Spacecraft Subsystems

11.1 Attitude Determination and Control

Control Modes and Requirements; Selection of Spacecraft Control Type; Quantify the Disturbance Environment; Select and Size ADCS Hardware; Define the Control Algorithms

11.2 Telemetry, Tracking, and Command Requirements; Designing the TT&C Subsystem

11.3 Command and Data Handling

Introduction to C&DH; C&DH System Sizing Process; C&DH Basics; A Final Note

11.4 Power

Power Sources; Energy Storage; Power Distribution; Power Regulation and Control

11.5 Thermal

Spacecraft Thermal Environment; Thermal Control Components; The Thermal Design and Development Process; Thermal Control Challenges; Heat Balance Estimation; Mass, Power, Telemetry Estimates

11.6 Structures and Mechanisms

Structural Requirements; Packaging and Configuring the Subsystem; Design Options; Structural Design Philosophy and Criteria; Preliminary Sizing of Structural Members; Structural Mechanics and Analysis; An Example Problem; Mechanisms and Déployables

11.7 Guidance and Navigation

System Definition Process; Orbit Determination Systems; Orbit Maintenance and Control; Sizing Autonomous Guidance and Navigation

This chapter provides design information for the spacecraft bus subsystems, emphasizing material most pertinent to the spacecraft engineer. It offers practical insight into the mission and interface requirements that drive how we configure spacecraft We include first-order approximations and describe hardware to show how each subsystem works and to help estimate the subsystem's size, weight, power requirements, and eventual cost. We also reference many chapters of this book to integrate concepts and subsystems. Chapter 17 discusses the propulsion subsystem, and Chap. 13 provides much of the communications theory. Chapter 10, Agrawal [1986], Chetty [1991], and Morgan and Gordon [1989] provide insight to the theory and practice of designing spacecraft subsystems.

In the rest of this chapter we will discuss these issues and an approach for estimating the size and configuration of spacecraft subsystems.

11.1 Attitude Determination and Control John S. Eterno, Ball Aerospace & Technologies Corporation

The attitude determination and control subsystem (ADCS) stabilizes the vehicle and orients it in desired directions during the mission despite the external disturbance torques acting on it This requires that the vehicle determine its attitude, using sensors, and control it, using actuators. The ADCS often is tightly coupled to other subsystems on board, especially the propulsion (Chap. 17) and navigation (Sec. 11.7) functions. Additional information on attitude determination and control can be found in Wertz [1978, 2001], Kaplan [1976], Agrawal [1986], Hughes [1986], Griffin and French [1990], Chobotov [1991], and Fortescue and Stark [1992],

We begin by discussing several useful concepts and definitions, including mass properties, disturbance torques, angular momentum, and reference vectors. The mass properties of a spacecraft are key in determining the size of control and disturbance torques. We typically need to know the location of the center of mass or gravity (eg) as well as the elements of the inertia matrix: the moments and products of inertia about chosen reference axes. (See Sec. 11.6 for examples of moment of inertia calculations.) The direction of the principal axes—those axes for which the inertia matrix is diagonal and the products of inertia are zero—are also of interest. Finally, we need to know how these properties change with time, as fuel or other consumables are used, or as appendages are moved or deployed.

A body in space is subject to small but persistent disturbance torques (e.g., 10-4 N-m) from a variety of sources. These torques are categorized as cyclic, varying in a sinusoidal manner during an orbit, or secular, accumulating with time, and not averaging out over an orbit. These torques would quickly reorient the vehicle unless resisted in some way. An ADCS system resists these torques either passively, by exploiting inherent inertia or magnetic properties to make the "disturbances" stabilizing and their effects tolerable, or actively, by sensing the resulting motion and applying corrective torques.

Angular momentum plays an important role in space, where torques typically are small and spacecraft are unconstrained. For a body initially at rest, an external torque will cause the body to angularly accelerate proportionally to the tOTque—resulting in an increasing angular velocity. Conversely, if die body is initially spinning about an axis perpendicular to the applied torque, then the body spin axis will precess, moving with a constant angular velocity proportional to the torque. Thus, spinning bodies act like gyroscopes, inherently resisting disturbance torques in 2 axes by responding with constant, rather than increasing, angular velocity. This property of spinning bodies, cal led gyroscopic stiffness, can be used to reduce the effect of small, cyclic disturbance torques. This is true whether the entire body spins or just a portion of it, such as a momentum wheel or spinning rotor.

Conservation of vehicle angular momentum requires that only external torques change the system net angular momentum. Thus, external disturbances must be resisted by external control torques (e.g., thrusters or magnetic torquers) or the resulting momentum buildup must be stored internally (e.g., by reaction wheels) without reorienting the vehicle beyond its allowable limits. The momentum buildup due to secular disturbances ultimately must be reduced by applying compensating external control torques.

Often, in addition to rejecting disturbances, the ADCS must reorient the vehicle (in slew maneuvers) to repoint the payload, solar arrays, or antennas. These periodic repointing requirements may drive the design to larger actuators than would be required for disturbance rejection alone.

To orient the vehicle correctly, external references must be used to determine the vehicle's absolute attitude. These references include the Sun, the Earth's IR horizon, the local magnetic field direction, and the stars. In addition, inertial sensors (gyroscopes) also can be carried to provide a short-term attitude reference between external updates. External references (e.g., Sun angles) are usually measured as body-centered angular distances to a vector. Each such vector measurement provides only two of the three independent parameters needed to specify the orientation of the spacecraft This results in the need for multiple sensor types on board most spacecraft.

Table 11-1 lists the steps for designing an ADCS for spacecraft The FireSat spacecraft, shown in Fig. 11-1, will be used to illustrate this process. The process must be iterative, with mission requirements and vehicle mass properties closely related to the ADCS approach. Also, a rough estimate of disturbance torques (see Chap. 10) is necessary before the type of control is selected (step 2), even though the type of control will help determine the real disturbance environment (step 3).

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one optional 30 deg maneuver per month to a target of opportunity

Fig. 11-1. Hypothetical Fire Sat Spacecraft. We use this simplified example to discuss key concepts throughout this section. See Fig. 5-1 for illustration of roll-pitch-yaw coordinates.

one optional 30 deg maneuver per month to a target of opportunity

Fig. 11-1. Hypothetical Fire Sat Spacecraft. We use this simplified example to discuss key concepts throughout this section. See Fig. 5-1 for illustration of roll-pitch-yaw coordinates.

11.1.1 Control Modes and Requirements

Tables 11-2 and 11-3 describe typical spacecraft control modes and requirements. The ADCS requirements are closely tied to mission needs and other subsystem characteristics, as shown in Fig. 11-2. These requirements may vary considerably with mission phase or modes, challenging the designer to develop a single hardware suite for different objectives.

For many spacecraft, the ADCS must control vehicle attitude during firing of large liquid or solid rocket motors, which may be used during orbit insertion or for orbit changes. Large motors create large disturbance torques, which can drive the design to larger actuators than are needed once on station.

TABLE 11-1. Control System Design Process. An iterative process is used for designing thi ADCS as part of the overall spacecraft system.

Step

Inputs

Outputs

FireSat Example

1a. Define control modes 1b. Defineor derive system-level requirements by control mode

Mission requirements, mission profile, type of insertion for launch vehicle

List of different control modes during mission (See Table 11-2)

Requirements and constraints (See Table 11-3)

Orbit Injection: none—provided by launch vehicle

Normal: nadir pointing, <0.1 deg; autonomous determination (Earth-relative)

Optional slew: One 30 deg maneuver per month to a target of opportunity

2. Select type of spacecraft control by attitude control mode (Sea 11.1.2)

Payload, thermal and power needs

Orbit, pointing direction

Disturbance environment

Method for stabilizing and control: 3-axis, spinning, or gravity gradient

Momentum bias stabilization with a pitch wheel, electromagnets for momentum dumping, and optionally, thrusters for slewing (shared with a/ system In navigation)

3. Quantify disturbance environment (Sec. 11.1.3)

Spacecraft geometry, orbit, solar/magnetic models, mission profile

Values for forces from gravity gradient, magnetic aerodynamics, solar pressure, Internal disturbances, and powered flight effects on control (eg offsets, slosh)

Gravity gradient 1.8 x 10"8 N-m normal pointing; 4.4 x 10-® N-m during target-of-opportunity mode

Magnetic: 4.5 x 10~6 N-m Solar. 6.6 x 10-8 N-m Aerodynamic: 3.4 x 10-® N-m

4. Select and size ADCS hardware (Sec. 11.1.4)

Spacecraft geometry, pointing accuracy, orbit conditions, mission requirements, lifetime, orbit, pointing direction, slew rates

Sensor suita- Earth, Sun, inertial, or other sensing devices

Control actuators, e.g., reaction wheels, thrusters, or magnetic torquers

Data processing electronics, if any, or processing requirements for other subsystems or ground computer

1 Momentum wheel, Momentum: 40 N-m-s

2 Horizon sensors, Scanning, 0.1 deg accuracy

3 Electromagnets, Dipole moment: 10 A-m2

4 Sun sensors, 0.1 deg accuracy

1 3-axis magnetometer, 1 deg accuracy

5. Define determination and control algorithms

AD of above

Algorithms, parameters, and logic for each determination and control mode

Determination: Horizon data filtered for pitch and roll. Magnetometer and Sun sensors used for yaw.

Control: Proportional-plus-derivative for pitch, Coupled roll-yaw control with electromagnets

S. Iterate and document

AO of above

Refined requirements and design

Subsystem specification

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