Astronomy Station Vibration Level

10 100 1,000 Hz

Fig. 11-26. Random Vibration. The plot shown on the left is an example of how acceleration from random vibration would vary overtime. The probability density curve (center), with the same vertical scale, describes the relative likelihood of acceleration being at a given value. This is a normal distribution, with each tick mark on the vertical scale representing a standard deviation, a. The figure on the right is a plot of power spectral density. It describes the frequency content of the vibration and Is equal to the mean-square acceleration (g2) in a selected frequency band divided by the width, in Hz, of that band.

In actual design, we combine math models of the spacecraft and launch vehicle to do a coupled loads analysis. In this analysis, we drive the coupled model with forcing functions (force as a function of time or frequency) that are based on measured launch-vehicle environments. Before we get to this point, though, we must configure the structure, select from our design options, and roughly size the structure based on estimated design loads.

11.6.2 Packaging and Configuring the Subsystem

Designers must trade the low weight of a high-density design against the need to access individual componente for testing or replacement before launch. The preliminary arrangement should account for every component in the design because the spacecraft's structure inevitably becomes heavier if it must accommodate new components. Added component mass multiplies through higher allowances for weight growth, heavier structure for support, and more onboard propellants for attitude control.

The payload and the attitude control approach most strongly influence a spacecraft's configuration, and the launch vehicle constrains it Chapter 9 discusses payloads and their requirements. Section 11.1 discusses spin- and 3-axis stabilization to control attitude. A spin-stabilized spacecraft influences packaging most because the mass moment of inertia (MOT)* about the spin axis must be greater than any other axis to maintain stability. In 3-axis stabilized spacecraft, we must separate the magnetic torque rods and the magnetometer (device that senses the Earth's magnetic field) to preclude any magnetic interference between them.

Sensing devices always require specific fields of view and pointing accuracy. The packaging designer must locate sensors to be unobstructed by antennas or solar arrays.

* / is the standard notation for both mass moment of inertia and area moment of inertia. In this section, we use MOI for mass moment of inertia and I for area moment of inertia.

Advanced composite materials often help mounting-structure designs meet requirements for rigidity and thermoelastic distortion.

Communication antennas also require rigidity, thermoelastic stability, and a clear field of view. One solution is to mount sensors and antennas on an appendage that is stowed during launch, deployed on orbit for an unobstructed view, and constructed of advanced composite materials for rigidity and thermoelastic stability.

Components for command and data handling are often vulnerable to the environments of outer space, so we usually bury them in the center of the spacecraft to shield against radiation. Interfacing wire bundles also weigh less when the processor, data bus, and other control components are close together.

Propulsion subsystems include reaction-control assemblies and orbital transfer stages. By purchasing thrusters in multiaxis combinations, called rocket engine modules, we can reduce the number of propellant-line welds needed on site during the spacecraft's assembly. We usually place those modules on the spacecraft's periphery—far from the spacecraft's center of mass—but we must keep them from contaminating sensors, antennas, and solar array cells with propellant gases. A propulsion system with a low operating pressure helps lessen the propellant tank's weight Another structural challenge is die need to support transfer stages, so the thrust vector remains aligned with the spacecraft's center of mass. These stages are usually heavy; so, placing them on the bottom of the spacecraft stack, near the launch vehicle interface, helps minimize structural weight

The configuration of the power subsystem varies with power requirements and orbital conditions. For example, we must determine where to stow solar panels during launch and where to deploy them in orbit, so they do not touch or rest in the shadow of other subsystems. By using fewer folds in the panels, we can keep the deployment mechanisms simple and more reliable. Finally, batteries should be accessible for pre-launch testing or replacement and placed where they will be at their optimum temperature. Thermal control specialists can place components, select materials, and suggest surrounding structure (open truss or closed skin panels with stiffeners) to help control temperature. These measures help us avoid using other active temperature control devices.

If we configure the structure and package components at the same time, we may make the components part of the load-carrying structure. This concurrent approach may also produce better symmetry in the structure, which satisfies frequency response requirements. By using common members and joints throughout the design, we can lower fabrication costs and more easily meet weight allocations. For example, beams that make the spacecraft rigid can also support components. Establishing design routes for wire bundles and propellant lines helps avoid the inefficiencies of cutting through structure later. We should design special joints to connect members made with different materials because their varying rates of thermal expansion and contraction can be detrimental, finally, the spacecraft adapter must transition smoothly from the spacecraft to the interface on the launch vehicle's upper stage.

11.63 Design Options

In designing a structure, we consider optional materials, types of structure, and methods of construction. To select from these options, we do trade studies to compare weight, cost, and risk.

A typical spacecraft structure contains metallic and nonmetallic materials. Most metals are very nearly homogeneous, having constant properties throughout their composition, and isotropic, having the same properties regardless of direction. Non-metals are usually formed with composites, or blends of more than one material. Composite materials are not homogeneous and are normally not isotropic. Materials are selected based on:

By far the most commonly used metal for spacecraft structure is aluminum alloy, of which there are many types and tempers. Aluminum is relatively lightweight, strong, readily available, easy to machine, and low in raw material cost The stiffness-to-weight ratio of aluminum is about the same as steel, but the strength-to-weight ratio is usually higher. The main advantage of aluminum over steel for flight structures is its lower density. For the same mass, an aluminum shell or plate would be thicker and thus able to carry a greater compressive load before it would buckle. If we need harder or denser materials, we normally choose steel or titanium.

Alloys are available in sheets, plates, extrusions, forgings, and castings. The primary source of material properties is MIL-HDBK-5, Metallic Materials and Elements for Aerospace Vehicle Structures [U.S. Air Force Materials Laboratory, 1994], which contains many properties and statistically guaranteed strengths for all commonly used aerospace metals.

One popular advanced composite is graphite-epoxy, which has graphite fibers for strength and stiffness in an epoxy matrix. Composite fabric layers normally bond together in designed fiber orientations, so they can provide properties not available in homogeneous metallic materials, including extremely high stiffhess-to-weight ratios and negligible expansion and contraction resulting from temperature gradients. Other fibers in these composites include boron, Kevlar™, and glass. Graphite and boron fibers also reinforce metal-matrix composites. Techniques to manufacture and apply metal-matrix composites are presently less advanced than for epoxy-matrix materials. Tsai [1987] and MIL-HDBK-17 [1977,1989] provide more information on composite materials.

Table 11-51 summarizes the advantages and disadvantages of the most commonly used materials in spacecraft design. Table 11-52 shows their representative properties.

Types of structures include skin panel assemblies, trusses, ring frames, pressure vessels, fittings, brackets, and equipment boxes. Sometimes only one meets objectives; but we usually have several options. We use monocoque structures, which are panels and shells without attached stiffening members, only if applied and reacted loads are spread out rather than concentrated. A semimonocoque shell has lightweight closely spaced stiffening members (stijfeners) that increase its buckling strength. Skin-stringer structures have longitudinal members (stringers) and lateral members (frames) to accept concentrated loads and skin to spread those loads out and to transfer shear. A truss is an assembly that remains stable under applied concentrated loads with its structural members loaded only axially. A sandwich structure is a panel or shell

• Thermal conductivity

• Thermal expansion

• Corrosion resistance

• Ductility (which can prevent cracks)

• Fracture toughness (ability to resist crack growth)

• Ease of fabrication

• Versatility of attachment options.

such as welding • Availability

TABLE 11 -51. Advantages and Disadvantages of Commonly Used Materials.





• High strength vs. weight

• Ductile; tolerant of concentrated stresses

• Low density, efficient in compression

• Relatively low strength vs. volume

• High coefficient of thermal expansion

• Wide range of strength, hardness, and ductility obtained by treatment

• Not efficient for stability (high density)

• Most are hard to machine

• Magnetic


• High strength vs. volume

• Strength retained at high temperatures

• Ductile

• Not efficient for stability (high density)

• Not as hard as some steels


• Low density—'very efficient for stability

• Susceptible to corrosion

• Low strength vs. volume


• High strength vs. weight

• Low coefficient of thermal expansion

• Poor fracture toughness if solution treated and aged


• High stiffness vs. density

• Low ductility & fracture toughness

• Low short transverse properties

• Toxic


• Can be tailored for high stiffness, high strength, and extremely low coefficient of thermal expansion

• Good In tension (e.g., pressurized tanks)

• Costly for low production volume; requires development program

• Strength depends on workmanship; usually requires individual proof testing

• Laminated composites are not as strong in compression

• Brittle; can be hard to attach

constructed of thin face sheets separated by a lightweight core; this form of construction efficiently adds bending stiffness and stability. Section 15.3 of Sarafin [1995] provides guidance for selecting the above types of structures.

We can attach structural elements with adhesive bonds, welds, or mechanical fasteners. But regardless of the selected structure type and method of attachment, much of the structural subsystem's weight will Ik in the fittings used to transfer load from one member to another.

Most composite material structures have metal end fittings or edge members attached by bonding, but the bond's strength depends on the process and workmanship. Normally, we must select a proper bonding process through development testing. We can use bolts instead; however, local stress concentrations around the fasteners can cause failure at load levels much lower than a composite material can otherwise sustain. Welding is also possible for most aluminum alloys, but heat from welding can lower material strength near welds by more than 50%. If we need stiffness more than strength, we may choose welding over mechanical joints. As with bonding, welding processes require strict development, control, and testing.

The strength of mechanical fasteners, such as rivets and bolts, is very dependable as a result of process controls, inspections, and frequent sample testing. But to fully

TABLE 11-52. Design Properties tor Commonly Used Metals [MIL-HDBK-5G, 1994]. The design allowable stresses given here are statistically guaranteed for at least 99% of ail material specimens. Strengths shown are for the longitudinal direction (rolling or extrusion direction) of the material; strengths are usually lower in the long-transverse (across width) and short-transverse (through thickness) directions.

TABLE 11-52. Design Properties tor Commonly Used Metals [MIL-HDBK-5G, 1994]. The design allowable stresses given here are statistically guaranteed for at least 99% of ail material specimens. Strengths shown are for the longitudinal direction (rolling or extrusion direction) of the material; strengths are usually lower in the long-transverse (across width) and short-transverse (through thickness) directions.

Material Alloy and Form

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