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Command and data handling systems are generally conservative, evolutionary designs due to their mission-critical nature. The baselining process presented provides the mission designer with an approach to making a first-order estimate of the necessary hardware based upon previous hardware developments. This approach will provide a realistic estimate to be used in mission resource budgeting.

It is important to identify the rationale and drivers for the baseline. If possible, each specification should be allocated a rating or confidence factor to indicate if the specification is required, flexible, or merely a place holder. This information may allow the mission design team more creativity in solving a given technical problem.

1133 C&DH Basics

This section is a list of details of great concern to command and data handling system design and operation. Many of these concerns are of absolute necessity when determining C&DH requirements and generating procurement specifications. Emphasis is placed on the command system because of the severity of the effects if these guidelines are not followed. Data handling basics such as data rates and the number of bits per sample are covered in Sec. 13.2.

Interfaces to other equipment must be protected so that their faults do not propagate into the command decoder.

It is paramount that no commands or any transient signals appear on command outputs during application ot removal of prime poweT, or during undeT/over prime power voltage conditions.

It is a basic philosophy of command decoder designs that if the integrity of a command message is in doubt, the command is not issued. It is rejected! This is especially true when firing an ordnance device or the spacecraft is launched from a manned vehicle. It is for this reason that received command messages are not corrected, although the capability exists, using error check bits.

For safety concerns, operations such as firing ordnance, an engine, or thruster, require multiple commands configured in series forming a logical AND function. No single command causes the operation to occur. In a typical ordnance application, three commands are required: safe, arm, and fire. In this case, safe and arm are relays that enable a high level discrete command, fire. The commands must (shall) be isolated within the command decoder such that no single component or physical failure results in inadvertent function execution. To achieve this, the Hamming distance of controlling command messages must be two or greater (for isolation in the decoding scheme), and command outputs must be physically isolated to the greatest extent possible using different decoding circuits and interface connectors.

It is advised not to have any commands that turn a command decoder off during flight In addition, there should be no commands that interrupt the uplink source to the command decoder.

In redundant applications, where command outputs are cross strapped, the interface circuits and interconnection have to be designed such that no single component or physical failure prevents the active output from functioning. Along the same lines, where telemetry inputs and serial interface outputs are cross strapped, the interface circuits and interconnections have to be designed such that no single component or physical failure prevents the interface from functioning.

The rising and falling edges of discrete command and serial telemetry outputs are often limited in frequency content so that they are not a source of noise emissions on the spacecraft

113.4 A Final Note

The C&DH subsystem is often one of the last on the spacecraft to be defined. It is a tool, used to configure, control, or program the payload and other spacecraft subsystems. It is the spacecraft's senses reporting internal environment, health, and status information. C&DH equipment cannot be completely defined until the requirements of other systems have been established. The mission designer's main task is that of listing the command, telemetry and other data needs for each spacecraft system. The list must also include the rate at which commands are issued and telemetry is gathered for determination of composite data rates. Issues such as data format, encoding, and security must then be addressed. At this point it may be advantageous to stop and take an overall view of the spacecraft for other functions, which if included in the C&DH, would simplify overall design. Remember that the C&DH interfaces to nearly all spacecraft functions. Next the impact of the mission environments, duration and required reliability on the C&DH hardware is assessed. When these tasks are complete the C&DH subsystem can be fully characterized.

11.4 Power

Joseph K. McDermott, Lockheed Martin Astronautics

As illustrated in Fig. 11-8, the electrical power subsystem (EPS) provides, stores, distributes, and controls spacecraft electrical power. Table 11-30 lists typical functions performed by the EPS. The most important sizing requirements are the demands for average and peak electrical power and the orbital profile (inclination and altitude). We must identify the electrical power loads for mission operations at beginning-of-life, BOL, and ertd-of-life, EOL.

For many missions, the end-of-life power demands must be reduced to compensate for solar array performance degradation. The average electrical power needed at EOL determines the size of the power source. Section 10.3 shows a sample power budget that we may use to begin the sizing process. We usually multiply average power by 2 or 3 to obtain peak power requirements for attitude control, payload, thermal, and EPS (when charging the batteries). Fortunately, all the systems do not require peak power at the same time during the mission.

Fig. 11-8. Functional Breakdown lor the Spacecraft's Power Subsystem. We start with these four functions and must determine requirements for the hardware, software, and interfaces for each.

Fig. 11-8. Functional Breakdown lor the Spacecraft's Power Subsystem. We start with these four functions and must determine requirements for the hardware, software, and interfaces for each.

TABLE 11-30. Typical Top-Level Power Subsystem Functions. Each of these functions consists of subfurtctions with a myriad design characteristics which we must develop to meet mission requirements.

• Supply a continuous source of electrical power to spacecraft loads during the mission life.

• Control and distribute electrical power to the spacecraft.

• Support power requirements for average and peak electrical load.

• Provide converters for ac and regulated dc power buses, if required.

• Provide command and telemetry capability for EPS health and status, as well as control by ground station or an autonomous system. "

• Protect the spacecraft payload against failures within the EPS.

• Suppress transient bus voltages and protect against bus faults.

• Provide ability to fire ordnance, if required.

Table 11-31 summarizes the power subsystem design process, which we discuss further in the following subsections, and Table 11-32 shows the principal effects of mission requirements on the power system design. We will work through the design process, beginning with die selection of a power source.

TABLE 11-31. The Preliminary Design Process for the Power Subsystem. All of these design steps must link back to mission requirements to satisfy the owner and users. Note that derived requirements may Impact previous design decisions and force designers to iterate the design process.

Step

Information Required

Derived Requirements

References

1. Identify Requirements

Top-level requirements, mission type (LEO, GEO), spacecraft configuration, mission life, payload definition

Design requirements, spacecraft electrical power profile (average and peak)

Sees. 10.1,10.2

2. Select and Size Power Source

Mission type, spacecraft configuration, average load requirements for electrical power

EOL power requirement, type of solar cell, mass and area of solar array, solar array configuration (2-axis tracking panel, body-mounted)

Sees. 10.1,10.2 Table 10-9 Sec. 11.4.1 Table 11-34

3. Select and Size Energy Storage

Mission orbital parameters, average and peak load requirements for electrical power

Eclipse and load-leveling energy storage requirement (battery capacity requirement), battery mass and volume, battery type

Sec. 11.4.2 Tables 11-3,11-4, 11-38,11-39,11-40 Fig. 11-11

4. Identify Power Regulation and Control

Power-source selection, mission life, requirements for regulating mission load, and thermal-control requirements

Peak-power tracker or direct-energy-transfer system, thermal-control requirements, bus-voltage quality, power control algorithms

Sec. 11.4.4

TABLE 11-32. Effects of System-Level Parameters on the Power Subsystem. Most aspects of the mission affect the power subsystem because so many other subsystems require specific power attributes.

Parameter

Effects on Design

Average Electrical Power Requirement

Sizes the power-generation system (e.g., number of solar cells, primary battery size) and possibly the energy-storage system given the eclipse period and depth of discharge

Peak Electrical Power Required

Sizes the energy-storage system (e.g., number of batteries, capacitor bank size) and the power-processing and distribution equipment

Mission Life

Longer mission life (> 7 yr) implies extra redundancy design, independent battery charging, larger capacity batteries, and larger arrays

Orbital Parameters

Defines incident solar energy, eclipse/Sun periods, and radiation environment

Spacecraft Configuration

Spinner typically implies body-mounted solar ceils; 3-axis stabilized typically implies tody-fixed and deptoyable solar panels

11.4.1 Power Sources

The power source generates electrical power within the spacecraft. Launch vehicles such as Titan IV or Delta use primary batteries (discussed in Sec. 11.4.2) as the power source for electrical loads because the batteries usually need to last less than an hour. But batteries alone are too massive for missions that last from weeks to years. These missions need a source that can generate power over many orbital cycles to support electrical loads and recharge the batteries.

Typically, we use four types of power sources for spacecraft Photovoltaic solar cells, the most common power source for Earth-orbiting spacecraft, convert incident solar radiation directly to electrical energy. Static power sources use a heat source —typically plutonium-238 or uranium-235 (nuclear reactor), for direct thermal-to-electric conversion. Dynamic power sources also use a heat source—typically concentrated solar radiation, plutonium-238, or enriched uranium—to produce electrical power using the Brayton, Stirling, or Rankine cycles. The fourth power source is fuel cells, used on manned space missions such as Gemini, Apollo, SkyLab, and the Space Shuttle. Table 11-33 provides a comparison of various power sources.

Static power conversion uses either a thermoelectric or a thermionic concept The most common static power source for spacecraft is the thermoelectric couple. This basic converter uses the temperature gradient between the p-n junction of individual thermoelectric cells connected in a series-parallel arrangement to provide the desired dc electrical output from each converter. This temperature gradient comes from slow decay of the radioactive source. The thermal-to-electric conversion efficiency for a thermoelectric source is typically 5-8%.

Thermionic energy conversion produces electricity through a hot electrode (emitter•) facing a cooler electrode (collector) inside a sealed enclosure that typically contains an ionized gas. Electrons emitted from the hot emitter flow across the inter-electrode gap to the cooler collector. There they condense and return to the emitter through the electrical load connected externally between the collector and the emitter. We choose the collector and emitter temperatures for best overall system performance. In choosing the collector temperature, we try to decrease the weight and size of thermal radiators, and we choose materials based on mission life requirements. Thermionic power sources usually rely on a reactor heat source because of the high temperature required for efficient thermionic conversion. Power efficiencies for a thermionic power conversion are typically 10-20%.

In contrast to static sources, dynamic power sources use a heat source and a heat exchanger to drive an engine in a thermodynamic power cycle. The heat source can be concentrated solar energy, radioisotopes, or a controlled nuclear-fission reaction. Heat from the source transfers to a working fluid, which drives an energy-conversion heat engine. For a dynamic solar-power source, the balance of energy remains as latent and sensible heat in a heat exchanger (molten eutectic salt), which provides continuous energy to the thermodynamic cycle during eclipse periods. A dynamic power source using a nuclear reactor or plutonium-238 decay does not require thermal-energy storage because the source provides continuous heat.

Dynamic power sources use one of three methods to generate electrical power. Stirling cycle, Rankine cycle, or Brayton cycle. Stirling-cycle engines use a single-phase working fluid as the working medium. The thermodynamic cycle consists of two isothermal processes (compression and expansion) and two constant-volume processes (heating and cooling). Power-conversion efficiencies for Stirling engines are 23-30%. Ranltine-cycle engines are dynamic devices that use a two-phase fluid system

TABLE 11-33. Matrix for Comparing Most Common Spacecraft Power Sources. We may use different factors to select the correct power source but specific power and specific cost are used extensively.

EPS Design Parameters

Solar Photovoltaic

Solar Thermal Dynamic

Radio-Isotope

Nuclear Reactor

Fuel Cell

Power Range (kW)

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