P

Mdry = spacecraft dry weight

Primary power is distributed in most unmanned spacecraft as low-voltage direct current But for systems with power needs above 10 kW, we should consider alternating current distribution, both sine wave and square wave, at several hundred volts.

Radioisotope thermoelectric generators (RTGs) have been designed for various power levels but have been applied only to low-power needs. In practice, the units consist of a radioisotope heat source which can produce power by thermoelectrics or provide thermal energy to a rotating generator. If we use one of these units, we must dispose of excess heat during all mission phases and particularly during launch preparation and boost We must also consider safety issues, but RTG sources are probably safer than most propellants. The design must ensure that the generator remains intact and shielded even during catastrophic launch failure.

Using rotating machines to generate primary power is another design with potential. Gosed-cycle, thermal engines should be nearly twice as efficient as solar cells, and rotating generators can provide sine-wave AC power with better regulation than solar-cell designs.

10.4.7 Structures and Mechanisms

The spacecraft structure carries and protects the spacecraft and payload equipment through the launch environment and deploys the spacecraft after orbit injection. The load-carrying structure of a spacecraft is primary structure, whereas brackets, closeout panels, and most deployable components are secondary structure.

We size primary structure based on the launch loads, with strength and stiffness dominating its design. The size of secondary structure depends on on-orbit factors rather than boost-phase loads. Secondary structure only has to survive but not function during boost, and we can usually cage and protect deployables throughout this phase.

Each of the launch boosters provides maximum acceleration levels to be used for design (see Chap. 18). These acceleration levels or load factors are typically 6 g's maximum axial acceleration and 3 g's maximum lateral acceleration. These levels work for conceptual design, but some designers prefer to increase them by as much as 50% during early design phases. During preliminary sizing, we must remember that the primary structure must carry some weight, such as kick motors and propellant, which will drop away before orbit injection. Section 11.6 discusses structural design and presents methods for preliminary structural sizing.

We use cylindrical and conical shell structures and trusses for primary structure, commonly building them out of aluminum and magnesium with titanium for end fittings and high-strength attachments. Composite materials have seen limited use in primary structure to date but they will become more common. We can size primary structure by modeling it as a cylindrical beam which is mass loaded by its own weight and the spacecraft's components. The lateral load factors applied to this beam produce a moment that is a function of axial location. Compression in the extreme section of the beam carries the moment By adding the axial load to the moment-induced, compressive load, we can estimate the critical load, which in turn sizes the primary structure (see Eq. 11-42). In these preliminary sizing calculations, we can exercise much license in assuming symmetry and in simplifying the loads. We can iterate the skin gage to withstand stress levels and check the tubular design for buckling (see Sec. 11.6.6).

We use a similar approach to size a truss-based primary structure. We reduce the truss to its simplest form by successively removing redundant members until we reach a statically determinant structure. Simply combining loading conditions allows us to size the truss members.

We must also locate and mount components on the basic, load-carrying cylinder or truss. Most electronic components have rectangular symmetry and are mounted with lugs or bosses integral to their housings. Mounting requirements include loads, good thermal contact with the mounting surface, and good electrical contact. Aluminum honeycomb is an excellent mount for components. It attaches to longeron-stringer frames to form a semi-monocoque (load-carrying skin) structure. Honeycomb sheets with composite faces occasionally substitute for other approaches.

Some components are not rectangular. For example, propellant tanks are normally spherical but may be elongated, have conical sections, or be toroidal. Electromechanical drives and reaction wheels are cylindrical, and control moment gyros are complex. These components include mounting provisions in their designs. Generally they have flanges, bosses, or lugs. In most cases, and particularly in the case of tanks and pressure vessels^ the mounting must avoid loading the component To do so, the mount must be statically determinant and component loads from deflection of the mount must be minimal.

Other components of complex geometric shape, such as thrusters and connectors, may mount through brackets specifically tailored to them. Hinges and similar items are machined fittings with integral flanges or mounting bosses. We can align components by shimming, but we must be careful not to disturb thermal and electrical bonds.

A set of data on spacecraft subsystem mass is presented in Appendix A. These data show the structural mass to be approximately 20% of the spacecraft total. However these data do not include all of the injected mass (apogee kick motors carried in the spacecraft are not included). Therefore one should be careful about using these data for estimating new designs. However, structural mass of 10% to 20% of spacecraft dry mass is a reasonable starting point

We must have an interstage structure to mount the spacecraft to the booster. This structure conforms to the booster provisions for mounting and carries loads during the boost phase. Both truss structures and conical adapters are common. Because this structure is designed for strength under high loads, it is an excellent candidate for high-strength materials and weight-efficient design. The spacecraft usually provides this interstage structure and incorporates a separation joint to release the spacecraft at orbit

Common methods of attachment at the separation plane are marmon clamps or separation bolts. In the marmon clamp, the separation joint is a continuous ring held together by an annular clamp. Release of clamp tension allows the joint to separate. In the case of separation bolts, the joint is held by several bolts which are released by either severing the bolt or by releasing a nut. Once the separation joint is free, springs impart a small velocity increment to the spacecraft After separation, the booster maneuvers to avoid accidental impact For spin-stabilized spacecraft, the interstage structure may incorporate a mechanism to impart spin while ejecting the spacecraft.

The Shuttle interfaces differently from the expendable boosters. It links with its payloads at a series of hard points located along the sill of the cargo bay and along the cargo bay's centerline (keelfittings). The payload and its upper stages usually require a cradle or fittings to translate the loads into these hard points. Mechanisms for deploying the spacecraft may be spring-powered or motor-driven. Chapter 11 presents weight-estimating relations for motor-driven mechanisms. Spring-powered mechanisms must meet stiffness requirements, but they weigh about half as much as their motor-driven equivalents.

0 0

Post a comment