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OTS PRICE
https://ntrs.nasa.gov/search.jsp?R=19630010070 2020-05-03T01:02:43+00:00Z
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Technica/ Report No. 32-424
Mariner Venus Power-Supply System
E. N. Costogue
/ I
G.E. Sweetnam, Chief
Spacecraft Secondary Power Section
JET PROPULSION LABORATORY
CALIFORNIA INSTITUTE OF TECHNOLOGY
PASADENA, CALIFORNIA
March 30, 1963
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Copyright © 1963
Jet Propulsion Laboratory
California Institute of Technology
Prepared Under Contract No. NAS 7-100
National Aeronautics & Space Administration
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JPL TECHNICAL REPORTNO. 32-424
CONTENTS
I. Introduction ..................... 1
A. Power-System Configuration .............. 1
B. Electrical Conversion ................. 3
1. Power Switch and Logic (PS&L) Module ......... 4
2. Booster-Regulator Module .............. 4
3. Synchronizer Module ................ 4
4. 2400-cps Power Amplifier .............. 4
5. 400-cps 3-phase-l-phase Power Amplifier ......... 4
6. Battery Charger .................. 4
C. Solar Panels .................... 4
D. Battery ...................... 5
E. Design Prob|ems and Restraints ............. 6
II. Operational Characteristics of the Power System ...... 7
Iil. Power-System Performance Before Launch ......... 8
IV. Power-System Performance from Launch to Encounter .... 8
V. Conclusion ..................... 21
III
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JPL TECHNICAL REPORT NO. 32-424
FIGURES
1. Power-supply block diagram ................ 2
2. Load profile ..................... 3
3. Solar-panel electrical connection .............. 5
4. Mariner solar-panel power and voltage ............ 5
5. Mariner solar-panel temperatures .............. 5
6. Solar-panel power ................... 7
7. Composite power ................... 7
8. Electrical connections of power sources ............ 7
9. Combined panel characteristics .............. 9
10. Panel voltage, average reading of 4A1 1 and 4A12
(Mariner 2) ...................... 10
11. Panelcurrent4A11 and4A12DSD8(Mariner2) ......... 11
12. Baffery voltage A2 (Mariner 2) ............... 12
13. Battery charge and drain characteristics ............ 13
14. Temperature of panel 4A12 front E8 (Mariner 2) ......... 14
15. Temperature of panel 4A11 front E7 (Mariner 2) ......... 15
16. Temperature of panel 4A1 1 back E9 (Mariner 2i ......... 16
17. Temperature of battery 4A14 E5 (Mariner 2) .......... 17
18. Temperature of booster regulator E1 (Mariner 2i ......... 18
19. Temperature of case V F4 (Mariner 2) ............. 19
V
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JPL TECHNICAL REPORT NO. 32-424
ABSTRACT
This Report covers the design, development, and integration of
the components and subsystems comprising the power supply for
Mariner Venus spacecraft. Data on the performance of the power
system from launch until the end of mission are also presented in
order to confirm the design.
I. INTRODUCTION
The objective of the Mariner Venus project was to
develop and launch two spacecraft to the near vicinity
of the planet Venus in 1962, to receive communications
from the spacecraft while in the vicinity of Venus, and
to perform scientific measurements of the planet and
interplanetary space. To support this project, a power
system was designed to provide usable power to the
spacecraft in the form of a 2400-cps square wave and
400-cps 3- and 1-phase sine waves. The power in flightwas to be derived from solar cells arranged upon erect-
able panels. The attitude of the panels with respect to
the Sun is dependent upon the attitude of the entire
vehicle. A battery source was to be available to support
the power requirements of the spacecraft from launch to
Sun acquisition and midcourse maneuver when solar
energy was not available because the spacecraft was
oriented away from the Sun.
A. Power System Configuration
In the Mariner Venus spacecraft power system, a
cleaner interface between user and power source was
attempted by distributing the power at a fixed ac voltageinstead of the various dc voltage levels required by the
users. The configuration for the Mariner power system
has the form shown in Fig. 1.
The system provided the following voltages: (1) 28-v
dc battery vo|tage which furnished power directly to the
various short-term high-power loads, such as valves, re-
lays, stepping motors, and squibs. (2) a 2400-cps square
wave for operating transformer-rectifier circuits to pro-
duce the dc voltage desired by the user. (3) a 400-cps
3-phase sine wave which supplied power to the attitude
control gyros. Single-phase antenna servos and scientific
motors were supplied from one of the 3-phase outputs.
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JPL TECHNICAL REPORT NO. 32-424
EE
0
0
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2
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JPL TECHNICAL REPORT NO. 32-424
With this arrangement the power supply development
proceeded on a realistic schedule because of reduced
interface problems with the users.
The power profile in Fig. 2 shows the raw power re-
quired from the 2400- and 400-cps systems, the battery
load requirements, and the solar power available from
the solar panels between Earth and Venus.
The input power during cruise mode was obtained
from two solar panels with a minimum raw-power capa-
bility of approximately 200 w in Earth space and, at
Venus, approximately 270 or 170 w, depending on the
amount of degradation of solar cells.
From the solar characteristics, it appeared that the
battery was expected to share the load with the solar
panels during periods 5, 7, and 12. The battery was ex-
pected to be completely charged during period 4 and to
have sufficient capacity to handle the periods of load
sharing.
The weight allocated for the power system was 105 |b:
24 ]b for electrical conversion equipment, 48 lb for solar
panels, and 33 lb for battery sources.
B. Electrical Conversion
The design and development of the electrical con-
verters and associated subassemblies were performed at
the Jet Propulsion Laboratory. The electrical conversion
equipment was designed to operate in an environment
of -10 to +55°C ambient; it was packaged into sixmodules of a combined volume of 727 in 2. A reliability
study on the design of the conversion equipment was
560
320
280 --
240 --
200
PERIOD
_1--i--- 2 _3"1-_4 _ 5 ----I'_ 6 ----I-- 7 --I.U-IO-+- I I --I-12-1--- 13
-- D 462-w PEAK, 2sec
F MA×IMUM SOLAR POWER
_H_MINIMUM SOLAR A
160
EIJJ
0CL 120
4O
SOLAR POWERREQUIRED
2w 0.5 w-
ILSUN ACQUIRED
S =_= DAYS
A = 20w, 3 PHASE, STARTING POWER, GYROS
B = 15w, SINGLE PHASE, I0 sec EVERY DAY, ANTENNA SERVO
C = lOw, SINGLE PHASE, 3 sec EVERY 16 hr, UPDATE SERVOD = B+C
E = 187w SINGLE PHASE, 2rain EVERY3Omin FOR 2doys
F = 5-.w PEAK, I0 sec/day
G = IO-w PEAK, 20 msec
BATTERY POWER
CONTINUOUS BATTERY POWER
MID-COURSE MANEUVER
G G GF F F F F
DAYS
Fig. 2. Load profile
3
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JPL TECHNICAL REPORTNO. 32-424
performed, and, from the derating factors used and the
environment of operation, a mean life expectancy of
3180 hr was calculated, based on available data. A brief
description of each subassembly follows.
1. Power Switch and Logic (PS&L) Module
This module received electrical power from the solar
source, the battery, and the ground power unit, selected
the power source, and provided the power to the booster
regulator, the battery charger, and the synchronizer.
Ground power was introduced into the power system
through a motor-driven switch that was controlled from
the ground. Another function of the module was to
provide telemetry information for the following:
1. Solar panel voltage No. 1
2. Solar panel voltage No. 2
3. Solar panel current No. 1
4. Solar panel current No. 2
5. Battery current drain
6. Battery voltage
7. Battery charging current
2. Booster-Regulator Module
This module received the variable power voltage from
the PS&L module, and provided regulated voltage -+-1%
to the 2400-cps power amplifier, 400-cps 3- and 1-phase
power amplifier, and the synchronizer. The pulsed-width-
modulation technique in the booster arrangement was
used for efficiency.
3, Synchronizer Module
This module received a 38.4 kc ±0.01% square-wave
signal from the control timer of the Central Computer
and Sequencer (CC&S) subsystem and, with the use of
countdown and magnetic oscillator circuits, provided
square-wave base drive to the 2400-cps amplifier, and
3-phase square-wave base drive to the 400-cps 3- and
I-phase power amplifier. The frequency of the drive wascontrolled to +0.01%. The synchronizer will free run at
2.2 kc in the absence of the control-timer signal.
4. 2400-eps Power Amplifier
This module received the frequency-controlled base
drive from the synchronizer and regulated-power voltagefrom the booster regulator in order to provide 50-v ___2%
rms square-wave voltage to the various transformer-
rectifier users. Tqae square wave output has an output
impedance of 1 ohm and a maximum capacity of 125 w.
The users of the 2400-cps square-wave output are asfollows:
Communications
Attitude control
Central computer & sequencerCommand decoder
Data encoder
Science
30w (max.)
29 w (max.)
7.0 w (max.)
4.7w (max.)
10.0 w (max.)
22.5 w (max.)
5. 400-cps 3-phase-l-phase Power Amplifier
This module received the frequency-controlled base
drive from the synchronizer and regulated-power voltage
from the booster regulator, in order to provide a 3-phase
26-v ±5% rms sine wave. By switching out two power
amplifier stages the module converted to a single-phase
power amplifier capable of providing 26-v ±15% rmssine wave.
The 3- and 1-phase sine-wave users are as follows:
3,/, gyros 13.5 w (max.)
14 antenna servos 1.5 w (max.)
l_b update servos 1.5 w ( max. )
l_b scientific load 4.1 w ( max. )
6. Battery Charger
This module provided a 1-amp maximum-current
constant-voltage charge to the battery. This charge was
reduced to a low rate and final]y to a trickle charge of
less than 5 ma as the battery voltage rose to the open
circuit voltage.
C. Solar Panels
The Mariner Venus spacecraft solar panels were de-
signed with the benefit of only a small amount of experi-
ence with the Ranger panels, plus additional laboratory
experimentation on the filter cement, the cell adhesive,
the insulation material, and the method of cell mounting.
The panels are approximately 60 in. long by 30 in. wide.The substrate was manufactured from aluminum for
light weight, and Mylar was used for insulation before
cells were positioned. The cells were applied to the panelwith an adhesive.
The panel circuit called for a total of 4896 cells with
102 in series and 48 strings in parallel. The cells are 1- by
2-cm areas using P-layer contacts at the corners along
the 1-cm edge with grids running the long way on the
ceil. The average electrical output of each cell was to be
22.9 mw at the maximum power point under tungsten
light of 100 mw/cm 2 and 28°C cell temperature. To
assure the power-output capability of the panel, selection
of cells was necessary for uniform efficiency. The elec-trical connections of the series and parallel strings are
shown in Fig. 3.
4
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JPL TECHNICAL REPORT NO. 32-424
Each panel has two parallel connected areas. Each ofthese areas has three sections which are connected in
series. This configuration was to be capable of producing
100 w per panel in Earth space, and 135 w per panel near
Venus, with a maximum power voltage ranging from49 v at Earth to no less than 35 v at Venus. Zener diodes
were installed on the panels to reduce the output voltage
of the solar source to 50 v. A power degradation during
the flights to Venus was anticipated because of the higher
stabilized temperature and by high-energy radiation.
CONNECTOR
-_ +POWER
__ -POWER
Fig. 3. Solar-panel electrical connection
Calculations of the above solar power capability were
based on solar-panel temperature of 35°C in Earth space.
However, Ranger 3 flight data indicated that the solar
panels operated at approximately 60°C, which was 15°C
higher than expected. Consequently, the power capa-
bility calculated at this higher temperature revealed apotential shortage of power. This power deficiency was
corrected by increasing the area of the solar array to
permit the addition of 918 solar cells, increasing the totalnumber of cells from 9792 to 10710. The increase in the
number of solar cells enlarged the solar panel capability
by 15 w at the maximum power point and 12 w at the
battery sharing point.
Curves of anticipated panel temperatures, maximum
power and maximum-power point voltage vs. Sun-probe
distance in millions of km are shown in Fig. 4 and 5,
respectively.
D. Battery
In the design of the battery, the basic requirements
were that this component must:
1. Provide power during launch
520
280 //- •(MAX.) POWER WITH ADDITIONAL ..- /
PANEL, i e, 102 SERIES /" ./ -I _/"
x 9 PARALLEL_ t./I"/_Z/"240 - /, "L'_"
200 .f-- / // (MAX) POWER --
oc .f ADDITIONAL PANEL.::t
"-J ./.--OPEN CIRCUIT VOLTAGE
t20 - _ 50
/-(MAX.) POWER
80 - _ (" VOLTAGE
40 ....
o EARTHI160
I VENUS-_
4O>
w
30°>
150 140 t30 120 I I0
SUN-PROBE DISTANCE, millions of km
20I00
Fig. 4. Mariner solar-panel power and voltage
290
270
250
%"_ 230E
>.-I.-
Z 210
I-z
fig
mo 190
17c
150
13C
160
SOLAR INTENSITY--_
S//
-_i/ .,,i-EARTH
/
I///
IIIII
x_/ ...... 130
9O
w
w
FRONT PANEL-- 70TEMPERATURE
z
BACK PANELTEMPERATURE
50
VENUS-P
150 140 130 120 I10
SUN-PROBE DISTANCE, millions of km
Fig. 5. Mariner solar-panel temperatures
30I00
5
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JPL TECHNICAL REPORT NO. 32-424
2. Recharge during first cruise mode
3. Support mideourse maneuver power requirements
4. Support two possible reacquisition modes before
encounter
5. Support a continuous load of less than 0.5 w
during the cruise mode after midcourse, and
6. Fully recharge before encounter
The Mariner Venus spacecraft battery was constructed
using sealed silver-zinc cells. It had approximately 40
amp-hr capacity and was recharged during the flight.Each unit was composed of an aluminum housing, two
cell blocks (a total of 9 cells per block ), two multiple-pin
connectors, and two large terminal posts. The total bat-
tery with the 18 cells weighed 33.3 lb and the energy-
to-weight ratio was approximately 40 w-hr/Ib.
The battery characteristics were as follows:
System Rechargeable, sealed, silver zinc
Voltage range 25.8 to 33.4 v
Current 15 amp at 32 v for 10 ms
Cycles 5
Temperature range 50 to 120°F
E. Design Problems and Restraints
In the usual spacecraft program, users of power begin
the development of their equipment at the same time the
power subsystem is developed. Thus, it is apparent that
the power system cannot be developed properly until
the load profile and voltage requirements are firmly
known. This is true if the power source produces the
variety of voltages required by the user. By applying afixed ac voltage to the user, the principal variable in
power-supply design is load uncertainty, and the voltage
problem is transferred across the interface to the user,
who must develop his own transformer rectifier system.
Some problems are created with this particular power
distribution method. One of the problems noted was in
the incompatibility between the peak power capability
of the 2.4 kc amplifier and the high instantaneous-current
demand of the capacitor input filters in several of the
user T-R units. This problem was solved by incorporating
choke input filters in those T-R units which were the
major users of power.
A second problem was the lower-than-expected eff-
ciencies exhibited by the user T-R units and regulators.
In establishing the design criteria for the Mariner power
system, a somewhat different and more logical approach
was used for allocating the power output to the users.
The input power to the T-R units, rather than the output
power, was used as the controlled quantity, which re-sulted in more careful evaluation of the T-R unit char-
acteristics by the user, particularly efficiency.
The requirements for the power-system design were to
provide the power for the spacecraft loads and to complywith the desired attributes in the order of the priority
given below:
1. Reliability
2. Effciency
3. Weight
4. Conducted and radiated interference in relation to
the interference susceptibility of the most critical
subsystem in the spacecraft
5. Balancing of heat dissipation
6. Size
A restraint in maximum charging current was neces-
sary to assure sufficient solar power for charging and
providing the spacecraft loads at all times.
6
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JPL TECHNICAL REPORT NO. 32-424
II. OPERATIONAL CHARACTERISTICS OF THE POWER SYSTEM
The operation of the power system was sensitive tothe orientation of the spacecraft and to the electrical
loads. During the launch phase, the battery supplied the
power, since the spacecraft solar panels were not orientedtoward the Sun. When the spacecraft became Sun-
oriented, the solar panels assumed the load and recharged
the battery. For the remainder of the mission, the modes
of operation are obvious; that is, when the spacecraft was
Sun-oriented, the panels were the prime source of elec-
trical energy; at other times, when the panels were notdirected toward the Sun, the battery was the primary
power source.
When the load exceeded the maximum power capa-
bility of the solar panels, or when the output from the
solar panels was decreased because of spacecraft orienta-tion, the solar panels and the battery shared the load. In
either case, sustained operation was expected to deplete
the store of energy in the battery and to result eventually
in spacecraft failure. It had been anticipated that nor-
mally the sharing mode would be ended either by Sun
reacquisition or reduction of load.
The mechanism of the sharing mode was of consider-able interest. For most load conditions, there were two
operating points for the solar-panel system: (1) high
voltage and low current, and (2) low voltage and high
current. Solar-panel characteristics were conveniently de-
scribed in terms of power vs. voltage, as shown in Fig. 6.
Analysis of the power system revealed that, for stable
operation, the slope of the power-voltage characteristicmust be negative. Clearly, then, the region to the right
of the maximum-power point on the solar panel charac-
teristic was stable, whereas the region to the left was
unstable. In the Mariner system, operation to the left of
the maximum-power point of the solar panel character-
istic was stabilized by setting the maximum-power point
above the battery voltage.
n,cffhdo
Io
_._,_-- -- -- -- PEAK
/ I \
" i IE, E2
VOLTAGE
Fig. 6. Solar-panel power
POWER
NEGATIVE SLOPE
POSITIVE SLOPE
The effect of combining battery and solar panel char-
acteristics is illustrated in Fig. 7, in which the composite
power is plotted as a function of voltage. The circuit
logic which resulted in this characteristic is shown in
Fig. 8.
hi
o
o
//
II
///
//
llT
J'\VOLTAGE, v
Fig. 7. Composite power
IIII h
..... l t] _"III
I I
k
SOLAR PANEL
rI SWITCHING LOGIC
I
" t-' III.L
I SOLAR PANELL t
]-
Fig. 8. Electrical connections of power sources
Note that the solar-panel characteristics were modified
by the shunt zener diode, which reduced the voltage
swing anticipated, but did not affect the stability of the
system. The composite characteristic is shown as a solid
line in Fig. 7, the solar-panel characteristic appears as a
dotted line, and stable regions of operation are labeled
I and III. Region II is unstable. The actual shape and
size of the solar-panel portion of the characteristic curve
were determined by the temperature of the panel, the
incident solar intensity, and the degradation sustained
because of radiation damage; hence, there was consider-
able uncertainty in the location of critical points on the
curve,
7
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JPL TECHNICAL REPORTNO. 32-424
For a given set of conditions (i.e., a fixed-composite
characteristic), it is enlightening to examine the opera-
tion. If the power required was represented by the level
A in Fig. 7, the operating point was at the intersectionof A and the characteristic curve and was seen to lie in
a stable region (III). If the load were increased from
A to C, the operating point would still have remained in
region III, However, were the load increased to level D,
the operating point was expected to jump into region I.
Between levels B and C, the voltage was multivalued;
however, the operating point was expected to be in stable
regions only. If the load was reduced from a value such
as D, the curve was followed down to point B, whence
the voltage jumped upon further reduction of load to a
stable point in region III. Note that, in region I, the bat-
tery and the solar panel were expected to share the load;
in region III, the solar panel carried the load alone.
In the sharing condition, the input voltage to the bat-
tery charger was the battery voltage; thus, no charging
could occur. Generally speaking, sharing was an unde-
sirable condition which, if continued, was expected to
lead to the depletion of the battery and failure of the
power system. The only way to escape the sharing mode
was to reduce the load (power demand) below point Bin Fig. 7.
In Mariner, the solar panel source was designed to
provide sufficient power in the battery-sharing voltage
so that the system would drop out of sharing when the
spacecraft load was reduced from the peak demand to
cruise mode operation. In addition, a capability of
switching off a load (science) was incorporated to assure
the drop-out from sharing if the panel characteristics
deteriorated beyond the predicted values.
I!1. POWER-SYSTEM PERFORMANCE BEFORE LAUNCH
All hardware of the power system was flight-acceptance
tested prior to delivery to the spacecraft assembly area.
One set of the hardware was type-approval tested. The
type-approval tests subjected the equipment to extensive
environmental testing to prove the electrical and me-
chanical design of the system.
The power system was installed in the spacecraft and
was subjected to system tests at the spacecraft assembly
facility. System, flight-acceptance, and type-approval test-
ing presented some problems requiring rework of flighthardware.
During evahmtion tests of the Mariner spacecraft, the
400-cps, 3-phase supply appeared to develop abnormal
voltages when the unit was switched from single-phase
to 3-phase operation. The problem was traced to high
current arcing through a relay contact, and was corrected
by replacing the relay with a new one which had the
required current capacity.
A second problem occurred in the booster regulator of
Mariner 2 when the module showed abnormal operation
during a high-load condition. The problem was traced
to the unmatching characteristics between portions of
the booster-regulator circuitry which, for packaging con-
venience, were housed in two modules. The problem was
corrected by requiring that the flight-acceptance-tested
modules be kept in pairs for proper operation.
Data on the leakage current of the solar-panel cells
revealed that the series-connected blocking diodes in the
power switch and logic modules could be removed. The
elimination of these diodes reduced the power require-
ment by approximately 5 w. This added to the safety
margin of solar-panel power operation during cruisecondition.
All of the above modifications to the power system
were effected before final system tests were performed.
IV. POWER-SYSTEM PERFORMANCE FROM LAUNCH TO ENCOUNTER
Mariner 2 spacecraft was successfully launched at
approximately 0653 GMT on August 27, 1962. The power-
supply system performed normally during all the phases
of the spacecraft flight, launch, cruise, midcourse ma-
neuver and encounter of Venus. Some problems were
encountered during the flight; however, they did not
hinder the operation of the power-supply system, because
of redundancy. These problems are mentioned below:
8
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JPL TECHNICAL REPORT NO. 32-424
1. October 31, 1962 (day 304): a short developed in
one of the solar panels causing the spacecraft volt-
age to drop approximately 8.4 v. Because of the
short, the power output of that panel dropped to
zero and the other panel was supporting the space-
craft loads. The direct parallel connection of the
panels (no diode isolation) caused a damping con-
dition to the solar source voltage near the maximum
power point of the power producing panel. This can
be seen in the combined characteristics of the power
producing and shorted panels shown in Fig. 9.
2. November 8, 1962 (day 312) : the short on the panel
recovered and all voltage returned to normal.
. November 14, 1962 (day 318) : the short on the same
panel reappeared, and remained until the end ofmission.
, The battery temperature exceeded the upper design-
temperature limit of 120°F. However, no failure
was seen during encounter.
. On December 30, 1962 (day 364) : the 2.4 kc power
supply frequency shifted to 2.2 kc. This frequency
was the free-running frequency of the power system
which indicated the loss of the synchronizing signal.
Curves of the power-system telemetry information
(solar-panel output voltage and current, battery voltage,
current and charge, and system temperature measure-
ments) are shown in Figs.10through 19. Figure 10 shows
the average readings o£ both panel voltages during the
flight. The estimated values are also shown. The abrupt
voltage drops were caused by the short developed in one
of the panels. Figure 11 shows the output current of each
panel. When the short developed in one of the panels,
the power output produced by that panel dropped to
zero, causing it to stop providing solar power to the
spacecraft, and the spacecraft began receiving current
z"hi
0-
28O
OM NAL PANELS
I ÷ 4AI2
240
160 , __
120
80 ........
0O 10 20 30 40 50 60
VOLTAGE, v
Fig. 9. Combined panel characteristics
from the other panel, as shown by the values marked.
The power-producing panel was clamped at about its
maximum power point. Figure 12 shows the battery
voltage during the flight. The voltage dropped after
80 days of flight when the battery charger became in-
operative because of the drop in solar source voltage.
Figure 13 shows the battery drain and charge current.
A1] activity in the battery current data stopped after the
battery was charged. Figures 14, 15, and 16 show the
temperature of the panels. The front temperature of both
panels and the back temperature of one panel were
measured. Figures 17, 18, and 19 show the temperature
of the battery, booster regulator, and power-conversion
equipment.
9
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JPL TECHNICAL REPORT NO. 32-424
50! .....
46 mm--
44
42>
LJ(._
C) 40 ...._>
.JbJZ
o- 38--
36
34
32--
0 POINTS OF ESTIMATED
PANEL VOLTAGE FROM
ANALYSIS PROGRAM DATA
ENCOUNTER -_
0
239
10
249
20 30 40 50 60 70 80 90 t00 ..... t10
FLIGHT, days
259 289 279 289 299 309 319 329 339 349
Fig. 10. Ponel voltoge, overoge reoding of 4All ond 4A12 (Nloriner 2)
120 I 130 140
365
359 4 14
10
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JPL TECHNICAL REPORT NO. 32-424
bJn-r,-
C)
/bJ
n
0
0
239
10
249
20
259
30 40
269 279
Fig. 1 i.
50 60 80 90 I00
FLIGH_ doys
289 299 309 319 329 339
Panel current 4A1 ! and 4A12 D5 D8 fMoriner 2J
I10 i20 I 130 140
365
349 359 4 14
11
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JPL TECHNICAL REPORTNO. 32-424
36
35
34
33
>32 r
<tI.-
>o 31_ -
)-hiLUp-
nl
"l \
29
28
0
259
10
249
20
259
3O
269
40 50 60 70 80
279 289 299 309 319
FLIGHT, days
Fig. 12. Battery voltage A2 (Mariner 2)
90
329
I00
339
It0 120 130
349 359 365 4
12
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IPL TECHNICAL REPORT NO, 32-424
7.0 1.0
x BATTERY DRAIN 0-7 omps D6
• BATTERY CHARGE 0-1 amp BB
(MARINER 2)
6.0 08
zbdn,-n-
5.0 0.6
iI
I
2.0 0.4 •
J8
,.o0.2 L illl_._',
i
0 O_ _ ..........
0 5
259
I0 20
249 259
30
269
BATTERY
.... ___== ...........
40 50 60
279 289 299
FLIGHT, days
DRAIN AND BATTERY CHARGE
70 80 90 I00 II0 120 130
309 319 329 339 349 359 365 4
Fig. 13. Battery charge and drain characteristics
13
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JPL TECHNICAL REPORT NO, 32-424
82.2 180
71.1 160 --
48.9 IZO0
239
I0
249
20
259
30 40 50 6O 70 80 9 0
FLIGHT, doys
269 279 289 299 309 319 329
0 POINTS OF ESTIMATED TEMPERATURE FROMANALYSIS PROGRAM DATA
/o)
E TIMATED
100
339
ENCOUNTER
I I0 12o
565
349 359
130
4
Fig. 14. Temperature of panel 4A12 front E8 (Mariner 2i
14u
14
14
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JPL TECHNICAL REPORTNO. 32-424
320
300
280
126.7 260
=u 115.lb. 240o
n- (2:
_,o4.,__o
IM WI..- _-
93.3 200
82,2 I 80
71.1 160
60 140
48.9 120
I
....
I
I_) 20
239 249 259
I I
PANEL SHORT--------_I=,._
_ tJn
t
t
PANEL SHORT
PANEL RECOVERED_ _ _
/
/
0 POINTS OF ESTIMATED TEMPERATURE FROM ----
ANALYSIS PROGRAM DATA
' __ --r t ..............
...=.
30 4o 5o eo 70 80 90FLIGHT, doys
269 279 289 299 309 319 329 339
Fig. ] 5. Temperoture of ponel 4A| 1 front E7 (Morlner 2)
_ (MAX.) D/N
ENcouNTER----
100 lib 120 1 130 140
365
349 359 4 14
15
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JPL TECHNICAL REPORTNO. 32-424
220
104 210 ....
95 200 ----
82 t80--
er_ 7t b_ 160o
g gk-
_n" 60 _ '40
W wP" F-
49 120
58 IOC
27 80 .........
r .............
I..,-Z
0uzbJ
16 60 ----
4 40
0
239
I0 20 30 40
249 259 269 279
50 60 70 80 90
FLIGHT, doys
289 299 309 319 329
100 ii0 120 130 140
565
339 349 359 4 14
Fig. 16. Temperature of panel 4A1 1 back E9 (Mariner 2J
16
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JPL TECHNICAL REPORTNO. 32-424
150
140
130 --
120
I10
Ld"
,,_ I00
I-- 9O
80
70
--f
/
60--
500 l0 20 :30
259 249 259 269
40 50 60 70 80 90
FLIGHT, days279 289 299 309 319 329
ENCOUNTER
11100 II0 120 I 130 140
365
339 349 359 4 14
Fig. 17 Temperature of battery 4A14 E5 (Mariner 2J
17
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JPL TECHNICAL REPORTNO. 32-424
150 -- --
120
I_. I10
g
d
<_ I00n...IJJQ..
ILlI'--
90
80
70
60
50
0
239
I0 20 30
249 259 269
L1
40 50 60 70 80 90 I00
FLIGHT, doys
279 289 299 309 319 329 339
Fig. 18. Temperature of booster regulator E1 (Mariner 2)
J
ENCOUNTER
rio
349
120 l 13.0 140
565
:359 4 14
18
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JPL TECHNICAL REPORT NO. 32-424
160
150
140
130 --
120
o
t_nr
I'-,,_ II0
ILl0..
hlI--"
lOC
9O
8O
/
70--
60
0
239
10
249
20
259
30
269
40 50 60 70 80 90 I00
FLIGHT, doys
279 289 299 309 319 329 339
Fig. 19. Temperature of case V F4 [Mariner 2)
I
ENCOUNTER
I10 120
349 359
I 1:30 140
565
4 14
19
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-i -T _7
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JPL TECHNICAL REPORT NO. 32-424
V. CONCLUSION
The power system provided all energy required to
operate the spacecraft and experiments. This has not
been, however, without some abnormal behavior. Factor-
ing out the abnormal experience, the performance has
corresponded to that predicted. Solar power output be-
fore the short developed was as predicted without any
degradation. Because of the high solar power available,
no sharing during the expected periods occurred. Battery
drain and charge were normal. Trickle charge of less
than 5 ma remained continuously on the battery for more
than 2 mo without destruction. Correspondence of the
predicted and telemetered values appeared to be within
and, considering telemetry accuracy of 3%, the designseems to be well confirmed.
21
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