Sa

Gjj]

Loads

Chargers

Boost Reg

Unregulated Bus Using Parallel Batteries

Unregulated Bus Using Linear Charge Current Control Recharge Control

Quasi-Regulated Bus with Constant Current Chargers

Systems Using a Fully Regulated Bus

Batteries

Fig. 11-13. Techniques for Power Regulation. The basic approaches are Peak Power Tracking (PPT), which places a regulator hi series with the solar arrays and the load, and Direct Energy Transfer (DET), which uses a regulator in parallel with the solar arrays and load.

loads or battery charging do not need power. Power subsystems with shunt regulation are extremely efficient. They dissipate little energy by simply shunting excess power at the array or through shunt resistor banks. A shunt-regulated subsystem has advantages: fewer parts, lower mass, and higher total efficiency at EOL.

Techniques for controlling bus voltage on electrical-power subsystems fall into three categories: unregulated, quasi-regulated, or fully regulated. Figure 11-13 illustrates the main differences between these techniques. An unregulated subsystem has a load bus voltage that varies significandy. The bus-voltage regulation derives from battery regulation, which varies about 20% from charge to discharge. In an unregulated subsystem, the load bus voltage is the voltage of the batteries.

Quasi-regulated subsystems regulate the bus voltage during battery charge but not during battery discharge. A battery charger is in series with each battery or group of parallel batteries. During charge the bus voltage fixes at a potential several volts above the batteries. As the batteries reach full charge, the drop across the chargers decreases, but the bus voltage is still constantly regulated. The bus becomes unregulated during discharge when the voltage is about a diode drop lower than the batteries and decreases as the batteries further discharge. A quasi-regulated power subsystem has low efficiency and high electromagnetic interference if used with a peak-power tracker.

Th a fully regulated power subsystem is inefficient, but it will work on a spacecraft that requires low power and a highly regulated bus. This subsystem employs charge and discharge regulators. We can design the regulators so the charge regulator uses linear technology and the discharge regulator is a switching converter, but for best efficiency both should be converters. The advantage of this type of power subsystem is that, when we connect it to the loads, the system behaves like a low-impedance power supply, making design integration a simple task. But it is the most complex type of power subsystem, with an inherent low efficiency and high electromagnetic interference when used with a PPT or boost converter.

We can charge batteries individually or in parallel. A parallel charging system is simpler and has the lower cost, but does not allow flexibility in vehicle integration. It can also stress batteries so they degrade faster. When batteries are charged in parallel, the voltage is the same but the current and temperature are not Because current is not rigidly controlled, one battery could receive all the available charge current, and a thermal runaway condition could result if we do not control the bus voltage from the hottest battery. Parallel batteries eventually end up balancing out, so we could use them for missions under five years. To ensure a battery life greater than five years, we should seriously consider independent chargers, such as the linear, charge-current-control (LC3) design in Fig. 11-13.

Batteries usually limit the life of a spacecraft. To support a seven-year life, we must charge the batteries independently to degrade the battery as little as possible. Individual charging optimizes the battery use by charging all the batteries to their own unique limits. It also forgives battery deviations in systems with several batteries. Unfortunately, individual chargers add impedance, electronic piece parts, and thermal dissipation not present in a parallel system. To design the Power Regulation and Control subsystem, follow the steps in Table 11-42.

TABLE 11-42. Steps In the Power Regulation and Control Subsystem Design.

Step

Consider

Possibilities

1. Determine the power source

• All spacecraft loads, their duty cycles, and special operating modes

• Primary batteries

Photovoltaic

• Dynamic power

2. Design the electrical control subsystem

• Battery charging

• Spacecraft heating

• Peak-power tracker

• Direct-energy transfer

3. Develop the electrical bus voltage control

• How much control does each load require?

• Battery voltage variation from charge to discharge

• Battery recharge subsystem

• Battery cycle life

• Total system mass

• Unregulated

• Quasi-regulated

• Fully regulated

• Parallel or individual charging

- > 5 yrs—independent charge

11.5 Thermal*

David G. Gilmore, Brian E. Hardt, Robert C. Prager, The Aerospace Corporation Eric W. Grob, Wes Ousley, Goddard Space Flight Center

The role of the thermal control subsystem (TCS) is to maintain all spacecraft and payload components and subsystems within their required temperature limits for each mission phase. Temperature limits include a cold temperature which the component must not go below and a hot temperature that it must not exceed. Two limits are frequently defined: operational limits that the component must remain within while operating and survival limits that the component must remain within at all times, even when not powered. Exceeding survival temperature limits can result in permanent equipment damage as opposed to out-of-tolerance performance when operational limits are exceeded. Table 11-43 gives typical component temperature ranges for representative spacecraft components. Thermal control is also used to ensure that temperature gradient requirements are met. An example of a gradient requirement is to ensure that one side of a structure does not get hotter or colder than the opposite side by more than, say, 30 dC. A larger gradient could cause structural deformation such that pointing is adversely impacted, possibly permanently.

TABLE 11-43. Examples of Typical Thermal Requirements for Spacecraft Components.

The thermal control subsystem Is required to maintain all spacecraft equipment within proper temperature ranges. Note that the temperature extremes on the outer portions of spacecraft can vary between + 200 °C.

TABLE 11-43. Examples of Typical Thermal Requirements for Spacecraft Components.

The thermal control subsystem Is required to maintain all spacecraft equipment within proper temperature ranges. Note that the temperature extremes on the outer portions of spacecraft can vary between + 200 °C.

Component

Typical Temperature Ranges (°C)

Operational

Survival

Batteries

0 to 15

-10 to 25

Power Box Baseplates

-10 to 50

-20 to 60

Reaction Wheels

-10 to 40

-20 to 50

Gyros/IMUs

0 to 40

-10 to 50

Star Trackers

0to30

-10 to 40

C&DH Box Baseplates

-20 to 60

-40 to 75

Hydrazine Tanks and Lines

15 to 40

5 to 50

Antenna Gimbals

-40 to 80

-50 to 90

Antennas

-100 to 100

-120 to 120

Solar Panels

-150 to 110

-200 to 130

Thermal control techniques are broadly divided into two categories. Passive thermal control makes use of materials, coatings, or surface finishes (such as blankets or second surface mirrors) to maintain temperature limits. Active thermal control, which is generally more complex and expensive, maintains the temperature by some active means, such as heaters or thermo-electric coolers. In general, low-cost thermal

'This section has been rewritten in its entirety as of the 5th printing, September 2003. The help and assistance of GWynne Gurevich of Space Exploration Technologies; Brian D'Souza of Microcosm, Inc.; Ted Swanson, Ted Michalek, George Daelemans, and Dan Buder of Goddard Space Flight Center in the preparation of the new section is greatly appreciated.

control systems are designed to keep spacecraft at the cool end of allowable temperature ranges. Cooler components generally last longer, and this allows for system power growth. Though this can require additional power, it decreases the number of expensive iterations on the thermal design and analysis (which happens anyway, of course).

Thermal control is critical to ensuring the performance and survival of spacecraft and payload equipment Consider your personal computer, for example. It typically operates at room temperature plus or minus a few 10's of degrees. The space environment can cause equipment to get as hot as 100 °C and as cold as -130 °C with the changes occurring in 10's of seconds or minutes. Your cell phone works poorly, if it works at all, after being kept in your black car during a hot summer afternoon. In this example, the environment reaches approximately 60° to 65 °C.

Table 11-44 summarizes the design process for the thermal control system. As always, we begin with the development of requirements and constraints, paying particular attention to specific equipment or events likely to cause problems, such as the need for maintaining cryogenic temperatures for a payload instrument or a long thruster firing that may cause significant radiant heating on nearby surfaces. Step 2 is to determine the overall thermal environment of the spacecraft, i.e., characterize the heat inputs throughout the entire life of the mission. The most important external heat source will nearly always be the Sun, which continuously provides 1367 W/m2 (called the solar constant) at the mean distance of the Earth from the Sun and which varies as 1/r2 with distance from the Sun. (See Fig. 11-14.) This input goes away whenever the spacecraft enters a period of eclipse as discussed in Sec. 5.1. However, the Earth or other nearby central body serves as a moderating thermal influence by radiating heat in the infrared, corresponding to the blackbody temperature of the central body. (See Sec. 9.3.1.)

Fig. 11-14. Satellite Thermal Environment The most significant external heat source is the Sun, but we must also include reflected solar energy (albedo) and Earth infrared In our calculations. The only way a spacecraft can get rid of heat is by radiating it to space.

Fig. 11-14. Satellite Thermal Environment The most significant external heat source is the Sun, but we must also include reflected solar energy (albedo) and Earth infrared In our calculations. The only way a spacecraft can get rid of heat is by radiating it to space.

In Step 3, we review thermal requirements and constraints, compare them with actual heat sources and equipment placement and identify situations where the maximum and minimum equilibrium temperatures of the equipment are outside the required limits. For example, challenges arise when we must deal with cryogenic

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