Chapter

Emery I. Reeves, United States Air Force Academy

10.1 Requirements, Constraints, and the Design Process

10.2 Spacecraft Configuration

10.3 Design Budgets

10.4 Designing the Spacecraft Bus

Propulsion Subsystem; Attitude Determination and Control Subsystem; Communications Subsystem; Command and Data Handling Subsystem; Thermal Subsystem; Power Subsystem; Structures and Mechanisms

10.5 Integrating the Spacecraft Design Spacecraft Size; Lifetime and Reliability

10.6 Examples

Over the past four decades the engineering design of spacecraft has evolved from infancy to well-defined techniques supported by analysis tools, manufacturing technology, and space-qualified hardware. This chapter summarizes these techniques, with emphasis on the conceptual design of the spacecraft vehicle. The following two chapters present more detailed design and manufacturing information. To design a spacecraft, we must understand the mission, including the payload's size and characteristics, plus significant system constraints such as orbit, lifetime, and operations. We then configure a space vehicle to cany the payload equipment and provide the functions necessary for mission success. The design process shown in Table 10-1 involves identifying these functions, choosing candidate approaches for each function, and selecting the best approaches. This chapter presents design methods with rules of thumb that will help us roughly estimate the spacecraft design [Agrawal, 1986; Chetty, 1991; Griffin and French, 1991].

An unmanned spacecraft consists of at least three elements: a payload, a spacecraft bus, and a booster adapter. The payload is the mission-peculiar equipment or instruments. The spacecraft bus carries the payload and provides housekeeping functions. The payload and spacecraft bus may be separate modules, or the vehicle may be an integrated design. The booster adapter provides the load-carrying interface with the boost vehicle. The spacecraft may also have a propellant load and a propulsion kick stage. The propellant, either compressed gas, liquid or solid fuel, is used for velocity corrections and attitude control. A kick stage,* if used, is a separate rocket motor or liquid stage used to inject the spacecraft into its mission orbit

* Also called apogee boost motor, propulsion module, or integral propulsion stage.

TABLE 10-1. Overview of Spacecraft Design and Sizing. The process is highly iterative, normally requiring several cycles through the table even for preliminary designs.

Step

References

1. Prepare list of design requirements and constraints

2. Select preliminary spacecraft design approach and overall configuration based on the above list

3. Establish budgets for spacecraft propellant, power, and weight

4. Develop preliminary subsystem designs

5. Develop baseline spacecraft configuration

6. Iterate, negotiate, and update requirements, constraints, design budgets

Sec. 10.3 Sec. 10.4 Sees. 10.4,10.5 Steps 1 to 5

The top-level requirements and constraints are dictated by the mission concept, mission architecture, and by payload operation. For instance, the selection of orbit is intimately tied to the selected mission and payload as described in Chaps. 6 and 7. From a spacecraft design standpoint, the orbit also affects attitude control, thermal design, and the electric power subsystem. However, most of these design effects are secondary to the effect that the orbit can have on payload performance. The designer therefore selects the orbit based on mission and payload performance, and computes the required spacecraft performance characteristics such as pointing, thermal control, power quantity, and duty cycle. The spacecraft is then sized to meet these requirements. We can summarize succinctly the spacecraft bus functions: support the payload mass; point the payload correctly; keep the payload at the right temperature; provide electric power, commands, and telemetry; put the payload in the right orbit and keep it there; and provide data storage and communications, if required. The spacecraft bus consists of subsystems or equipment groups which provide these functions. Table 10-2 lists the somewhat arbitrary definitions of subsystems used here and in Chap. 11. The table also includes alternate terminology and groupings you may encounter, along with references to more detailed information. Sometimes the payload is also treated as a subsystem. Chapters 9 and 13 discuss payload design.

The propulsion subsystem provides thrust for changing the spacecraft's transla-tional velocity or applying torques to change its angular momentum. The simplest spacecraft do not require thrust and hence have no propulsion equipment But most spacecraft need some controlled thrust so their design includes some form of metered propulsion—a propulsion system that can be turned on and off in small increments. We use thrusting to change orbital parameters, correct velocity errors, maneuver, counter disturbance forces (e.g., drag), control attitude during thrusting, and control and correct angular momentum. The equipment in the propulsion subsystem includes a propellant supply (propellant tankage, distribution system, pressurant and propel-lant controls) and thrusters or engines. Compressed gasses, such as nitrogen, and liquids, such as monopropellant hydrazine, are common propellants. Significant sizing parameters for the subsystem are the total impulse and the number, orientation, and thrust levels of the thrusters. Chapter 17 describes design and equipment for propulsion subsystems.

The attitude determination and control subsystem measures and controls the spacecraft's angular orientation (pointing direction), or, in the case of a guidance, navigation, and control system, both its orientation and linear velocity (which affects its orbit). The simplest spacecraft are either uncontrolled or achieve control by passive

TABLE 10-2. Spacecraft Subsystems. A spacecraft consists of functional groups of equipment or subsystems.

Subsystem

Principal Functions

Other Names

References

propulsion

Provides thrust to adjust orbit and attitude, and to manage angular momentum

Reaction Control System (RCS)

Sec. 10.4.1, Chap. 17

Attitude

Determination & Control System (ADCS)

Provides determination and control of attitude and orbit position, plus pointing of spacecraft and appendages

Attitude Control System (ACS), Guidance, Navigation, & Control (GN&C) System, Control System

Sees.

10.42,11.1, 11.7

Communication (Comm)

Communicates with ground & other spacecraft; spacecraft tracking

Tracking, Telemetry, & Command (TT&C)

Sees.

10.4.3,112

Command & Data Handling (CSDH)

Processes and distributes commands; processes, stores, and formats data

Spacecraft Computer System, Spacecraft Processor

Sees.

104.4,11.3, Chap. 16

Thermal

Maintains equipment within allowed temperature ranges

Environmental Control System

Sees.

10.4.5,11.5

Power

Generates, stores, regulates, and distributes electric power

10.4.6,11.4

Structures and Mechanisms

Provides support structure, booster adapter, and moving parts

10.4.7,11.6

methods such as spinning or interacting with the Earth's magnetic or gravity fields. These may or may not use sensors to measure the attitude or position. More complex systems employ controllers to process the spacecraft attitude, and actuators, torquers, or propulsion subsystem thrusters to change attitude, velocity, or angular momentum. Spacecraft may have several bodies or appendages, such as solar arrays or communication antennas, that require individual attitude pointing. To control the appendages' attitude, we use actuators, sometimes with separate sensors and controllers. The capability of the attitude control subsystem depends on the number of body axes and appendages to be controlled, control accuracy and speed of response, maneuvering requirements, and the disturbance environment. Section 11.1 discusses design of the attitude determination and control subsystem.

The communications subsystem links the spacecraft with the ground or other spacecraft. Information flowing to the spacecraft (uplink or forward link) consists of commands and ranging tones. Information flowing from the spacecraft (downlink or return link) consists of status telemetry and ranging tones and may include payload data. The basic communication subsystem consists of a receiver, a transmitter, and a wide-angle (hemispheric or omnidirectional) antenna. Systems with high data rates may also use a directional antenna. The communications subsystem receives and demodulates commands, modulates and transmits telemetry and payload data, and receives and retransmits range tones—modulation that allows signal turnaround time delay and hence range to be measured. The subsystem may also provide coherence between uplink and downlink signals, allowing us to measure range-rate Doppler shifts. We size the communications subsystem by data rate, allowable error rate, communication path length, and RF frequency. Section 11.2 and Chap. 13 discuss design of the communications subsystem.

The command and data handling subsystem distributes commands and accumulates, stores, and formats data from the spacecraft and payload. For simpler systems, we combine these functions with the communications subsystem as a tracking, telemetry, and command subsystem. This arrangement assumes that distributing commands and formatting telemetry are baseband extensions of communications modulation and demodulation. In its more general form, the subsystem includes a central processor (computer), data buses, remote interface units, and data storage units to implement its functions. It may also handle sequenced or programmed events. For the most part, data volume and data rate determine the subsystem's size. Section 11.3 discusses subsystem design, and Chap. 16 covers computers and software.

The power subsystem provides electric power for the equipment on the spacecraft and payload. It consists of a power source (usually solar cells), power storage (batteries), and power conversion and distribution equipment. The power needed to operate the equipment and the power duty cycle determine the subsystem's size, but we must also consider power requirements during eclipses and peak power consumption. Because solar cells and batteries have limited lives, our design must account for power requirements at begirming-of-life (BOL) and end-qf-life (EOL). Section 11.4 discusses design of the power subsystem.

The thermal subsystem controls the spacecraft equipment's temperatures. It does so by the physical arrangement of equipment and using thermal insulation and coatings to balance heat from power dissipation, absorption from the Earth and Sun, and radiation to space. Sometimes passive, thermal-balance techniques are not enough. In this case, electrical heaters and high-capacity heat conductors, or heat pipes, actively control equipment temperatures. Hie amount of heat dissipation and temperatures required for equipment to operate and survive determine the subsystem's size. Section 11.5 discusses temperature control in more detail.

The structurât subsystem carries, supports, and mechanically aligns the spacecraft equipment It also cages and protects folded components during boost and deploys them in orbit The main load-carrying structure, or primary structure, is sized by either (1) the strength needed to carry the spacecraft mass through launch accelerations and transient events during launch or (2) stiffness needed to avoid dynamic interaction between the spacecraft and the launch vehicle structures. Secondary structure, which consists of déployables and supports for components is designed for compact packaging and convenience of assembly. Section 11.6 discusses structural design.

10.1 Requirements, Constraints, and the Design Process

In designing spacecraft, we begin by developing baseline requirements and constraints such as those in Table 10-3. If some of the information is not available, we may need to assume values or use typical values such as those presented here or in the following chapters. For successful design, we must document all assumptions and revisit them until we establish an acceptable baseline.

To get a feel for the size and complexity of a spacecraft design, we must understand the space mission: its concept of operations, duration, overall architecture, and constraints on cost and schedule. Even if we select a mission concept arbitrarily from several good candidates, clearly defining it allows us to complete the spacecraft design and evaluate its performance.

The payload is the single most significant driver of spacecraft design. Its physical parameters—size, weight, and power—dominate the physical parameters of the

TABLE 10-3. Principal Requirements and Constraints for Spacecraft Design. These parameters typically drive the design of a baseline system.

Requirements and Constraints

Information Needed

Reference

Mission: Operations Concept Spacecraft Life & Reliability Comm Architecture Security

Programmatic Constraints

Type, mission approach Mission duration, success criteria Command, control, comm approach Level, requirements Cost and schedule profiles

Sees. 1.4,10.5,19.2 Sec. 13.1 Sees. 13.1,15.4 Chaps. 1,20

Payload: Physical Parameters

Operations Pointing Slewing Environment

Size, weight, shape, power

Duty cycle, data rates, fields of view Reference, accuracy, stability Magnitude, frequency Max and min temperatures, cleanliness

Chaps. 9,13

Sec. 9.5

Orbit: Defining Parameters Eclipses

Lighting Conditions Maneuvers

Altitude, inclination, eccentricity Maximum duration, frequency Sun angle and viewing conditions Size, frequency

Chaps. 6,7 Sees. 7.4,7.5 Sea 5.1 Sees. 5.1,5.2 Sees. 6.5,7.3

Environment: Radiation Dosage Particles & Meteoroids Space Debris Hostile Environment

Average, peak Size, density

Density, probability of impact Type, level of threat

Chap. 8 Sees. 8.1,82 Sees. 8.1,21.2 Sec. 212 Sec. 8 2

Launch: Launch Strategy

Boosted Weight Envelope Environments Interfaces Launch Sites

Single, dual; dedicated, shared; use of upper kick stage

Launch capabilities

Size, shape g's, vibration, acoustics, temperature Electrical and mechanical Locations, allowed launch azimuths

Sees. 18.1,18.2 Sec. 18.3 Sec. 18.3 Sec. 18.3 Sec. 18.1

Ground-System Interface: Degree of Autonomy Ground Stations Space Links

Required autonomous operations Number, locations, performance Space-to-space link, performance

Chaps. 14,15 Sees. 15.4,16.1 Sees. 15.1,15.5 Sees. 13.3,13.4

spacecraft. Payload operations and support are key requirements for the spacecraft's subsystems, as well. The payload may also impose significant special requirements that drive the design approach, such as cryogenic temperatures or avoidance of contamination. Fortunately, we often understand the payload's characteristics better than the spacecraft's overall characteristics in the early design phases. Thus, we can infer many important design features by understanding the payload and how it operates.

Chapters 6 and 7 show how orbital characteristics affect the mission. In spacecraft design, the orbit affects propulsion, attitude control, thermal design, and the electric power subsystem. Most of these design effects are secondary to the orbit's effect on payload performance. Therefore, we select the orbit based on mission and payload performance. Then we compute the performance characteristics needed for the spacecraft, such as pointing, thermal control, and power quantity and duty cycle. Finally, we size the spacecraft to meet these needs.

The natural space environment—especially radiation—limits two aspects of spacecraft design: usable materials or piece parts, and spacecraft lifetime. Radiation levels and dose must be considered in the design, but they do not normally affect the system's configuration or ability. Chapter 8 provides useful space environment information. However, some types of hostile (weapon) environments may affect countermeasures, configuration, shielding, or maneuvering ability.

Selecting a boost vehicle and the possible use of kick stages are central issues in designing a spacecraft We must select a booster that can put at least the minimum version of our spacecraft into its required orbit Chapter 18 describes available boosters, all of which have limited weight-lifting ability. In most cases, we must extrapolate published data to meet our mission requirements. Chapters 6 and 7 present the laws of orbital mechanics and the techniques of trajectory design. These include methods for computing velocity increments and guidance techniques. In some cases, the spacecraft must provide large amounts of velocity just to reach orbit or to guide the flight path. Chapter 17 presents performance characteristics for solid and liquid propulsion kick stages to implement these functions. Common nomenclature for a kick stage used to inject a spacecraft into transfer orbit is a perigee lack motor (PKM), whereas a kick stage used to circularize at high altitude is called an apogee kick motor (AKM).

The booster selection will also affect a spacecraft's linear dimensions. An aerodynamic cover, called a fairing, or shroud, protects the spacecraft as it travels through the atmosphere. The fairing's diameter and length limit the spacecraft's size—at least while it is attached to the booster. Chapter 18 presents the size of standard fairings for various boosters. If the on-orbit spacecraft is larger than the fairing, it must be folded or stowed to fit within the fairing and unfolded or deployed on orbit To design an item with a large area but small intrinsic mass, such as a solar array or antenna, we make the item as light as possible, fold it and protect it during boost and deploy it (unfold, pull, or stretch it into shape) on orbit Solar cells may rest on lightweight substrates or even on film that is folded or rolled for storage. Antenna reflectors have consisted of folded rigid panels or of fabric, either film or mesh. Thus, we meet the launch vehicle's demands for a smaller spacecraft by using a stowed configuration and then deploying the spacecraft to meet the full size needed on orbit We use weight efficiently by caging and protecting the light-weight deployables during boost

The ground system interface determines how much ground operators and the spacecraft can interact—an important part of design. The periods of visibility between ground stations and the spacecraft limit ground control of spacecraft operations or corrections of errant behavior. Visibility periods and ground coverage issues are described in Chap. 5. If ground operations cost too much, we may want the spacecraft to operate autonomously—another major design decision.

Table 10-4 lists initial configuration decisions or trade-offs designers often face. Weight, size, and power requirements for the payload place lower limits on spacecraft's weight, size, and power. The spacecraft's overall size may depend on such payload parameters as antenna size or optical system diameter. Our approach to spacecraft design must match these dimensions and provide fields of view appropriate to the payload functions. The spacecraft must generate enough power to satisfy the payload needs as well as its own requirements. The amount of power and the duty cycle will dictate the size and shape of solar arrays and the requirements for the battery.

TABLE 10-4. Initial Spacecraft Design Decisions or Trade-offs. Further discussion of these trades is In Sec. 10.2.

Design Approach or Aspect

Where Discussed

Principal Options or Key Issues

Spacecraft Weight

Table 10-10

Must allow for spacecraft bus weight and payload weight

Spacecraft Power

Tables 10-8, 10-9

Must meet power requirements of payload and bus.

Spacecraft Size

Sec. 10.1

Is there an item such as a payload antenna or optical system that dominates the spacecraft's physical size? Can the spacecraft be folded to fit within the booster diameter? Spacecraft size can be estimated from weight and power requirements.

Attitude Control Approach

Sees. 10.2, 10.4.2,11.1

Options include no control, spin stabilization, or 3-axis control: selection of sensors and control torquers. Key issues are number of items to be controlled, accuracy, and amount of scanning or slewing required.

Solar Array Approach

Sees. 10.2, 10.4.6,11.4

Options include planar, cylindrical, and omnidirectional arrays either body mounted or offset

Kick Stage Use

Chaps. 17,18

Use of a kick stage can raise Injected weight Options include solid and liquid stages.

Propulsion Approach

Sees. 10.2, 10.4.1,17.2, 17.3

Is metered propulsion required? Options Include no propulsion, compressed gas, liquid monopropeKant or bipropellanL

Field-of-view and pointing considerations influence how we configure the spacecraft Instruments, sensors, solar arrays, and thermal radiators all have pointing and field-of-view requirements that must be satisfied by their mounting on the spacecraft and the spacecraft's orientation. In the simplest case, all items are fixed to ¿be body, and control of the body's attitude points the field of view. In more complex cases, single or two-degree-of-freedom mechanisms articulate the field of view.

We must also establish how to configure the spacecraft's propulsion early in the design process. Although interaction with the Earth's gravity or magnetic field can control attitude, it cannot change the spacecraft's velocity state. If spacecraft velocity control is needed, some form of metered propulsion must be used. If we decide to use metered propulsion, we should look at using this system for such functions as attitude control or as an orbit transfer stage. Most attitude-control systems use metered pro pulsion to exert external torque on the spacecraft Selecting a propulsion approach depends on the total impulse requirement, and the propulsion system's performance, as discussed in Chap. 17.

102 Spacecraft Configuration

To estimate the size and structure of a spacecraft we select a design approach, develop a spacecraft configuration (overall arrangement) and make performance allocations to the spacecraft subsystems. We then evaluate the resulting design and reconfigure or reallocate as needed. Subsequent iterations add design detail and provide tetter allocations. The process of allocating design requirements involves two mutually supporting techniques. First the allocated design requirements are dictated by considering the overall spacecraft design—a top-down approach. Alternatively, the allocated design requirements are developed by gathering detailed design information—a bottom-up approach. For instance, we may allocate 100 kg for structural weight based on 10% of the overall spacecraft weight This is a top-down allocation. However, a detailed design of the structure may require 120 kg if aluminum is used and 90 kg if composites are used. These are bottom-up allocations, providing us with the opportunity to trade off alternatives and reallocate requirements to optimize the design. Most of the allocation methods presented in this chapter are top-down. They provide a starting point for the allocation process. However, we should use them in conjunction with bottom-up design from the more detailed information given in Chaps. 11,16, and 17.

Figure 10-1 shows different spacecraft configurations. First observe that each of these spacecraft has a central body or equipment compartment that houses most of the spacecraft equipment. Second, note that these spacecraft all have solar arrays either mounted on external panels or on the skin of the equipment compartment. And finally note that some of the spacecraft have appendages carrying instruments or antennas attached to the main compartment Let's examine each of these configuration features in more detail.

Table 10-5 lists the factors called configuration drivers leading to the various configurations. The weight size and shape of the payload, and the boost vehicle diameter drive the size and shape of the equipment compartment Table 10-5 also presents rules of thumb based on analysis of a large number of spacecraft designs. This analysis shows that the average spacecraft bus dry weight (spacecraft weight excluding propel-lant) is approximately twice that of the payload. The minimum spacecraft bus dry weight is equal to the payload weight and is achieved only when the payload is massive and compact At the other extreme, low-density payloads or those consisting of multiple instruments can lead to a spacecraft bus as massive as 6 times the payload. Although this is a large range of possible spacecraft bus weights, these ratios are at least a bound. Section 103 shows how to refine the estimate.

The spacecraft equipment compartment volume can be estimated from its weight For 75 spacecraft launched between 1975 and 1984, the average spacecraft in launch configuration with propellant loaded and all appendages folded had a density of only 79 kg/m3 with a maximum of 172 kg/m3 and a minimum of 20 kg/m3. However, appendages are usually lightweight so the weight of the equipment compartment is only slightly less than the total spacecraft weight We can use this experience to estimate the spacecraft size (volume and dimensions) by the steps shown in Table 10-6. We start with payload weight to obtain an estimate of spacecraft bus weight (e.g.,

A. Spin-Stabilized Spacecraft

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