Space Station in orbit. These two factors involve maintenance and replacement of major components and addition of large structures like the solar array wings. All of these activities will cause large variations in the array and battery outputs, resulting in substantial differences between the daytime and nighttime bus power capabilities. The communication satellites operate either in spin-stabilized or three-axis stabilized orientation. Because of launch vehicle limitations, the higher power satellites (above one kW) have resorted to three-axis stabilization using sun-oriented flat foldout solar arrays. Relative to the power system, typical operational modes on current comsats (3] appear to be the following: (1) during the sun seasons (76% of the year) the batteries are continuously trickle charged and then reconditioned just prior to entering the eclipse season, and (2) during the eclipse seasons, the batteries and loads are monitored so that the batteries are not discharged greater than the design limits (normally about 75% for NiCd). Load Demand and Profile Spacecraft electrical loads are conveniently divided into three groups: payloads, crew systems (for manned vehicles) and housekeeping. Payloads are generally the most important equipment, and they include experiments, special sensors, and orbit construction equipment. The crew systems loads consist of life support systems, food processors, waste processors, and control and display. For the Space Station, many of these crew systems loads that affect the life and safety of astronauts automatically become the highest priority loads and cannot be turned off. The housekeeping loads are identified by the spacecraft disciplines: attitude control, communication, data management, propulsion, thermal control, and electrical power. All loads also may be classified by input voltage type (ac or de and voltage levels) to facilitate power distribution and control. Collectively, the housekeeping loads represent a relatively small fraction of the total spacecraft loads on the Space Station and telecommunication satellites. In the DANMOE concept, the most important issue is whether or not there would be sufficient user loads that are controllable in the manner described earlier, i.e., reduction of nighttime loads and/or load shifting from eclipse to sunlight period. For the Space Station about 73% (55 kW) of the initial station load of 75 kW bus power is for the payloads, and 27% for the housekeeping and crew systems [2]. For the final stage, the bulk of the 150-kW bus load is estimated to be in the payloads and other loads staying at about 30 kW. This range of ratios between the payloads and other user equipment suggests that DANMOE is very feasible under normal operation. Good candidates for daytime-only (or mostly daytime) loads include equipment and habitation heaters, food and waste processors, interior illumination, refrigerators, freezers, various types of experiments, and construction equipment. Heaters, in particular, may have to be managed by the centralized controller to facilitate control of their various operational modes. Even some of the large life support equipment loads that are non-essential from a crew safety standpoint could be placed under DANMOE control. The SKYLAB, flown into LEO in 1973, offers the only source of power management data for an extended mission powered by the photovoltaic system. During its liftoff, one of the two solar panel wings (over 5 kW each) had torn off, creating a 2-kW shortage at the bus. The mission controllers were able to quickly implement load management procedures that involved removal of non-essential loads to perform earth resources experiments and careful monitoring of the condition of 26 batteries [4]. Key lessons learned from the SKYLAB experience are: (1) electrical consumables management is a vital part of mission planning and must be carefully addressed starting at the early design phase rather than later in the program, (2) well-designed power distribution and controls to facilitate electrical load profile management at the main bus is essential, (3) mission operations requirements involving equipment sequencing and operation were largely ignored during the design phase; they should be identified early enough to be implemented in both power and thermal system designs, and (4) battery load management was possible via appropriate sequencing of selected user loads. Power System Performance Characteristics The performance capability of any solar array/battery system operating in orbit is usually defined as the average power available at the main bus during one orbit. This parameter is identified as Pq in Eq. (3) for the case where P„ = P . The available bus power is limited by the available array power, Pg^, and the predetetermined battery depth of discharge (DOD) limit. In addition to array sizing, Eq. (3) is thus also very effective in arriving at the power capability of the system for a given orbit. In actual practice, however, the daytime load is generally different from the nighttime load so that Eqs. (5) and (6) are needed to find P . In such cases, either P or P must be defined or known in order to calculate P^. The bus power capability will vary as a function of the orbit as illustrated in Fig. 10 for the LEO spacecraft. This available bus power, calculated using Eq. (3) and the eclipse profile shown in Fig. 8, can increase by more than twice the end of mission bus power whenever the eclipse duration goes to zero. This situation occurs in only about 5% of the orbits within a year (see Table 4) so it cannot be relied upon on a continual basis. Nevertheless, on the LEO spacecraft a substantial amount of bus power is available every 50 to 100 days for several consecutive
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