Bart F. Zalewski

ZIN Technologies, Inc., Middleburg Heights, Ohio 

Brett A. Bednarcyk

Glenn Research Center, Cleveland, Ohio

ACT Payload Shroud Structural 

Concept Analysis and Optimization

NASA/TM—2010-216942

December 2010
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Bart F. Zalewski

ZIN Technologies, Inc., Middleburg Heights, Ohio 

Brett A. Bednarcyk

Glenn Research Center, Cleveland, Ohio

ACT Payload Shroud Structural 

Concept Analysis and Optimization

NASA/TM—2010-216942

December 2010

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

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NASA/TM—2010-216942 1

ACT Payload Shroud Structural 

Concept Analysis and Optimization

Bart F. Zalewski

ZIN Technologies, Inc.

Middleburg Heights, Ohio 44130

Brett A. Bednarcyk

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Abstract

Aerospace structural applications demand a weight efficient design to perform in a cost effective 

manner. This is particularly true for launch vehicle structures, where weight is the dominant design driver. 

The design process typically requires many iterations to ensure that a satisfactory minimum weight has 

been obtained. Although metallic structures can be weight efficient, composite structures can provide 

additional weight savings due to their lower density and additional design flexibility. This work presents 

structural analysis and weight optimization of a composite payload shroud for NASA’s Ares V heavy lift 

vehicle. Two concepts, which were previously determined to be efficient for such a structure are 

evaluated: a hat stiffened/corrugated panel and a fiber reinforced foam sandwich panel. A composite 

structural optimization code, HyperSizer, is used to optimize the panel geometry, composite material ply 

orientations, and sandwich core material. HyperSizer enables an efficient evaluation of thousands of 

potential designs versus multiple strength and stability-based failure criteria across multiple load cases. 

HyperSizer sizing process uses a global finite element model to obtain element forces, which are 

statistically processed to arrive at panel-level design-to loads. These loads are then used to analyze each 

candidate panel design. A near optimum design is selected as the one with the lowest weight that also 

provides all positive margins of safety. The stiffness of each newly sized panel or beam component is 

taken into account in the subsequent finite element analysis. Iteration of analysis/optimization is 

performed to ensure a converged design. Sizing results for the hat stiffened panel concept and the fiber 

reinforced foam sandwich concept are presented.

Introduction

The design of structural components involves ensuring their capability to withstand strength, 

stiffness, and stability requirements. However, one may design a structure that will not fail, and may not 

be significantly over-conservative, in a manner that it is unnecessarily heavy and therefore inefficient. 

Although, in traditional structural design, the efficiency of the design is recommended, the importance of 

structural component efficiency becomes more apparent for aerospace systems. The effort to optimize the 

structure in terms of its overall performance is mainly driven by two factors: mission objectives and cost. 

One can view the first factor as the lower bound of the project objective as it gives the minimum mission 

requirements, which must be met. The second factor can be viewed as an upper bound on the project 

objective as it gives the upper bound on the overall cost of the project. It may very well happen that the 

mission requirements are not feasible with the established budgetary constraints either due to a low 

budget, unrealistic mission objectives, and/or limited or undeveloped technology. Assuming that cost is 

guiding the project, it is important to develop a system, which achieves the project goals with a 

reasonable monetary constraint. For aerospace structures, like the one considered in this work, one of the 

main objectives is to produce the lightest weight, yet safe, design. Assuming that the technology 

development to achieve this goal doesn’t produce unrealistic negative time and cost effects, structural 
NASA/TM—2010-216942 2

weight optimization provides a solution to both primary project drivers, namely: mission objectives and 

cost.

This work presents structural optimization of the payload shroud of NASA’s next generation heavy 

lift vehicle, Ares V. Ares V has a mission objective of carrying approximately 1.6 times the mass of 

Saturn V’s capability to the low Earth orbit (LEO) and 17 times that of space shuttle’s capability. The 

payload shroud, or payload fairing, will protect the payload, which currently consists of ALTAIR lunar 

lander, from the exterior environment. This exterior environment includes, but is not limited to, 

aerodynamic, acoustic, and thermal loading. Previous studies showed that choosing a correct geometry for 

the outer mold line (OML) allows for minimization of the weight of the total engineering system 

including structural weight, weight of thermal protection, and weight of acoustic blankets. The shape 

chosen to minimize the overall weight of these three aspects is a tangent ogive. This work will elaborate 

on the structural optimization of this shape to obtain the optimum design. Thermal and acoustic 

environments were not considered in the analysis.

In this work structural optimization consisted of a material and a geometry optimization. Due to 

improvements in material characteristics and manufacturing quality over the last two decades, composite 

materials were chosen to provide additional weight savings when compared to metallic construction. The 

material chosen for this study was an IM7/977-3 graphite epoxy composite with pristine autoclave tape 

properties. The study compares two composite panel architectures, which after preliminary analysis, 

showed great potential to provide a low-weight design for the payload shroud, the hat stiffened panel and 

the fiber reinforced foam sandwich panel (FRF). Although, the hat stiffened panel can be made from 

fabric material, especially in the stiffeners, and the FRF concept is an out-of-autoclave concept, the 

material choice was fixed for closer comparison of the two architectures and not their manufacturing 

heritage. Further work will be needed to obtain true material properties of the manufacturable concepts. 

These two concepts were sized using sizing software coupled to finite element solver. Sizing strategies

and basics of the sizing software, HyperSizer, will be discussed in the subsequent sections.

Description of the Hat Stiffened Panel

The hat stiffened panel (Figure 1) is currently used as a fuselage on Boeing 787 and Airbus A350 as 

well as a concept for airplane wing construction. Its design variables include panel height, facesheet 

thickness, web thickness, stiffener spacing, crown width, flange width, and web angle (Figure 2). Due to 

its high efficiency in carrying axial loads, the hat stiffened panel with co-cured stiffeners has been 

considered for the construction of the Ares V payload shroud. A significant portion of the shroud, the 

cylindrical barrel, is subjected to such loading. The hat stiffened panel can also have a high buckling 

resistance if the panel is combined with cylindrical ring frames, as it usually is. Global buckling in the 

cylindrical sections is usually mitigated, provided the ring frames are stiff enough such that buckling 

occurs between the ring frames. This effectively reduces the buckling length of the panel to that of the 

ring spacing. An addition of ring frames may not have a significant impact on the mass of the structure, as 

even light ring frames can significantly increase the circumferential stiffness of the panel. However, the 

hat stiffened panel is not as efficient as some of the sandwich panels under biaxial bending loads. In this 

type of loading, the hat concept may be more susceptible to buckling and addition of ring frames may not 

be adequate to preclude such behavior. In the shroud, the ogive section is subjected to high biaxial

bending loads. Therefore, due to its additional bottom facesheet, a corrugated concept was chosen to 

better treat the effects of bending in the ogive region (Figure 3 to Figure 5). The decision of choosing a

corrugated sandwich was based on the assumption that the transition between the hat stiffened panel and 

the corrugated sandwich can be performed without a circumferential joint due to the similarity of the two 

concepts. In addition, the forward end of the ogive is composed of a thicker solid laminate. This was 

dictated by manufacturing as the stiffeners in the corrugated panel cannot converge to a point (Figure 3). 

The high curvature of the shroud in this region allowed the use of solid laminate to be fairly efficient.
NASA/TM—2010-216942 3

Figure 1.—Hat stiffened panel.

Figure 2.—Hat stiffened panel design variables.

Figure 3.—Shroud concepts.

Figure 4.—Corrugated panel.

Solid laminate

Hat stiffened panel

Corrugated panel

Separation pails
NASA/TM—2010-216942 4

Figure 5.—Corrugated panel design variables.

The corrugated panel (Figure 5) design variables include panel height, top facesheet thickness, bottom 

facesheet thickness, web thickness, stiffener spacing, crown width, and web angle. Due to the many 

similarities in cross section geometry and design variables between the hat stiffened concept and the 

corrugated concept, it is assumed that the two can be joined readily and do not require a circumferential 

joint.

Description of the Fiber Reinforced Foam Sandwich Panel

The primary application of the FRF concept to date has been in wind turbine blades, where the 

concept reduces cost and improves performance compared to balsa and PVC foam designs. Because cost 

is a primary driver in this application, glass fibers are typically used rather than carbon fibers. FRF is also 

currently employed in marine structural applications, such as hulls, decks, bulk-heads, cockpit floors, hard 

tops, and fishing tower platforms. Additional non-aerospace applications include specialty shelters, high 

impact matting, and bridge decks. Aerospace applications for FRF include the ATK shroud boat tail 

demonstration article, an impact-resistant turbine fan case designed and fabricated with NASA Glenn 

Research Center, and a weapons bay door for a Boeing UCAV, all of which were designed and 

manufactured using carbon fibers. In addition, a FRF sandwich panel preliminary design for the NASA 

Crew Exploration Vehicle Crew Module was developed (Bednarcyk et al. 2007).

Fiber Reinforced Foam (FRF) (WebCore Technologies, LLC) is a novel sandwich panel concept that 

combines aspects of foam sandwich panels and stiffened panels (Figure 6). The panel is constructed by 

starting with long strips of closed-cell structural foam, such as Rohacell (Evonik Industries), with 

rectangular cross sections. These strips are wrapped with dry carbon fibers and placed adjacent to each 

other on a tool with dry carbon fiber facesheet preforms on the top and bottom. The panel is then 

infiltrated with resin via vacuum-assisted resin transfer molding (VARTM) (Berg and Higgins 2008).

Once the panel is infiltrated and cured, the fiber overwraps located between the foam strips form integral 

composite webs (Figure 6). The webs provide significantly improved through-thickness shear strength 

compared to foam sandwich panels, while the foam provides support for the webs against local buckling. 

The webs also provide some limited axial strength and stiffness, however, this is not their primary 

purpose as they typically do not contain axial fibers, as the wrapping procedure requires a minimum 

helical angle of approximately 5°. Facesheets provide most of the panel’s bending stiffness as they are 

located farther from the panel neutral axis. The FRF panel design variables include panel height, top 

facesheet thickness, bottom facesheet thickness, web thickness, web spacing, and core type (Figure 7). 

Compared to aluminum honeycomb sandwich panels, FRF sandwich panels can have the advantage of not 

requiring an adhesive between the facesheet and core, which adds parasitic mass, and debonding between 

facesheet and core is not typically an issue with FRF. Furthermore, FRF sandwich panels provide 

improved through-thickness shear strength in the web direction compared to aluminum honeycomb 

sandwich panels and they do not exhibit reduced shear strength as the panel thickness increases. Finally, 
NASA/TM—2010-216942 5

the FRF sandwich concept is highly damage tolerant as its multiple webs, which are supported by the 

foam, provide redundant load paths, and, as stated above, the facesheets do not tend to delaminate from 

the core. The FRF preform can also be stitched prior to resin infiltration to further improve damage 

tolerance. Since FRF incorporates advantages of both sandwich and stiffened panels, the shroud design 

studied was composed entirely of this architecture (Figure 8).

Figure 6.—Fiber reinforced foam panel.

Figure 7.—Fiber reinforced foam panel design variables.

Figure 8.—Fiber reinforced foam panel design variables.

FRF panel

FRF panel

FRF panel

Separation rails
NASA/TM—2010-216942 6

Description of the Payload Shroud

The payload shroud, or fairing, safeguards the payload from the external environment, which can 

include external loading such as aero loads, acoustic vibrations, and thermal effects, among others. Some 

of the effects of these external loads can be mitigated by the shroud's geometry. From previous studies, 

the geometry that provides the greatest overall mitigation to the examined loads was a cylinder/tangent 

ogive configuration (Figure 9).

The OML diameter of the barrel as well as the aft portion of the ogive measures 33 ft (10.0584 m) 

(Figure 10). The height of the barrel is 8 m and the height of the ogive is 14 m (Figure 10). To improve 

buckling characteristics of the barrel section, ring frames spaced at roughly 1 m were added in both 

concepts (Figure 10 to Figure 11). Two ring frames were also added in the ogive section for the FRF 

concept (Figure 11) to decrease buckling susceptibility in this region. The corrugated construction did not 

require this modification.

Both concepts included a ring frame at the interface of the cylinder and ogive sections as well as four 

separation rails oriented in the axial direction. Five penetrations were considered in this shroud 

configuration, one in the ogive, the other four in the barrel. Buildup regions were also added around the 

penetrations. The penetrations and separation rails were not sized in this work.

Figure 9.—Shroud geometry.

Figure 10.—Ring stiffeners for the hat stiffened panel.
NASA/TM—2010-216942 7

Figure 11.—Ring stiffeners for the FRF panel.

(a)

(b)

Figure 12.—Applied pressure load: (a) Max Q loading; and (b) Max 

Q-α loading.

Description of Loading, Safety Factors, and Failure Criteria

The finite element shroud model was subjected to three types of loading conditions: max G loading 

condition, max Q loading condition, and max Q-α loading condition. Four orientations of the max Q-α

loading were chosen for this preliminary sizing for various Mach numbers; however, these may not lead 

to the worst case behavior of the shroud. In future work, more orientations of the max Q-α loading should 

be considered for more accurate mass estimates. Applied pressures contours for max Q and max Q-α are 

shown (Figure 12). Exaggerated deflection shapes due to each loading condition are also shown 

(Figure 13 and Figure 14), although the scaling factors among the figures are not the same. Material 

properties at an elevated temperature of 220 °F were used when considering the max G loading condition.

Boundary conditions restraining all translational and rotational degrees of freedom were applied at the aft

(barrel) end. The mechanical limit safety factor was chosen to be 1 and the ultimate limit safety factor 

was chosen to be 1.4. While there was no local buckling knockdown, the global buckling knockdown was 

set at 0.65.
NASA/TM—2010-216942 8

 Figure 13.—Hat panel deformation: Max G loading, Max Q loading, Max Q-α loading. Note that the scaling factors are 

not the same.

 Figure 14.—FRF panel deformation: Max G loading, Max Q loading, Max Q-α loading. Note that the scaling factors 

are not the same.

Various failure criteria were used to determine the failure mechanism of the panels. Although other 

failure criteria are available, HyperSizer 5.8.11 standard failure criteria were used at this stage. It was 

determined that these failure criteria are sufficient at this sizing stage. These include checks for the:

• Composite strength, max strain 11, 22, and 12 directions

• Composite strength, max stress 11, 22, and 12 directions

• Composite strength, Tsai-Hill

• Composite strength, Tsai-Wu

• Composite strength, Tsai-Hahn

• Composite strength, Hoffman

• Composite strength, LaRC03 Matrix Cracking

• Composite strength, LaRC03 Fiber Failure

• Composite strength, crippling MIL-HDBK-17-3E including Dij terms

• Global buckling, curved panel

• Local buckling, longitudinal, transverse, and shear directions

Description of the Analysis Process

The goal of the analysis was to obtain preliminary mass estimates for the Ares V payload shroud. 

These estimates were obtained from a sized acreage panel design of the hat stiffened/corrugated panel and 

the fiber reinforced foam sandwich panel. This design aims to accommodate most of the structural loads 

with the assumption that design details and stress concentrations will not alter the final configuration 

significantly. The analysis process was composed of two parts, which were performed iteratively, finite 

element analysis and structural sizing. For the first finite element analysis, solid laminate properties were 

used to obtain preliminary elemental force and moment distribution. The analysis was performed using 
NASA/TM—2010-216942 9

NX NASTRAN for the hat stiffened/corrugated concept and using MSC NASTRAN for the FRF concept. 

Only shell elements for composite panels and bar elements for beams were used in the model. These 

element types were used since their element forces can be imported into the structural sizing software. 

Also, the cross section of both elements is only accounted for in the stiffness matrix formulation and does 

not alter the mesh, making these elements convenient to use for iterative sizing.

Sizing of the structure was performed using sizing software HyperSizer 5.8.11 (Collier 1994, Collier 

1996, Collier et al. 1997), which uses a brute force optimization to find the lightest design that also has 

positive margins of safety. HyperSizer 5.8.11 obtains element forces and moments from a finite element 

analysis. The FE computed generalized forces are then subjected to statistical analysis which determines 

the mean of the loads and their standard deviation. For this sizing, a mean load plus two standard 

deviations are applied to the panel concept. These statistical loads are then used in strength of materials 

type calculations using composite material properties to distribute them to the various parts of the panel 

whose limits and range on the geometry are defined. Hence, free body diagram calculations are performed 

on numerous panel configurations. Buckling and strength criteria are checked for panels ordered from 

lightest to heaviest such that the lightest panel with positive margins can be chosen without performing 

calculations for the heavier panels. If no panels within the limits have positive margins, the user needs to 

augment the limits or range of variables. Assuming the optimal panel is found, A, B, and D matrices are 

computed for a unit thickness shell element such that the correct stiffness properties are captured in that 

element. The procedure is analogous to the method of transformed sections commonly used in composite 

beam calculations. These properties are imported into a finite element code to update the element stiffness 

such that new updated force and moment distribution can be computed. The updated element forces and 

moments are imported into HyperSizer 5.8.11 where a new optimization is performed. The process of 

optimization and finite element analysis continues until the mass of the structure stabilizes. At this point 

the geometry of the design should also stabilize, although this is not a convergence criterion and should 

be checked by the analyst. The HyperSizer optimization process is depicted below (Figure 15).

The payload shroud was divided into several components. Each component was sized individually in 

HyperSizer 5.8.11 and thus had its own cross section geometry. The component subdivision for the hat 

stiffened/corrugated and FRF panels are shown below (Figure 16 and Figure 17).

Figure 15.—HyperSizer sizing process.
NASA/TM—2010-216942 10

Figure 16.—Hat stiffened/corrugated sizing components.

Figure 17.—FRF sizing components.

For the hat stiffened/corrugated concept, seven different major sizing components were considered

(Figure 16), four of them in the ogive and three in the barrel. The ogive nose was sized using a solid 

laminate while the rest of the ogive used a corrugated panel concept. The barrel was sized using a hat 

stiffened panel concept. Separate solid laminate components around the penetrations were also 

considered. The FRF concept had additional components in the barrel (Figure 17). Each of the hat 

stiffened components was further subdivided into four components circumferentially. This allowed for the 

better design-to load assessment, which benefitted the FRF design, since statistical load analysis is 

performed for an individual component. The four circumferential sections were then linked to ensure that 

they are composed of the same cross section. Due to different manufacturing constraints, the FRF concept 

penetration buildups and ogive nose were sized as FRF panels. For both concepts, ring frames were sized 

such that they precluded global buckling. The barrel ring frames were considered as being one 

component, and therefore had the same geometry, while the remaining ring frames were all sized 

independently. Separation rails and penetrations were not sized. The material used in this sizing was a 

graphite epoxy composite IM7/977-3 with the following properties at 75 °F and 220 °F. At 75 °F the 

properties are:

• ETension11 = 22.18 Msi [152.93 GPa]

• ETension22 = 1.29 Msi [8.89 GPa]

• ECompression11 = 21.73 Msi [149.83 GPa]

• ECompression22 = 1.29 Msi [8.89 GPa]
NASA/TM—2010-216942 11

• Poisson’s RatioTension12 = 0.329 GPa]

• Poisson’s RatioCompression12 = 0.329

• Shear modulus12 = 0.71 Msi [4.90 GPa]

• Allowable tensile stress11 = 275.06 ksi [1896.54 GPa]

• Allowable tensile stress22 = 15.765 ksi [108.70 GPa]

• Allowable compressive stress11 = 227.285 ksi [1567.13 GPa]

• Allowable compressive stress22 = 19.443 ksi [134.06 GPa]

• Allowable shear stress12 = 9.149 ksi [63.08 GPa]

At 220 °F the properties are:

• ETension11 = 22.18 Msi [152.93 GPa]

• ETension22 = 1.12 Msi [7.72 GPa]

• ECompression11 = 21.73 Msi [149.83 GPa]

• ECompression22 = 1.12 Msi [7.72 GPa]

• Poisson’s RatioTension12 = 0.383

• Poisson’s RatioCompression12 = 0.383

• Shear modulus12 = 0.54 Msi [3.72 GPa]

• Allowable tensile stress11 = 275.454 ksi [1899.26 GPa]

• Allowable tensile stress22 = 12.25 ksi [84.46 GPa]

• Allowable compressive stress11 = 152.966 ksi [1054.70 GPa]

• Allowable compressive stress 22 = 9.951 ksi [68.61 GPa]

• Allowable shear stress12 = 6.837 ksi [47.14 GPa]

Both concepts were sized using equivalent orthotropic materials with different ply percentages (Table 1).

TABLE 1.—EQUIVALENT LAMINATES USED IN SIZING

Laminate Percent of 45 

plies

Percent of 0 

plies

Percent of 90 

plies

1 20 20 60

2 30 30 40

3 20 70 10

4 40 50 10

5 50 25 25

Analysis Results

The majority of both structures experienced failure by either local buckling or global buckling 

(Figure 18 and Figure 19). Buckling analysis for both concepts was performed using finite element 

analysis (Figure 20 and Figure 21). The total mass of the hat stiffened/corrugated shroud was 2495 kg and

the total mass of the FRF shroud was 2163 kg. The general sizing summary is shown (Table 2 to Table 5).

Note that the mass of the separation rails and the mass of the penetration panels, which were not sized, is

excluded. It is noted that the hat stiffened/corrugated shroud ring frames totaled 436 kg whereas the FRF 

ring frames totaled 73 kg. This 363 kg ring frame mass savings accounts for the entirety of the mass 

savings of the FRF design, which was 332 kg lighter than the hat stiffened/corrugated design. It is 

expected that the hat stiffened barrel should require heavier ring frames compared to the FRF barrel as the 

hat stiffeners provide no circumferential stiffness or strength.
NASA/TM—2010-216942 12

Figure 18.—Hat stiffened/corrugated failure modes.

Figure 19.—FRF failure modes.

Figure 20.—Hat stiffened/corrugated first buckling mode (Eigenvalue = 2.43).
NASA/TM—2010-216942 13

Figure 21.—FRF first buckling mode (Eigenvalue = 2.23).

TABLE 2.—HAT/CORRUGATED PANEL SIZING RESULTS SUMMARY

Component Area

(m2)

Unit mass

(kg/m2)

Mass

(kg)

Lowest

MS

Controlling

failure mode

Tip 1.99 1.28 2.54 2.539 Composite strength, Tsai-Wu

Ogive fwd 46.48 3.91 181.66 0.02121 Local buckling

Ogive mid 100.06 3.93 393.13 0.08008 Curved panel buckling

Ogive aft 185.99 3.73 694.45 0.0417 Local buckling

Barrel fwd 81.54 2.46 200.71 0.007769 Local buckling

Barrel mid 84.27 2.96 249.52 0.0316 Curved panel buckling

Barrel aft 81.37 3.58 291.48 0.009815 Curved panel buckling

Ogive buildup 0.46 7.70 3.54 0.04347 Curved panel buckling

Barrel fwd buildup 1.19 13.22 15.73 0.001398 Curved panel buckling

Barrel aft big buildup 0.85 13.64 11.59 0.02598 Curved panel buckling

Barrel aft small buildup 1.02 14.49 14.78 0.02739 Curved panel buckling

TABLE 3.—HAT/CORRUGATED BEAM SIZING RESULTS SUMMARY

Component Length

(m)

Unit Mass

(kg/m)

Mass

(kg)

Lowest

MS

Controlling

failure mode

Barrel ring frames 252.77 1.68 425.06 0.004364 Stiffness requirement

Barrel/ogive frame 31.61 0.35 10.93 0.01343 Stiffness requirement

TABLE 4.—FRF PANEL SIZING RESULTS SUMMARY

Component Area

(m2)

Unit mass

(kg/m2)

Mass

(kg)

Lowest

MS

Controlling

failure mode

Tip 1.99 3.114 6.204 0.2339 Curved panel buckling

Ogive fwd 46.48 3.709 172.4 0.1114 Curved panel buckling

Ogive mid 100.06 3.907 390.8 0.11 Curved panel buckling

Ogive aft 185.99 3.709 689.9 0.1291 Curved panel buckling

Barrel fwd 81.54 2.915 237.7 0.1402 Curved panel buckling

Barrel mid 84.27 3.114 262.4 0.1202 Curved panel buckling

Barrel aft 81.37 3.907 318.0 0.08725 Curved panel buckling

Ogive buildup 0.46 3.709 1.718 0.2055 Curved panel buckling

Barrel fwd buildup 1.19 3.114 3.719 0.1666 Curved panel buckling

Barrel aft big buildup 0.85 3.907 3.332 0.2235 Facesheet wrinkling

Barrel aft small buildup 1.02 3.907 3.999 0.1795 Curved panel buckling
NASA/TM—2010-216942 14

TABLE 5.—FRF BEAM SIZING RESULTS SUMMARY

Component Length

(m)

Unit mass

(kg/m)

Mass

(kg)

Lowest

MS

Controlling

failure mode

Forward ogive frame 17.49 0.8388 14.67 0.01972 Stiffness requirement

Mid ogive ring frame 26.74 0.7456 19.94 0.00742 Local buckling

Aft ogive ring frame 31.60 0.08563 2.706 0.4286 Composite strength

Barrel ring frames 252.77 0.1304 32.97 1.213 Stiffness requirement

Barrel aft ring frame 31.61 0.08563 2.706 10000 Composite strength

Summary

This paper presents a preliminary design/sizing summary for the Ares V payload shroud. An 8 m 

high cylindrical barrel with a 33 ft (10.0584 m) diameter and a 14 m high tangent ogive with a 33 ft 

(10.0584 m) aft diameter was sized using a hat stiffened/corrugated concept and a fiber reinforce foam 

concept. Sizing using HyperSizer 5.8.11 and linear static analysis using NX NASTRAN and MSC 

NASTRAN were performed to obtain an optimum shroud design. To preclude buckling in the barrel, 

circumferential ring frames were added. The total mass of the hat stiffened/corrugated shroud is 2495 kg

while the mass of the fiber reinforced foam shroud is 2163 kg, 13 percent lighter than the hat/corrugated 

design. The hat stiffened construction is very beneficial in the barrel portion of the shroud where the 

uniaxial loads dominate; however, the FRF construction is more efficient than the relatively heavy 

corrugated architecture in the ogive section where high biaxial loads are present. The mass of the two 

concepts is competitive; however, detailed analysis is needed to verify the preliminary mass estimates.

References

Bednarcyk, B.A., Arnold, S.M., Collier, C.S., and Yarrington, P.W. (2007) “Preliminary Structural Sizing 

and Alternative Material Trade Study of CEV Crew Module,” NASA/TM—2007-214947.

Berg, J.S. and Higgins, J. (2008) “VARTM Infusion: Processing Large Carbon/Epoxy Space Structures 

Out of the Autoclave,” SAMPE Journal, Vol. 44, No. 6, pp. 41–44.

Collier, C.S., “Thermoelastic Formulation of Stiffened, Unsymmetric Composite Panels for Finite 

Element Analysis of High Speed Aircraft,” 35th AIAA/ASME/ASCE/AHS/ACS Structures, 

Dynamics, & Materials Conference, Hilton Head, SC, Apr. 1994, AIAA–94–1579.

Collier, C.S., “An Integrated FEA and Design Optimization System for Composite Structures,” The 1996 

MSC World Users’ Conference, Newport Beach, CA, Jun. 1996.

Collier, C.S., Pickenheim, M., and Yarrington, P.W., “Next Generation Structural Optimization Today,”

the 1997 MSC Aerospace Users Conference, Newport Beach, CA, Nov. 1997.

Evonik Industries, Darmstadt, Germany, http://www.rohacell.com/en/performanceplastics8344.html.

WebCore Technologies, LLC, Miamisburg, OH, http://www.webcoreonline.com.
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4. TITLE AND SUBTITLE

ACT Payload Shroud Structural Concept Analysis and Optimization 

5a. CONTRACT NUMBER

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6. AUTHOR(S)

Zalewski, Bart, F.; Bednarcyk, Brett, A. 

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WBS 727950.04.03.22 

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National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field 

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Available electronically at http://gltrs.grc.nasa.gov 

This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

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14. ABSTRACT

Aerospace structural applications demand a weight efficient design to perform in a cost effective manner. This is particularly true for launch 

vehicle structures, where weight is the dominant design driver. The design process typically requires many iterations to ensure that a 

satisfactory minimum weight has been obtained. Although metallic structures can be weight efficient, composite structures can provide 

additional weight savings due to their lower density and additional design flexibility. This work presents structural analysis and weight 

optimization of a composite payload shroud for NASA’s Ares V heavy lift vehicle. Two concepts, which were previously determined to be 

efficient for such a structure are evaluated: a hat stiffened/corrugated panel and a fiber reinforced foam sandwich panel. A composite 

structural optimization code, HyperSizer, is used to optimize the panel geometry, composite material ply orientations, and sandwich core 

material. HyperSizer enables an efficient evaluation of thousands of potential designs versus multiple strength and stability-based failure 

criteria across multiple load cases. HyperSizer sizing process uses a global finite element model to obtain element forces, which are 

statistically processed to arrive at panel-level design-to loads. These loads are then used to analyze each candidate panel design. A near 

optimum design is selected as the one with the lowest weight that also provides all positive margins of safety. The stiffness of each newly 

sized panel or beam component is taken into account in the subsequent finite element analysis. Iteration of analysis/optimization is 

performed to ensure a converged design. Sizing results for the hat stiffened panel concept and the fiber reinforced foam sandwich concept 

are presented. 

15. SUBJECT TERMS

Structural design; Loads; Fiber composites; Foams; Finite element method; Launch vehicles; Composite structures; Structural 

analysis; Shrouds; Ares V cargo launch vehicle; Stability; Failure; Panels 

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