ots price - nasa€¦ · cleaner interface between user and power source was attempted by...

OTS PRICE

https://ntrs.nasa.gov/search.jsp?R=19630010070 2020-05-03T01:02:43+00:00Z

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

Copyright © 1963

Jet Propulsion Laboratory

California Institute of Technology

Prepared Under Contract No. NAS 7-100

National Aeronautics & Space Administration

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

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


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.


EE

0

0

@

0L

2


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


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


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


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


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


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


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


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


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


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

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

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


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


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


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


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


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

-i -T _7


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

ots price - nasa€¦ · cleaner interface between user and power source was attempted by...

Documents