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Team Gains Experience as it Builds Innovative Composite Spacecraft
 
Even as NASA’s Constellation program is designing the Orion crew module that will carry astronauts back to the moon and beyond, another NASA team is exploring alternate materials and fabrication techniques that promise even more advances for future spacecraft. Aerospace composites, not unlike the materials used in open-wheeled racecars and high-end sporting equipment, are lightweight and tough. And - this fiber and resin-filled material can be fabricated into complex shapes with significantly fewer tools and parts with much less labor than conventional metal-based construction techniques.

Of course, if it were easy to do, it would have already been done. Using new technologies to build a space-rated pressure vessel to safely carry humans beyond earth orbit is a technical challenge that has yet to be demonstrated.

That’s why the NASA Engineering and Safety Center (NESC) is building a composite crew module. NASA’s Orion crew module was picked as the pattern for the composites demo project, in part, to take advantage of the wealth of engineering data already developed by the Constellation program for NASA’s next crewed spacecraft. The NESC composites team has taken the basic size, shape, and other design elements of Orion, but is building a highly-modified version as a high-tech technology demonstrator.

When the NESC composites team is done, it will have gained hands-on design, build, and test experience in anticipation that future exploration systems may be made of composite materials.

Here are project details and progress to date.

Problem:
In 2006, the NESC studied the feasibility of a composite crew module for the Crew Exploration Vehicle. The overall finding indicated that a composite crew module was feasible but that a detailed design would be necessary to quantify technical characteristics, particularly in the area of mass and manufacturability. Accordingly, the NASA Administrator, Associate Administrator for Exploration, and the Constellation Program manager chartered the NESC to design, build, and test a composite crew module with the goal of developing a network of engineers within the agency with hands on experience using composites on a habitable spacecraft design.

NESC Contribution:
The NESC Composite Crew Module Project Objective is to design, build, and test a structural test article of the Crew Exploration Vehicle Crew Module primary structure. The project was chartered in January of 2007, with a goal of delivering a test article for structural testing in July of 2008, 18 months after project initiation. The project team is a partnership between NASA and industry and includes design, manufacturing, and tooling expertise. Partners include civil servants from ARC, DFRC, GRC, GSFC, JSC, JPL, KSC, LaRC, MSFC, Air Force Research Laboratories and contractors from Alcore, Alliant Technologies, Bally Ribbon Mills, Collier Corporation, Genesis Engineering, Janicki Industries, Lockheed Martin, and Northrop Grumman.

Result:
The composite crew module team operates in a predominantly virtual environment electronically connecting participants across the geographic span. During the design phase, the team co-located at GSFC for 1 or 2 weeks approximately every 6 to 8 weeks. The team constrained the design to match interfaces with the current Orion crew module including the internal packaging constraints that utilize a backbone for securing internal components. During the first 2 months of the effort, the team evaluated design solutions and focused in on a design that utilizes predominantly aluminum honeycomb sandwich and solid laminate material systems. One unique feature of the composite crew module design was the integration of the packaging backbone structurally with the floor and walls of the pressure shell. This provides a load path that accommodates load sharing with the heatshield for water landing load cases. Another unique feature of the composite design is the use of lobes between the webs of the backbone. This feature puts the floor into a membrane type loading resulting in a lower mass solution. Connecting the floor to the backbone, and placing lobes into the floor resulted in mass savings of approximately 150 lbs to the overall primary structure design. The design is constructed in two primary parts - an upper and lower pressure shell. The two halves are joined together in an out of autoclave process to enable subsystem packaging of large or complex subsystems if deemed beneficial by the project. The initial design concepts were reviewed by an independent review panel in March 2007, followed by another independent review of the preliminary design in June 2007. The project conducted building block testing of critical areas and the results of the building block testing were used to validate critical assumptions for the final design. The detailed design was reviewed again in December 2007 by an independent review panel and based on the review, proceeded into the manufacturing phase of the project. A manufacturing readiness review was conducted in May 2008 and provided independent feedback on the detailed manufacturing instructions for the various components of the design. Fabrication of the upper and lower pressure shells began in February 2008, and post cure assembly operations started in May 2008. The current project plan is to statically test the upper and lower shells individually; to verify that the analysis models predict the response of the structure under load; and then to join the two halves together and repeat the static tests with internal pressure. Testing on the upper shell is currently scheduled to occur at LaRC in October 2008.

Findings to date

Design Findings
  • Non-autoclave splice allows concurrent fabrication, assembly, and integration of major structural components and subsystems and provides lower cost cure tooling option
  • Membrane lobed floor integrated with backbone subsystem packaging feature offer weight savings (~ 150lbs) through complex shapes enabled by composites
  • State of the art Pi-preforms leveraged from DoD offer robust orthogonal composite joints
  • Honeycomb design approach combined with mature secondary attachment technology provides flexibility and robustness in secondary attachment locations
  • Inner mold line tooling offers opportunity to optimize or change design through tailoring of layups or core density, as loads and environments change with program maturation
  • Composite solutions offer lower part count resulting in a lower drawing count (~47) which helps reduce overall life cycle costs
  • Estimated mass based on March 2007 loads, environments, and interfaces:
  • PDR (20% WGA) = 291 lbs (metallics) + 1149 lbs (composites) = 1441 lbs total
  • CDR (12% WGA) = 386 lbs (metallics) + 944 lbs (composites) = 1330 lbs total
  • MR (~4 % WGA) = 459 lbs (metallics) + 1005 lbs (composites) = 1464 lbs total
  • Note: at MR - margins were increased in metallics to push test article failures to composites

Materials Findings
  • Leveraged existing material databases, by using IM7/977-2 fiber and resin, a mature material system with extensive government and industry experience
  • Building block program accelerated by emphasis on element level testing over materials level (lamina/laminate)
  • Flight qualification gap not expected to change design
  • Backfill materials level coupon data to substantiate literature generated design allowable and design rules
  • Additional element level testing to validate environmental effects and off-nominal manufacturing processes
  • Demonstrate life with credible damage and off nominal configuration per MSFC 3479

Analysis Findings
  • Comprehensive analysis of entire structure
  • Numerous (>160) load cases that envelope March 2007 flight environment
  • Numerous (>15) analytical models using various modeling techniques with overlaps to verify results
  • Element testing confirmed failure mode and failure load predictions
  • Thermal and dynamic differences from aluminum being investigated; preliminary estimates do not indicate that composite create any system level issues
  • Schedule constraints required extensive use of engineering experience from other flight vehicles backed up by detailed analysis

Manufacturing and Tooling Findings
  • Integrated electronic process flow exists between Design, Manufacturing, and Inspection via CAD, FiberSIM, flat patterns, laser projection, layup, inspection
  • Proven mature composite manufacturing processes using autoclave, oven, hand layup, with composite tooling
  • Part specific process development was applied to manage manufacturing risk
  • Minimal (< 44) cataloged tools for the entire composite crew module primary structure
  • Total touch labor for manufacturing engineering test unit of primary structure ~6000 hours
  • Affordable (<$1M) and quick (< 5 months) state of the art, precision, autoclave, multi-cycle cure tooling for full scale pressure shell

Non-Destructive Examination Findings
  • Mature commercially available inspection equipment; IR thermography, ultrasound, and X-ray
  • 100% non-destructive inspection currently expected for primary structure
  • Documented inspection criteria consistent with material, analysis, and design assumptions

Test and Evaluation Findings
  • All element tests needed to preserve critical path have been completed
  • Additional tests to be conducted as needed to verify manufacturing processes
  • Component test fixture, design, analysis completed with fabrication underway
  • Full scale test design and analysis underway

Image showing that Pressure module = Upper pressure shell + splice joint + lower pressure shell


Image showing Cutaway of Pressure Module Showing Backbone internal structure Cutaway of Pressure Module Showing Backbone internal structure


Cutaway of Pressure Module Showing Backbone internal structure


Image showing Upper Pressure Shell cure tool and Tool Proof and Janicki Industries Upper Pressure Shell cure tool and Tool Proof and Janicki Industries


Image showing Upper Pressure Shell IML layup, bagging, and Autoclave cure at ATK in Iuka, MS Upper Pressure Shell IML layup, bagging, and Autoclave cure at ATK in Iuka, MS


Image showing IR Thermagraphy of inner sandwich skin at ATK in Iuka, MS IR Thermagraphy of inner sandwich skin at ATK in Iuka, MS


Image showing Installation of Film Adhesive on inner skin for core
Installation of Film Adhesive on inner skin for core



Image showing Installation of honeycomb core on Inner sandwich skin
Installation of honeycomb core on Inner sandwich skin



Image showing installation of outer sandwich skin
Installation of outer sandwich skin



Image showing Cured Upper Pressure Shell Sandwich at ATK in Iuka, MS Cured Upper Pressure Shell Sandwich at ATK in Iuka, MS


Image showing Automated Ultrasonic Inspection of Upper Pressure Shell Sandwich at ATK in Iuka, MS Automated Ultrasonic Inspection of Upper Pressure Shell Sandwich at ATK in Iuka, MS


Image showing Cured inner sandwich skin of Lower Pressure Shell at ATK in Iuka, MS Cured inner sandwich skin of Lower Pressure Shell at ATK in Iuka, MS


Image showing Engineering Model of planned static test fixture for Upper Pressure Shell. The same test fixture is planned to be used for the pressure module assembly. Engineering Model of planned static test fixture for Upper Pressure Shell. The same test fixture is planned to be used for the pressure module assembly.