K700k

Insufficient Data

Hardness

- Natural Radiation -Nuclear Threat

- Laser Threat

- Pellets

Low-Medium Medium Medium Low

High High High Medium

Very high Very high Very high Very high

Very high Very high Very high Very high

High High High Medium

Stability and Maneuverability

Low

Medium

High

High

High

Low-orbit Drag

High

High

Low

Medium (due to radiator)

Low

Degradation Over Life

Medium

Medium

Low

Low

Low

Storage Required for Solar Eclipse

Yes

Yes

No

No

No

Sensitivity to Sun Angle

Medium

High

None

None

Spacecraft

Shadowing

Low (with bypass diodes)

High

None

None

Spacecraft

Viewing

High

High

Low

Medium (due to radiator)

None

Fuel Availability

Unlimited

Unlimited

Very low

Very low

Medium

Safety Analysis Reporting

Minimal

Minimal

Routine

Extensive

Routine

IR Signature

Low

Medium

Medium

High

Medium

Principal Applications

Earth-orbiting spacecraft

Interplanetary, Earth-orbiting spacecraft

Interplanetary

Interplanetary

Interplanetary

employing a boiler, turbine, alternator, condenser, and pump. This power-conversion cycle is essentially the same as that used to generate electricity from fossil and nuclear energy on Earth. Power-conversion efficiencies for Rankine-cycle engines are 15-20%. Brayton-cycle engines are dynamic devices that use a single, compressible working fluid as the working medium. The thermodynamic cycle consists of adiabatic compression and expansion stages separated and coupled by stages that add or reject heat at constant pressure. Placed after the turbine, a recuperator-heat exchanger improves the cycle's efficiency. Power conversion efficiencies for the Bray ton cycle are 20-35%.

Fuel cells convert the chemical energy of an oxidation reaction to electricity. They are self-contained generators that operate continuously without sunlight, but must carry their own reactant supply, usually. The longer the mission, the larger the reactant tanks. The most popular version for space applications is the hydrogen-oxygen (referred to as "alkaline" because of the KOH electrolyte) fuel cell because of its relatively high specific power (275 W/kg on the Space Shutde), low reactant mass (hydrogen and oxygen), and useful by-product (water).

A typical single cell produces a voltage of 0.8 Vdc. In combination, a fuel cell unit can create many kilowatts of power (each Shuttle fuel cell produces 16 kW peak or 12 kW continuous). The energy conversion efficiency can run as high as 80% for low current draws, but as current increases, the efficiency drops to 50-60%, due to activation overpotential and electrical resistance in the electrolyte solution between electrodes. However, compared with other power sources, fuel cell efficiencies are high.

The three Space Shuttle fuel cells are state-of-the-art power generators that produce all of the Shuttle electricity for the 28 Vdc bus. Their high efficiency (70%), low weight (118 kg), and excellent reliability (> 99% available) attest to their quality. Other important factors are their 15-min start-up time, instantaneous shutdown, and long lifetime (2,400 hours before refurbishment). Besides electricity, these fuel cells produce crew drinking water, at a rate of 0.36 kg/kWh, or about 104 kg a day.*

Research is underway to solve the short-mission limit with fuel cells, caused by carrying large reactant masses. Because the fuel-cell reaction is reversible, we can use electrolysis to create more reactants from the water by-product To optimize each process, however, we have to use separate units for generating electricity and separating the water. Any long-duration mission could use this regenerative system if it had some input electricity from solar cells, nuclear generators, or other power system during periods of low electrical load.

Earth-orbiting spacecraft at low-Earth to geosynchronous orbits have usually employed photovoltaics as their power source. Often, photovoltaics were the only real candidate for these low-power missions (less than 15 kW) because solar cells were well-known and reliable. Photovoltaic sources are not attractive for interplanetary missions to the outer planets because solar radiation decreases, thus reducing the available energy from a solar array. To configure and size a solar array, we must understand cell types and characteristics; solar-array design issues, types, sizing calculations, configurations, regulation; and radiation and thermal environments. Key design issues for solar arrays include spacecraft configuration, required power level (peak and average), operating temperatures, shadowing, radiation environment illumination or orientation, mission life, mass and area, cost, and risk. Table 11-34 shows the solar array design process.

Step 1. Mission life and the average power requirement are the two key design considerations in sizing the solar array for most spacecraft We size a photovoltaic system to meet power requirements at EOL, with the resulting solar array often oversized for power requirements at BOL. This excess power at BOL requires coordinated systems engineering to avoid thermal problems. The longer the mission life, the larger the difference between power requirements at EOL and BOL. We usually consider photovoltaics a poor power source for missions lasting more than 10 years because of natural degradation in the solar array. Section 11.4.4 discusses how we manage excess power from the solar array. The average power requirement can be obtained from Sees. 10.1 and 10.2.

* Telephone conversation with Jay Garrows, International Fuel Cells, Inc., Oct. 98.

TABLE 11-34. Solar Array Design Process. In the FlreSat example column, ld represents inherent degradation, 6 is the Sun incidence angle, Ldis life degradation, and X and Xd represent the efficiencies of the power distribution paths. The material following the table further explains these quantities.

TABLE 11-34. Solar Array Design Process. In the FlreSat example column, ld represents inherent degradation, 6 is the Sun incidence angle, Ldis life degradation, and X and Xd represent the efficiencies of the power distribution paths. The material following the table further explains these quantities.

Step

Reference

FlreSat Example

1. Determine requirements and constraints for power subsystem solar array design

• Average power required during daylight and eclipse

• Orbit altitude and eclipse duration

• Design lifetime

Input parameter, Sees. 10.1,10.2

Input parameter, end papers

Chaps. 2, 3

110 W during daylight and eclipse

700 km 35.3 min

5 yr

2. Calculate amount of power that must be produced by the solar arrays, P^

Step 1

Eq. 11-5

Assume a peak power tracking regulation scheme with Xe = 0.6 and Xd = 0.8 Psa = 239.4 W

3. Select type of solar cell and estimate power output, P0. with the Sun normal to the surface oi the cells

P0 = 0.22X1,367 W/m2 = 301 W/m2

Si solar cells Pa = 202 W/m2

4. Determine the beglnning-of-life (BOL) power production capability, Pbol • per unit area of the array

Table 11-35 Eq. 5-7 Eq. 11-6

0 = 23.5 deg (worst case) Pbol = 143 W/m2

5. Determine the end-of-life (EOL) power production capability, Peol . 'or the solar array

Performance degradation Si: 3.75% per yr, GaAs: 2.75% per yr, Multijunction: 0.5% per yr Eq. 11-7 Eq. 11-8

Performance degradation is 3.75% per year

¡-4= 0.826 for 5 yr mission PE0L =118.1 W/m2

6. Estimate the solar array area, a^, required to produce the necessary power, Pgg, based on Peol an alternate approach

Eq.11-9 Eq. 10-12t

Ajg = 2.0 m2 Ajg = 2.5 m2

7. Estimate the mass of the solar array

Eq. 10-13t

ma = 9.6 kg

8. Document assumptions

0 0

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