upper 

1.1. Layup and cure IML skin 

1.2. Cure Honeycomb core to IML skin 

1.3. Layup and cure OML skin 

1.4. Install Docking Tunnel 

1.5. Install Gussets 

1.6. Install Gusset caps 

1.7. Install main parachute fittings 

1.8. Machine window and hatch cutouts 

1.9. Install window and hatch frames 
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2. CCM Lower 

2.1. Layup and cure IML skin 

2.2. Install and cure Honeycomb core 

2.3. Layup and cure OML skin 

2.4. Layup and cure backbone panels 

2.5. Assemble backbone panels 

2.6. Install and cure backbone to lower shell 

2.7. Layup and cure backbone cap 

2.8. Install backbone in lower shell assembly 

2.9. Install metallic interface fittings on Lower shell 

3. CCM Assembly 

3.1. Splice longeron regions of assembly 

3.2. Splice acreage regions of assembly 

The Orion design was especially different in operations. The lower barrel, aft dome, ceiling, and 

docking tunnel were each one piece components machined to final dimension in a computer 

controlled machining center. The Orion conic was fabricated from six segments with hatch, 

windows, and longerons integrally machined. The individual components were assembled using 

automated friction stir welding techniques with built in inspection. 

The CCM approach could be improved through the use of automated fiber placement. However, 

this would only improve the layup and cure of the upper and lower sandwich systems, which 

represented only 2 months of the 9 month assembly process. 

It was determined by both the Orion and CCM teams that manufacturability comparisons yielded 

no measureable benefit between the two different material systems. 

One exception was the use of the assembly joint between the upper and lower pressure shell in 

the CCM. This joint was intended to offer potential benefits during subsystem integration as part 

of the overall spacecraft integration process. The Orion team developed a similar clam shell 

approach for their design, but it was later discarded as not being advantageous in the overall 

system integration flow. 

13.3 Other Technology Gaps 

In addition to the mass and manufacturing comparisons, there were numerous technology gaps 

identified by the Orion team with respect to adopting the CCM design. These technology gaps 

represented a risk to the Orion project. Given that CCM did not provide a quantifiable
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advantage, and represented numerous additional risks, the decision was made not to incorporate 

the CCM into the Orion baseline. 

Since completion of the study in November 2008, some of the technology gaps have evolved 

and, in some cases, have been retired. 

13.3.1 Damage Tolerance for Leak Integrity 

The CCM leak integrity approach was to use predominantly sandwich systems with unvented 

core. The approach provided two barriers against through thickness permeability through the use 

of the IML skin and the OML skin. However, it was shown that exceedingly low impact 

damage, less than 2 ft/lb, could cause leakage and would not be detectable visibly or with NDE 

techniques. Given the extensive subsystem integration both on the inside and the outside of the 

CCM, there was a possibility of undetected damage coincident on the inside and outside that 

could provide an undetected leak path. 

Some damage tolerant liners were studied during the CCM project and were shown to prevent 

leakage even after impact on biaxially loaded test coupons. However, this work was limited to 

coupon testing and was not demonstrated on the full-scale assembly. This work is discussed in 

more detail in the NESC 06-019 Composite Crew Module: Materials and Processes Report. 

13.3.2 Damage Tolerance as Related to Structural Integrity

At the time of the Orion study, little if any damage tolerance work had been done on the CCM 

project except for the use of damage tolerant material allowable. However, the subsequent test 

project included numerous damage tolerant evaluation tests, including the effect of 6 ft/lb 

impacts at 18 different locations followed by inspection, testing to ultimate design levels, and 

then cyclic testing through four life times. Additional impact sites were impacted to 26 ft/lb and 

inspection, testing to ultimate design levels, and then cyclic testing through four life times and 

were repeated. In all cases, no detrimental growth of damage was shown. 

13.3.3 Inspection of CCM Pressure Vessel 

Inspections of the CCM pressure vessel were closely tied to the damage tolerance program for 

the structure. On CCM, impact damage was induced, measured visually, and then measured 

using different state-of-the-art NDE techniques. As mentioned earlier, in all cases, no 

detrimental growth of damage was shown. 

On CCM, there are bonded joints in the primary load path that cannot be verified acceptable by 

inspection, particularly the docking tunnel interface ring and the main assembly splice joint.

Since these joints are process-dependent, it was expected that some level of acceptance testing 

would be necessary for each assembly prior to flight. The team envisioned some level of proof 

test for the acceptance test, but did not finalize a position on this issue. 
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13.3.4 Splice Joint Feasibility 

During the Orion study, the viability of the out-of-autoclave splice joint was in question because 

the operation had yet to be performed on the CCM project. There was particular concern about 

the manufacturing tolerance, leakage, structural integrity, and out-of-autoclave approach. As 

demonstrated by the final assembly and test results, the manufacturing tolerance, structural 

integrity, and out-of-autoclave technique were validated. The leakage, or more accurately, 

permeability in the out-of-autoclave material system, was managed by using a material system 

appropriate for out-of-autoclave curing, or by the use of a secondary barrier as discussed in 

Section 13.3.1. 

13.3.5 Repair Techniques for Pressure Containment and Structural Integrity 

CCM repair occurred numerous times during the development of the manufacturing process. 

Repair techniques were applied to correct for manufacturing anomalies such as tooling shift 

during cure of Pi-preforms, improper cure of splice doublers, excess void in docking tunnel 

doublers, and removal and replacement of tunnel gusset Pi joints following disbond from over 

constrained tooling. Additionally, repairs were made following impact damage caused by tool 

drop. These repairs are addressed more comprehensively in the NESC 06-019 Composite Crew 

Module: Materials and Processes Report 

13.3.6 Verification and Certification 

Verification and certification of composite primary structure for human spaceflight was not in 

the scope of the CCM project.

13.3.7 Material System Solution 

Technology gaps associated with space environments and mission duration impacts and 

compatibility were cited as well as with the unvented core approach to leak protection versus the 

lunar mission requirements for leakage. These gaps are well within NASA’s knowledge base 

and are not viewed as major risks for future composite spacecraft. 

13.3.8 Penetrations to the Pressure Shell – Maintaining Leak Integrity 

The CCM contains fastener penetrations through the pressure shell. The design baselined the use 

of polysulfide sealant, wet installed with the fasteners, as the method for managing leak integrity 

around the penetrations. The polysulfide passed many of NASA test Standard 6001 for 

compatibility, but represented an odor problem. It is believed that an alternative that preserves 

the advantages of polysulfide, but without the obnoxious odor, could be identified and 

demonstrated if a composite pressure shell was adopted for a human spaceflight program. 
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14.0 Findings and NESC Recommendations 

14.1 Findings 

The following findings were identified: 

F-1. Many of the design, analysis, materials, and manufacturing lessons learned were not 

evident in the coupon, element, or subcomponent testing and only became evident when 

building the full-scale assembly. 

F-2. Small agile teams—emphasizing “in-line” work—working in parallel with the flight 

project can achieve significant results through rapid decision making. These teams can 

exploit unforeseen opportunities and minimize consequence of risk with quick recovery 

to problems that occur. 

F-3. Project collocations served as a valuable tool for developing relationships among team 

members and for achieving significant technical results quickly. 

F-4. Frequent telecoms, WebExTM, instant messaging, and desktop sharing were critical tools 

enabling virtual project management when the team was not collocated. 

F-5. Critical path project tools were the key method for managing overall project priorities, 

schedule performance, and cost performance. 

F-6. Small, but focused, independent technical reviews with subject matter experts were 

effective for challenging team assumptions and technical decisions resulting in increased 

innovation and a higher motivated team. 

F-7. Overall travel expenditures were insignificant expenses relative to the project value even 

though significant time was spent in team collocations throughout the design, 

development, manufacturing, and test phases. 

F-8. Through thickness leakage or permeability, there remains an area for demonstrating 

solutions at the full-scale complex shape level, especially in applications exposed to 

possible impact damage. 

14.2 NESC Recommendations 

The following NESC recommendations were identified and directed towards the ESMD unless 

otherwise identified: 

R-1. It is recommended that Programs/Projects consider implementing full-scale 

manufacturing development early in the conceptual design phase so that manufacturing 

lessons can be incorporated as the design evolves. (F-1)

R-2. It is recommended that future Human Spaceflight Programs demonstrate permeability 

and leakage solutions on complex polymeric composite components such as the CCM.

(F-8)
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15.0 Alternate Viewpoints 

There were no alternate viewpoints identified during the course of this assessment by the NESC 

team or the NRB quorum. 

16.0 Other Deliverables 

No unique hardware, software, or data packages, outside those contained in this report, were 

disseminated to other parties outside this assessment. 

17.0 Lessons Learned 

LL-1. Small agile teams—emphasizing “in-line” work—working in parallel with, but outside 

the umbrella of a flight project can achieve significant results through rapid decision 

making. These teams can exploit unforeseen opportunities and minimize consequence of 

risk with quick recovery to problems that may occur.

LL-2. Engineering models predicted mass and structural response well. 

LL-3. Engineering models did not always predict production challenges. 

18.0 Definition of Terms 

Corrective Actions Changes to design processes, work instructions, workmanship practices, 

training, inspections, tests, procedures, specifications, drawings, tools, 

equipment, facilities, resources, or material that result in preventing, 

minimizing, or limiting the potential for recurrence of a problem. 

Finding A conclusion based on facts established by the investigating authority. 

Lessons Learned Knowledge or understanding gained by experience. The experience may 

be positive, as in a successful test or mission, or negative, as in a mishap 

or failure. A lesson must be significant in that it has real or assumed 

impact on operations; valid in that it is factually and technically correct; 

and applicable in that it identifies a specific design, process, or decision 

that reduces or limits the potential for failures and mishaps, or reinforces a 

positive result. 

Observation A factor, event, or circumstance identified during the assessment that did 

not contribute to the problem, but if left uncorrected has the potential to 

cause a mishap, injury, or increase the severity should a mishap occur. 

Alternatively, an observation could be a positive acknowledgement of a 

Center/Program/Project/Organization’s operational structure, tools, and/or 

support provided. 
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Problem The subject of the independent technical assessment. 

Proximate Cause The event(s) that occurred, including any condition(s) that existed 

immediately before the undesired outcome, directly resulted in its 

occurrence and, if eliminated or modified, would have prevented the 

undesired outcome. 

Recommendation An action identified by the NESC to correct a root cause or deficiency 

identified during the investigation. The recommendations may be used by 

the responsible Center/Program/Project/Organization in the preparation of 

a corrective action plan. 

Root Cause One of multiple factors (events, conditions, or organizational factors) that 

contributed to or created the proximate cause and subsequent undesired 

outcome and, if eliminated or modified, would have prevented the 

undesired outcome. Typically, multiple root causes contribute to an 

undesired outcome. 

19.0 Acronyms List 

3D Three Dimensional 

AA Associate Administrator 

AC Actual Cost 

ALAS Alternate Launch Abort System 

Al-Li Aluminum Lithium 

ATK Alliant Techsystems, Inc. 

CAD Computer-Aided Design 

CCM Composite Crew Module 

CEV Crew Exploration Vehicle 

CM Crew Module 

COLTS Combined Loads Test System 

CxP Constellation Program 

DCO Document Change Order 

EAC Estimate to Complete 

EO Engineering Order 

ESMD Exploration Systems Mission Directorate 

ETC Estimate to Complete 

EV Earned Value 

EVM Earned Value Management 

FTE Full-Time Equivalent 

GSFC Goddard Space Flight Center 

ICE Integrated Collaborative Environment 

IML Inner Mold Line 
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IR Infrared 

ITAR International Traffic In Arms Regulations 

JSC Johnson Space Center 

KSC Kennedy Space Center 

LaRC Langley Research Center 

LAS Launch Abort System 

lbs pounds 

MDT Manufacturing Development Team

MIL-STD Military Standard 

MSFC Marshall Space Flight Center 

NASA National Aeronautics and Space Administration 

NASTRAN® NASA Structural Analysis System 

NDE Non-Destructive Evaluation 

NESC NASA Engineering and Safety Center 

NRB NESC Review Board 

OML Outer Mold Line 

Pro-E® Pro Engineering®

PTC Parametric Technology Corporation 

PV Planned Value 

SBT Smart Buyer Team 

SM Service Module 

WYE Work Year Equivalent 

20.0 References 

1. Oscar Hoffman, The Brittle Strength of Orthotropic Materials, Journal of Composite 

Materials, April 1967, pp 200-206.

21.0 Appendices 

Appendix A. Major Project Milestones 

Appendix B. CCM Team List 
NASA Engineering and Safety Center 

Technical Assessment Report

Document #: 

NESC-RP06-019

Version: 

1.0

Title: 

Composite Crew Module Primary Structure 

Page #: 

48 of 52 

NESC Request No.: 06-019-I (Primary Structure)

Appendix A. Major Project Milestones
NASA Engineering and Safety Center 

Technical Assessment Report

Document #: 

NESC-RP06-019

Version: 

1.0

Title: 

Composite Crew Module Primary Structure 

Page #: 

49 of 52 

NESC Request No.: 06-019-I (Primary Structure)

Appendix B. CCM Team List
NASA Engineering and Safety Center 

Technical Assessment Report

Document #: 

NESC-RP06-019

Version: 

1.0

Title: 

Composite Crew Module Primary Structure 

Page #: 

50 of 52 

NESC Request No.: 06-019-I (Primary Structure)
NASA Engineering and Safety Center 

Technical Assessment Report

Document #: 

NESC-RP06-019

Version: 

1.0

Title: 

Composite Crew Module Primary Structure 

Page #: 

51 of 52 

NESC Request No.: 06-019-I (Primary Structure)
NASA Engineering and Safety Center 

Technical Assessment Report

Document #: 

NESC-RP06-019

Version: 

1.0

Title: 

Composite Crew Module Primary Structure 

Page #: 

52 of 52 

NESC Request No.: 06-019-I (Primary Structure)
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

2. REPORT TYPE 

Technical Memorandum

 4. TITLE AND SUBTITLE

Composite Crew Module: Primary Structure

5a. CONTRACT NUMBER

 6. AUTHOR(S)

Kirsch, Michael T.

 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-20085 NESC-RP-06-019

10. SPONSOR/MONITOR'S ACRONYM(S)

NASA

13. SUPPLEMENTARY NOTES

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Category 16 Space Transportation and Safety

Availability: NASA CASI (443) 757-5802

19a. NAME OF RESPONSIBLE PERSON

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

14. ABSTRACT

In January 2007, the NASA Administrator and Associate Administrator for the Exploration Systems Mission Directorate 

chartered the NASA Engineering and Safety Center to design, build, and test a full-scale crew module primary structure, using 

carbon fiber reinforced epoxy based composite materials. The overall goal of the Composite Crew Module project was to 

develop a team from the NASA family with hands-on experience in composite design, manufacturing, and testing in 

anticipation of future space exploration systems being made of composite materials. The CCM project was planned to run 

concurrently with the Orion project's baseline metallic design within the Constellation Program so that features could be 

compared and discussed without inducing risk to the overall Program. This report discusses the project management aspects of 

the project including team organization, decision making, independent technical reviews, and cost and schedule management 

approach.

15. SUBJECT TERMS

Composite Crew Module; NASA Engineering and Safety Center; Alternate Launch Abort System; Exploration Systems 

Mission Directorate

18. NUMBER

 OF 

 PAGES

57

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)

January 2007 - May 2011

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

869021.05.07.07.03

11. SPONSOR/MONITOR'S REPORT

 NUMBER(S)

NASA/TM-2011-217185

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PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS.

1. REPORT DATE (DD-MM-YYYY)

01- 11 - 2011