Apollo Program
The first Apollo mission launch date was October 27, 1961 and it verified the Saturn I's aerodynamic and structural design. A series of five more launches were conducted before the first manned Apollo launch was made on October 11, 1968. See Table 4.1 for details of Apollo missions [1]. On July 16, 1969, man landed on the moon (Apollo 11 mission) followed by five more moon landings. The last launch was on December 7, 1972. See Table I for details of Apollo missions.
Electrical Power System of Apollo
The Apollo command and service module electrical power system (CSM-EPS)
was designed to operate from any combination of seven direct-current sources
[2-5]. The electrical sources are:
. Fuel Cells (3, 575 kilowatt-hours each)
. Entry Batteries (3, 40 ampere-hours each, silver oxide-zinc)
Service Module Battery (1, 400 ampere-hours, added after
Apollo 13)
The three fuel cells located in the service module (SM) supplied the primary source of power; two of the three entry batteries, located in the command CM, supplemented the fuel cells during high electrical demand; and the service module battery could be used if a fuel cell failed. Because of the failure of the cryogenic oxygen system in Apollo 13, a 400 ampere-hour service module battery was installed in the remaining Apollo missions. If required, this battery could have provided 12 kilowatt hours of additional or emergency energy via the command module main buses.
After the spacecraft attained orbit, the entry batteries were disconnected, recharged, and used to supplement the fuel cells during service propulsion system (SFS) burns. The load varied from 60 to 80 amperes between SPS bums, which is well within the fuel cell rating; however, during SPS burns when the gimbals were operated, the load current could reach a level of 120 amperes, which required the additional capacity of the two entry batteries.
The EPS of the Apollo CSM was designed to deliver nominal 28 volts dc and three-phase 400 hertz 115 volts ac derived from one of three inverters each having sufficient capacity to supply all alternating-current power required by the system.
The basic dc distribution system as shown in Figure 4-1 has two redundant buses and a single point ground that is connected to the spacecraft structure [61. The two main dc buses, marked A and B, are energized by the fuel cells and/or the entry and post landing batteries labeled A, B and C. Battery buses A and B are powered by their respective entry and post landing battery. The third battery C can be connected to either or both buses in the event that batteries A or B fail.
The flight and post landing bus was energized from both main dc buses and diodes or directly by the three entry and post landing batteries via diode pairs.
The flight bus received power from both main buses A and B through isolation diodes and the nonessential bus (marked 1 and 2) was energized from either main bus A or B depending on the position of the mechnically coupled single-pole double-throw switch.
The pyrotechnic buses A and B, which were isolated from the main electrical via a normally open switch, are powered by the pyrotechnic batteries. If the pyrotechnic batteries malfunctioned, entry batteries could be connected to pyrotechnic bus A or B.
The battery charger was a constant-voltage current-limited charger with the current limited to 2.8 amperes for a battery voltage less than 36 volts.
The charger operated in a continuous mode. At 36 volts the battery charger entered a cycling mode. The internal impedance of the battery increased with increasing battery voltage causing the charging current to decrease. At 39
volts minimum, the current was negligible and the battery reached its fully charged state.
The ac power distribution system illustrated in Figure 4-2 was a three phase four-wire system with the ac neutral connected to the single point ground. Two ac redundant buses, 1 and 2 provided power to the ac spacecraft loads.
Ac power was supplied by one or two solid state inverters rated at 115/200 volts 400 hertz. They produced 1250 volt-amperes each. Inverter 1 and 2 were respectively powered through main bus A and B and inverter 3 through either main A or B. The AC control (6 motor switches) operated contacts to connect or disconnect the inverter from the ac buses such that no two inverters were connected to the same ac bus at the same time. Inverters were automatically disconnected if an overvoltage or overload were present.
The inverter was designed to meet the following specifications [2]:
. Frequency: 400 hertz with 6400 hertz external timing or 400 + 7 hertz when free running
A major portion of the ac generated was used to power the fuel cell pump motors which presented a highly inductive load to the inverters. A capacitor bank was added to compensate for the lagging power factor of the inductive loads. When the fuel cell pump motors were redesigned with a larger power factor (less inductive), it was demonstrated that the power factor correction bank was overcompensating. Some of the capacitance was removed instead of redesigning the box.
The Lunar Excursion Module (LEM) electrical power system [7] supplied all required power for the LEM during its lunar mission. Prior to separation from the orbiting portion of the Apollo spacecraft, power was provided by the
CSM-LEM docking umbilical cable. The electrical power system of the LEM
consisted of a dc and ac section with the primary dc power being supplied by six silver oxide-zinc batteries (four in the descent stage and two in the ascent stage).
During the descent phase, all four batteries, rated at 400 ampere-hours each at a nominal output of 28 volts, supplied the electrical power in order for the LEM to complete its mission exclusive of the ascent phase. If only three descent batteries were functional, a protracted mission could be executed. However, if two of the four descent batteries were nonoperational, the LEM mission would be aborted. Either ascent battery, rated at 300
ampere-hours, was capable of supplying all ascent electrical power demands during normal mission operations and an abort.
The ac electrical power was generated by two solid-state inverters with each inverter rated at 115 ± 2 volts rms with an input of 28 ± 4 volts dc. A 6400 hertz master timing pulse supplied by the LEM guidance computer set the output frequency of the inverter at 400 ± 4 hertz. In the absence of the timing pulse, the output frequency tolerance increased to ± 10 hertz. The inverter output waveform was sinusoidal with less than 5% total harmonic distortion.
The dc and ac systems are shown in Figures 4-3 and 4-4.
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