Holding It All Together: Space Structures

Eric Hall

Space structure design is characterized by an almost fanatical concern with reducing weight, accommodating diverse loads, and facilitating integration with other subsystems.

Beneath the shiny exterior of any space vehicle is the subsystem that holds it all together: the structure. Most easily envisioned as the skeleton, the structural subsystem forms the framework for the space vehicle and provides attachment points for all other subsystems.

The design goal for a spacecraft structure is simple in theory: provide enough material to carry expected loads with enough stiffness to maintain the relative position of attached components. In practice, the process is complicated by the need to minimize weight, accommodate diverse and unique loads and environments, and achieve a high degree of integration with other subsystems.

The development of a space structure is a long and complex process that involves many design iterations. There is a large penalty for mistakes, so a process to periodically review and assess the maturity and completeness of the design from an independent perspective is essential. Aerospace serves this function on many space programs. Drawing upon longstanding expertise in conceptual design, design criteria, structural analysis, material characterization, and large-scale structural testing, Aerospace has developed and tailored the theoretical methods and analytical tools of structural mechanics to tackle the challenges of structural design over the life cycle of a space vehicle.

The classic form of a space vehicle structure includes a stiff central truss or cylinder (shown in magenta) that has a direct connection to the launch vehicle adapter and lightweight panels (blue), which carry and enclose other subsystems. Solar arrays (green) and large payloads are mounted on the exterior.

Basic Concerns

Space-system structures are ultimately assessed in terms of strength, stiffness, and stability. The science in space structure design is in understanding the expected environment, identifying all potential failure modes, and quantifying how much capability exists in a structure relative to its expected environment.

One of the core technical missions of Aerospace is to characterize the loads and environments that affect a space vehicle during its life. Space structures can be extremely delicate, and even simple handling, transportation, and deployment testing can cause severe damage. The launch environment is particularly stressful, exposing structures to a variety of forces, including inertial loads from atmospheric flight and dynamic excitation, acoustic pressure, staging and separation shocks, and depressurization as the vehicle leaves the atmosphere. On-orbit loads can also present unique challenges; although the basic mechanical loads are minimal, large thermal loads can cause static deformation, and radiation can cause material degradation. So, the first step in the structural design process is to estimate the maximum expected loads and environments.

Next, the theoretical methods of structural mechanics are applied to determine whether a particular configuration will survive without damage. This assessment is typically rendered in terms of a load margin, which represents the increase in load that a structure can tolerate before reaching a limiting condition such as permanent deformation or rupture. Statistical limits are typically determined through various test methods and are represented in terms of an "allowable"—the load at which a given structure can survive with a given certainty. Aerospace supports this process through extensive work in characterizing known failure modes and identifying new failure modes in advanced materials.

One complication in this process is that the structure is a dependent subsystem, and its requirements are typically traceable not to mission requirements but to a flow-down of requirements from other subsystems. For example, the forces that a structure must carry depend on the mass of all of the subsystems that it supports, including the structure itself. On top of this, the loads that a structure experiences during launch are dependent upon the design of that structure. A further complication is that the dividing line between structures and other subsystems can be ambiguous in a highly integrated system such as a space vehicle. Thus, the design process is highly iterative and requires a conservative approach to prevent drastic fluctuations as the design matures. Aerospace has documented best practices for managing such conservatism in guidelines such as TOR-2003(8583)-2894, subsequently published as AIAA Standard S-110-2005, Space Systems—Structures, Structural Components, and Structural Assemblies.

Unique Considerations

The inherent difficulty and complexity of spaceflight presents unique challenges for the structural designer. In particular, the designer must address stringent volumetric and mass constraints, the need for autonomous operation in a harsh environment, low production rate, cost constraints, and risk avoidance.

The size and mass of a space vehicle is naturally constrained by the launch vehicle. Every pound is critical, so designers seek to reduce mass to the absolute minimum. The desire to minimize mass runs contrary to the desire to maximize structural integrity. Reconciling these two conflicting goals typically results in designs with low margins for strength. As a result, even prelaunch handling loads can be significant. Similarly, volumetric limitations constrain the geometric layout of structures and can lead to undesirable load paths.

A space vehicle must perform without maintenance for a long period in a harsh environment. On orbit, vacuum causes the release of trapped gases from organic materials and induces stress in components that have unvented volumes. Radiation can cause material degradation, and extreme temperature fluctuations create large internal stresses and geometric distortions. These considerations limit the materials available to the space structure designer and can influence the basic configuration.

Many space vehicles are unique, and even large constellations are built at a rate of only a few vehicles per year. This means that, in essence, each structural design is unique, and must be tested rigorously to demonstrate adequacy for each new mission. Requirements can change during a long development period, and technology can advance between production of satellites in the same series. For this reason, even satellites in the same constellation are generally not identical, and old designs cannot simply be reused for new missions.

Pressure to control cost can limit the selection of materials, launch vehicles, and methods for verifying structural integrity. Cost trades are an essential part of the design process, and there is always pressure to reduce manpower, production time, and complexity. However, a well-learned lesson is that cost savings from shortcuts in the verification program are easily destroyed by the cost of an anomaly investigation.

Space mission planners are loathe to add risk to an already risky endeavor, especially for important national security space missions. A conservative approach is therefore favored. Flight-qualified technology is preferred, and tangible evidence is required before new technology is considered. All flight systems undergo extensive testing, and small modifications (such as changing a material supplier) can trigger the need for requalification. For some programs, the design cannot be proven until the vehicle is on orbit and operational. Thus, there is high demand for analytical simulations and tests of space structures to manage and mitigate potential risks.

The Development Process

The designer of a space vehicle structure usually begins with just a few parameters, perhaps only supported mass, design load factors, volume, and overall system mass. From this starting point, the designer follows an iterative process of design, analysis, and assessment. The level of fidelity increases with each new iteration until finally, flight-qualified hardware is available for launch. Throughout this process, Aerospace applies various analytical tools to verify the adequacy of structural designs. Some of these specialized tools do not exist elsewhere in the industry—for example, tools for progressive damage analysis of high-temperature composite applications and fatigue assessments of electronic components.

Conceptual Design

The first step in structural design is to develop a concept that can be used for mission and payload trade studies. The goal is to assess the structural feasibility in terms of size, weight, power, cost, and technology. A general configuration is proposed based on defined requirements including size, weight, volume, fields of view, payload geometry, survivability, and service life. Typically, a similar spacecraft is used as a template to produce a weight breakdown and component list that will ultimately be used to derive subtier requirements for specific structural elements. Structural members are sized using similarity and empirical or analytical relations that include some margin for growth and allow sufficient flexibility to accommodate alternate designs. Typically, what is needed is a mass estimate for the primary structure and deployables, a layout of spacecraft mass distribution and inertia, and a size budget for the solar, radiator, and payload panels.

At this point, key decisions focus on the development approach (old or new design?), the general shape of the primary structure (thrust cylinder or truss?), the class of materials required (metal or composite?), and configuration changes (fixed or deployable components?). These are determined by means of historical data, rules of thumb, hand calculations, and spring-mass models. Various configurations are typically simulated to establish weight budgets in support of trade studies. The analytical simulations also enable the designer to gain insight into the major design issues and to identify the greatest risks.

Theoretical methods and computer models enable the simulation of structural response. Here, Matthew Keough investigates a sensor platform deformed by dynamic loads. These simulations are also used to identify the areas of a structure most susceptible to failure.

Requirements Specification

Once a baseline configuration has been established, it is possible to continue the requirements allocation process through the flow-down of subsystem requirements. In the structures area, iterative allocations of key parameters such as loads, operating temperatures, and component stiffness are evaluated. As allocations mature, a significant number of studies are performed to assess high-level trades (such as cost versus performance) and low-level trades (such as minimum frequency versus structural weight). The structural designer then translates the resulting requirements into design goals or directives—for example, to transfer the accelerations necessary for orbital insertion to all space vehicle components, to maintain the alignment of all components, to provide environmental protection for equipment, and to provide a foundation for deployables. As the design matures, these design goals become requirements for individual structural elements.

The designer begins to specify the component forms, such as honeycomb panels, truss tubes, solid shells, solid panels, tubes, brackets, and flexures. Analytical simulations are valuable at this stage not only to assess the design but to ensure that the allocated requirements are reasonable and achievable. An assessment is typically made regarding the satisfaction of the major structural requirements—mass, geometric constraints, stowed volume, dynamic envelope in the launch vehicle fairing, component and system stiffness, on-orbit stability, and stress margins.

Preliminary Design

Once a baseline vehicle configuration has been established, the structural designer will assess structural load paths, refine specifications related to subsystem geometry and interfaces, and further define the dimensions of structural elements. The greatest emphasis is on the "quasi-static" loads related to inertial loading and steady-state aerodynamic forces encountered during launch and ascent. These loads typically dictate the size of structural members, the type of failure modes, and the allowable mass distribution and center-of-gravity location for the vehicle in the stowed configuration. Other priorities would include dynamic considerations, such as the fundamental frequencies for the entire vehicle and selected structural elements such as equipment panels. The on-orbit thermal environment can also be a significant factor in material selection, joining techniques, and subsystem interactions. Secondary design drivers may include manufacturability, ground handling, damage tolerance, testability, the radiation environment, micrometeoroids, on-orbit geometric stability, material degradation, fracture control, gravity deformation, and contamination.

In developing a preliminary design, the structural designer must assess the feasibility of mass allocations, subsystem interfaces, and the potential manufacturing approach. The objective is to develop element designs by specifying the materials, form, and dimensions for major structural elements. It is also important to assess the structural performance relative to allowables, dynamic response, and fundamental frequencies. At this stage, it is also important to consider the verification approach, design sensitivities, failure modes, and long-lead procurements. Other intangible criteria are also assessed, including analyzability, accessibility, inspectability, testability, and reliability. An important step in this process is to clearly define the responsibility for mechanical design of interfaces.

Finite-element models, such as the one shown here, are used extensively in the design phase. They help the structural designer specify the materials, form, and dimensions for major structural elements.

Detailed Design

As the design coalesces, the focus will shift from design to analysis. Designs are refined, and comprehensive finite-element models are developed for both stress analyses and coupled-loads analyses. At this point, the focus is on detailed simulation of the structural response to loads. The goal is to identify failure modes, failure loads, and elastic responses. As the design matures, more detailed analyses may refine the launch and ascent loads. Similarly, multidisciplinary analysis may be used to explore and iterate allocations with other subsystems.

By the end of the detailed design phase, the structure will be specified down to the component level, including the specific nuts and bolts, preloads, material specifications, and manufacturing and assembly tolerances. All of these will be captured via design drawings, which can now be released to manufacturing. All structural margins have been fully assessed, and all assembly, integration, and test details will be fairly mature, including test fixture designs, off-load fixture designs, alignment procedures, and allocated facilities.

Integration, Test, and Verification

Demonstrating structural integrity is critical, but it is impractical to fully simulate a launch during ground testing. Thus, the standard practice is to separate the test and evaluation process into discrete tests that apply the most stressful conditions to various levels of assembly. The launch and ascent loads are most relevant to the primary structure and are traditionally simulated by applying static loads to the structure that envelop predicted inertial loads. Vibroacoustics are most important for large, lightweight panels and compact electronics, and these are tested by exposure to direct (acoustic) or transmitted (base-shake) vibration. For some structures, such as propellant tanks, pressure loads can be critical, and internal pressurization is usually achieved using water and ullage pressure. Structures that are sensitive to thermally induced loads can be tested using heating blankets or cooling systems. What's important is to ensure that all of the critical structural failure modes are adequately tested.

The goal of the test program for launch and ascent loads is to accurately, yet conservatively, simulate the expected structural loads. The challenge is that the loads themselves are dependent on the design of the space and launch vehicles, and therefore the process is by necessity iterative. The test and evaluation process must reflect this iterative nature and include a combination of analyses and tests. Structural tests have three objectives: to obtain data for model correlation, to qualify a structural design, and to verify the adequacy of flight hardware. As for the second objective, it has become fairly common industry practice to consider structural qualification by analysis only, with higher design factors of safety, as an equal alternative to full-scale qualification testing. Aerospace does not support this approach for primary structures; not only does it increase weight, it also increases risk (in spite of the higher factors of safety) through the introduction of analytical unknowns. It also precludes the opportunity to learn from the test. Even successful tests can identify potential design and manufacturing flaws and lead to significant product improvements. There is no technically equivalent substitute for testing in the flight configuration.

Safety and structural reliability considerations mandate proof or acceptance testing of many structures. Acceptance and proof tests are similar to qualification tests, except that they are performed on flight hardware to verify workmanship or capability. Pressure vessels and pressurized structures commonly require proof testing, and other structures that should be proof tested whenever possible include fracture-critical fittings or bolts, bonded structures, and composite structures. Proof testing of these structures adds a measure of protection against defects and damage introduced through manufacturing and assembly. Acceptance tests may require only nominal loading to screen out poor workmanship or preclude a hazardous condition. Proof tests always require that the flight structure be subjected to at least the maximum expected load.

Summary

Structural failures in flight are quite rare in large part because of the tradition of well-designed structural tests in conjunction with rigorous loads and strength analysis and vigilant quality control; however, failures in structural test are still fairly common, as are costly and time-consuming retests necessitated by postqualification deviations. Following sound engineering principles can help minimize these problems (see sidebar, Sound Engineering).

New technology continues to improve space structures, but the use of new technology is no substitute for a rigorous development process. For example, advanced composites have replaced metals as the structural materials of choice because they offer numerous advantages in terms of weight, customization, and reduced part counts; however, their use often adds complexity and uncertainty to the development and qualification processes. Although advanced finite-element analysis has exposed a number of behaviors and failure modes not previously considered, mistakes are made every day in the application of complex computer tools. The structural designer must always fall back on good engineering judgment to build a solid skeleton for space vehicles.