Optimizing the Ares V Payload Shroud

Sizing software allows rapid analysis of candidate composite architecture for highly loaded and weight-sensitive launch vehicle component.

Design Results

The former Constellation mission replaced the aging Space Shuttle program. Notably, the program marked the first time that composites were considered for manned mission hardware.

NASA designed and built a full-scale Composite Crew Module (CCM) in parallel with the mostly metallic Orion crew exploration capsule that was to fly as part of Constellation. In addition, the payload shroud, interstage and intertank structures of the Ares V heavy-lift launch vehicle, also part of the Constellation concept, were conceptually designed in composites.

The Ares V is a three-stage rocket: Like the Space Shuttle, it was designed to use a pair of solid-fuel first-stage rocket boosters that burn simultaneously with the liquid-fuel second (core) stage. The upper stage, dubbed the Earth Departure Stage (EDS), is powered by a liquid hydrogen/liquid oxygen rocket engine. At 116m/380 ft long and 10m/33 ft in diameter, Ares V would be the world’s largest launch vehicle. It would also be the most powerful, able to lift 188,000 kg/415,000 lb into low earth orbit, or 71,000 kg/157,000 lb to the Moon.

Design of the Ares V components were facilitated by Collier Research, using its HyperSizer software. First developed in the early 1990s at NASA’s Langley Research Center (Hampton, Virginia), HyperSizer does not perform finite element analysis (FEA) and is not a computer-aided design (CAD) program, stresses Craig Collier of Collier Research Corporation (Hampton, Virginia). “It’s a tool for composite designers that provides automated structural analysis, material selection and design optimization.” The software is designed to help the user select the best composite materials for an application and determine margins of safety, for the lightest possible structure. Collier’s company was part of a NASA team that aimed to determine the optimum structural concepts for the largest composite component on Ares V, the 33-ft/10m-diameter and nearly 72-ft/22m-long payload shroud that surrounded the lunar landing hardware.

Illustration: Karl Reque
Illustration: Karl Reque

Optimizing composite laminates for strength

The shroud was envisioned as a four-petal, ogive-shaped enclosure that, at the appointed time in flight, that would explosively separate along a frangible horizontal joint above the interstage and along four “rails,” jettisoning the four petal segments to release the lunar lander. For this component, Collier and the NASA team first developed a finite element (FE) model, using Nastran and Abaqus FEA software from SIMULIA (Providence, Rhode Island); the model mesh is shown at upper right, with dark lines representing the ring frames and the four petal-separation rails. The external dynamic flight pressures were resolved through the FE model mesh to determine the internal “design-to” loads for the HyperSizer software, explains Collier, noting that the software also works without FEA, as long as design-to loads are known.

The FEA calculated the stress resultants or shell element unit forces in both the x and y directions, and showed that the highest loads were in the axial (compressive) direction on the order of –454 kg/–1,000 lb per inch of vehicle circumference, as expected. But, says Collier, “the higher flight pressures in the nose region caused relatively high internal hoop compression [Ny] loads, which are particularly challenging for sizing stiffened panels.” In-plane shear loads (Nxy) for the shroud were not significant.

With the design-to loads established, the team conceived an initial design of composite panels between each ringframe and spanning the distance between the four petal separation rails (the ringframes and rails themselves were envisioned as composite profiles. Panel candidates included cored sandwich panels and stiffened panels, the latter featuring uncored, monolithic laminates with hat stiffeners bonded or comolded with the panel to prevent buckling. HyperSizer provided a means to quickly establish the best — i.e., the lightest — panel concept.

Shell Element Mesh
A shell element mesh was developed so that finite element analysis (FEA) could resolve the flight loads on the payload shroud. The bold lines represent the ringframes and petal-separation rails. Source: Collier Research

The software’s Material Manager module allowed the designers to select from an integrated database of specific composite materials, which included some of the materials found in the Composites Material Handbook (CMH-17), but also permitted them to import specific, user-defined materials and custom layups. The designers built composite laminates, systematically stacking different material forms or types simply by using Windows’ cut, paste and copy functions for quick ply insertions and practical layup arrangements. HyperSizer’s analytical modules (HyperSizer Basic and Pro) subjected the material layups to detailed stress analyses, using hundreds of failure modes.

Collier pointed out that the team considered ply-based and laminate-based approaches for optimization. Essentially, ply-based analysis determines the strength of the total laminate by computing the stress/strain in each ply. To do so, each individual ply’s material angle is a critical input into the calculation — typically, engineers assume that the total laminate fails whenever the first ply fails. On the other hand, laminate analyses do not try to establish the stress/strain in each individual ply, but instead calculate the effective stress/strain of the total laminate and compare it to an empirically based allowable curve. “The NASA team chose the ply-based approach to failure analysis. The software allowed either approach.

Sandwich vs. stiffened panel

The designers selected various panel concepts from the materials database, using the cut-and-paste features. These included a two-stack, honeycomb-cored sandwich, with a variety of facesheet materials and core thicknesses, as well as open-section and closed-section stiffeners, with varying fiber architecture in the skin versus the stiffener. Then, for each panel concept, the software rapidly analyzed more than 100 composites-specific failure modes, including flat-wise tension, facesheet wrinkling, intracell dimpling, core shear and crimping for sandwich concepts, and local buckling, crippling, and panel buckling for stiffened panels. For each panel design, the software calculated whether that particular combination of materials and layup could bear the known loads with an acceptable margin of safety, by automatically graphing the failure envelopes and the stress/strain profiles, using standard quadratic failure criteria, such as Tsai-Wu.


When “large acreage” design was employed, that is, large areas of stiffened panels that had identical layups (screen shot on left), the dark purple color indicates that many panels were overdesigned, with margins of safety well over 1.0. The optimized design (on the right) featured tailored layups that met flight specifications but reduced component weight. Source: Collier Research

When the failure results were examined, Collier notes, it became obvious that a design approach in which “large acreage” areas of the shroud were assumed to require identical layups resulted in overdesign, as indicated by the dark purple color in the figure at right. Says Collier, “These areas represent margins of safety well above 1.0, meaning the entire large area must be able to withstand the load of any small, localized stressed area.” In the figure on the right, he explains, ”the lighter colors mean panels have been optimized by taking material out, with ply drops, and with customized materials, for a lighter overall design.”

According to Collier, the optimization process yielded lighter results with stiffened panels because their greater complexity offers more opportunity to save weight, due to the additional sizing variables. For example, the stiffeners can have primarily 0° plies while the skin can be dominated by 45° and 90° fibers. Further, detailed ply-level tailoring is possible for greater weight savings than with a sandwich design: “While sandwich panels do perform well in biaxial compression loading, the core material and associated bonding adhesive weight is parasitic — that is, it doesn’t directly contribute to the panel’s strength, which is a disadvantage.” In contrast, Collier points out, all of the material in a stiffened panel carries load, but laminate thickness and material layups must be precisely designed to prevent compressive load instability in the individual panel segments between the stiffeners.

Optimizing for manufacturability

When the panels were optimized for strength and positive margins of safety were confirmed for all potential failure modes, HyperSizer was used to optimize the entire structure for “ply compatibility.” That is, the software determined not only the best design but the most practical layup as well. “This capability proves manufacturability by determining the most efficient layup sequence, and can help develop programs for automated processes as well,” notes Collier.

The software generates a series of global sublaminate stacks (GSS) within which all plies are continuous throughout the entire panel or part. These are placed on the tool side (either OML or IML). If ply drop-offs or ply-adds are required, they are placed near the laminate midplane, between GSS. This minimizes drop-offs on the tool, making it much easier to manufacture the part.

An example of HyperSizer’s ply compatibility function, where the software optimizes ply drop-offs and adds to arrive at the most manufacturable design. Source: Collier Research
An example of HyperSizer’s ply compatibility function, where the software optimizes ply drop-offs and adds to arrive at the most manufacturable design. Source: Collier Research

The software also prioritizes ply drop-off order, avoiding “stair step” drops that can greatly diminish laminate performance. “Well over a million candidate layup arrangements across each panel surface were evaluated so that the lightest design was also the most manufacturable,” Collier reports.

Finally, HyperSizer’s PC-compatibility enabled multiple team members to run the program on different computers, all with access to the same database. Each engineer’s results were documented and maintained, and complete stress reports were available for later airworthiness documentation.

NASA conducted a rigorous “figure of merit” (FOM) process to score the candidate shroud concepts, with weight savings as the most important metric. If the Ares V shroud had been designed as stiffened panels, the skins would have been thicker than expected for an axially loaded structure: It would have required a high percentage of 90° plies if it was to react the tension hoop load caused by internal pressurization yet also withstand crush pressure. The axial loading during launch requires special attention to prevent panel buckling — the thick skins and stiffener webs with +/-45° plies would counteract such buckling forces.

The fact that composites used for the Ares V Shroud indicates that the technology is maturing. “The composites optimization process for Ares V,” concluded Collier, “has shown potential to reduce the weight of traditionally accepted launch vehicle designs by 30 percent.”