History of Orbiting Solar Observatory OSO-2

8/7/2019 History of Orbiting Solar Observatory OSO-2

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HISTORY OFORBITING SOLAR OBSERVATORYoso-2

GPO PRICE $CFSTI PRICE(S) $ -i_-.

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I..ITHRUI&67 (ACCESSION NUMBER)11368, 0 I -l a - !* ICODEI

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4 CCATEOORYI(NASA C ~ R ~ RMX O R AD NUMB E R) I.

1 - GODDARD_ SPACE FLIGHT CENTERc

_-__ GREENBELT, #ID.,\


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1X-440- 66-322

HISTORY OFORBITING SOLAR OBSERVATORYoso-2

April 1966

GODDARD SPACE FLIGHT CENTERGreenbelt, Maryland


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CONTENTS

SECTION 1 - INTRODUCTIONPage-

1.1 INTRODUCTION 1-11 .2 GENERAL DESCRIPTION O F O S 0 1-2

1 . 2 . 1 SPACECRAFT 1-21 . 2 . 2 EXPERIMENTS . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . .. . . . . . . . . . . . .. . . . . . . . . . . . . . . . . . .1-14. . . . . . . . . . . . . . . . . . 1-14

1.3.1 OSO-2 MISSION OBJECTIVES 1-141-15

1 . 3 MISSION OBJECTIVES

. . . . . . . . . . .1.3.2 OSO-2 EXPERIMENTS . . . . . . . . . . . . . . .SECTION 2 - OSO-B DISASTER2 . 1 EVENTS LEADING U P TO THE DISASTER 2-1. . . . . . . .2 . 2 THE DISASTER 2- 1. . . . . . . . . . . . . . . . . . . .2 . 3 ACCIDENT INVESTIGATION . . . . . . . . . . . . . . . 2-2

. . . . . . . . .. 4 CONCLUSIONS AND RECOMMENDATIONS 2-5. . . . . . . . . . .. 5 SPIN BALANCE FACILITY REWORK 2-5SECTION 3 - DEVELOPMENT OF OSO-23 . 1 SPACECRAFT . . . . . . . . . . . . . . . . . . . . . . 3- 1

3 . 1 . 1 WHEEL STRUCTURE . . . . . . . . . . . . . . . 3- 13 . 1 . 2 SAILSTRUCTURE 3-23 . 1 . 3 BEARINGS . . . . . . . . . . . . . . . . . . . . 3-23 . 1 . 4 ATTITUDE CONTROL SYSTEM 3-23 . 1 . 5 COMMAND SYSTEM . . . . . . . . . . . . . . . . 3-33 . 1 . 6 TELEMETRY SYSTEM . . . . . . . . . . . . . . 3- 33 . 1 . 7 ELECTRICAL POWER SYSTEM . . . . . . . . . . 3-4

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CONTENTS vPage

SECTION 3 - DEVELOPMENT OF OSO-2 (Continued)3.2 EXPERIMENTS. 3-5

3.2.1 POINTEDEXPERIMENTS . 3- 63.2.2 WHEELEXPERIMENTS - 3- 6

3.3 OBSERVATORY . . . . . . . . . . . . . . . . . . . . . 3-83.3.1 WEIGHT ANDBALANCE . . . . . . . . . . . . . . 3-83.3.2 ACCEPTANCETESTS 0 0 3-8

3.4 SHIPMENTTOCAPEKENNEDY . 3- 103.5 PRE-LAUNCH PREPARATIONS 0 . . 3- 12

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3.5.1 HANGARCHECKS . . . e . . 3-123.5.2 SPIN' AND BALANCE CHECKS . . . . . . . . . . . 3- 153.5.3 LAUNCH TOWER CHECKS . . . . . . . . . . . . . 3-16

SECTION 4 - LAUNCH AND EARLY ORBIT OPERATIONS4.1 LIFT-OFFANDFIRSTORBIT . . . . . . . . . . . . . . . 4- 14.2 EXPERIMENTTURN-ON . . . . . . . . . . . . . . . . . 4-3SECTION 5 - NORMAL OPERATION5.1 SPACECRAFT . . . .

5.1.1 ATTITUDE CONTROL SYSTEM . . . . . . . . . . .5.1.2 TELEMETRY SYSTEM . . . . . . . . . . . . . . .5.1.3 COMMAND SYSTEM . . . . . . . . . . . . . . . .5.1.4 ELECTRICALPOWERSYSTEM . . . .5.1.5 THERMAL MONITORING SYSTEM . . . . . . . . .

5.2 EXPERIMENTS . . . . . . . . . . . . . . . . . . . . . .5.2.1 NAVAL RESEARCH LABORATORY ULTRAVIOLET

TELESCOPE AND CORONAGRAPH . . . . . . . . .

5-15-15-95-105-125-125-16

5-16

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CONTENTS

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Page-SECTION 5 - NORMAL OPERATION (Continued)

5.2.2 NAVAL RESEARCH LABORATORY X-RAY TELESCOPE 5-265.2.3 HARVARD COLLEGE OBSERVATORY ULTRAVIOLET. . . . . . . . . . . . . .PECTROHELIOGRAPH 5-275.2.4 GODDARD SPACE FLIGHT CENTER ULTRAVIOLET

SPECTROPHOTOMETER . . . . . . . . . . . . . .GODDARD SPACE FLIGHT CENTER LOW ENERGYGAMMA-RAYTELESCOPE . . . . . . . . . . . . ..2.5

5.2.6 AMES RESEARCH CENTER EMISSIVITYDETECTORS . . . . . . . . . . . . . . . . . . .UNIVERSITY OF MINNESOTA ZODIACAL LIGHTTELESCOPES . . . . . . . . . . . . . . . . . . ..2.7

5.2.8 UNIVERSITY O F NEW MEXICO HIGH ENERGYGAMMA-RAY TELESCOPE . . . . . . . . . . . .

5.3 TRACKING AND DATA . . . . . . . . . . . . . . . . . .5.3.1 DATA ACQUISITION . . . . . . . . . . . . . . . .5.3.2 DATA PROCESSING . . . . . . . . . . . . . . . .

SECTION 6 - TERMINAL OPERATION

5-27

5-30

5- 305-31

5- 325-325-325-34

6 . 1 G E N E R A L . . . . . . . . . . . . . . . . . . . . . . . . 6-16.2 TERMINAL OPERATION PLAN . . . . . . . . . . . . . . 6-16.3 TERMINAL OPERATION AND RESULTS . . . . . . . . . . 6-2

6.3.1 FIRST DAY - SEPTEMBER 24, 1965 6-26.3.2 SECOND DAY - SEPTEMBER 25, 1965 6- 36.3.3 THIRD DAY - SEPTEMBER 26, 1965 6-36.3.4 FOURTH DAY - SEPTEMBER 27, 1965 6-46.3.5 SIXTH DAY - SEPTEMBER 29, 1965 6-46.3.6 SEVENTH DAY - SEPTEMBER 30, 1965 6-5

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CONTENTSPage

SECTION 6 - TERMINAL OPERATION (Continued)6.3.7 EIGHTH DAY - OCTOBER 1, 1965 . . . . . . . . . . 6-56.3.8 NINTH DAY - OCTOBER 2 , 1965 . . . . . . . . . . 6-66.3.9 FOURTEENTH DAY - OCTOBER 7, 1965 . . . . . . . 6-6

6.4 CONCLUSIONS . . . . . . . . . . . . . . . . . . . . . . 6-8SECTION 7 - STOW OPERATION7.1 PURPOSE . . . . . . . . . . . . . . . . . . . . . . . . 7-17.2 OPERATION . . . . . . . . . . . . . . . . . . . . . . 7-17.3 OPERATIONS AFTER OBSERVATORY STOW . . . . . . . . 7-17.4 RENEWED OSO-2 OPERATION, JUNE 1966 . . . . . . . . . 7-6

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LIST O F ILLUSTRATIONS

..

FigureFroniispiece

1-11-21-31-41-51-61-71-81-91-101-113-13-23- 3

4- 15- 15-25-35-45-55-65-75-85-9

nnn n n......... CLU D U - L O ~ ~ L c ; G : C ; L Q L L . . . . . . . . . . . . . . . .Major Features of Wheel Structure . . . . . . . . .Major Features of Sail Structu re . . . . . . . . . .Azimuth Shaft Assembly . . . . . . . . . . . . .Launch Sequence Events . . . . . . . . . . . . .Gas Control Systems . . . . . . . . . . . . . . .Control System . . . . . . . . . . . . . . . . .Raster Scan . . . . . . . . . . . . . . . . . . .Nutation Damper . . . . . . . . . . . . . . . . .Oriented Experiments . . . . . . . . . . . . . .Oriented Experiment Operational Modes . . . . . .Wheel Experiments . . . . . . . . . . . . . . . .OS0 Observatory in Ball Bro thers Balance Machine .OSO-2 Vibration Test Set-up . . . . . . . . . . .OSO-2 Loading at Lowry A i r Force Base .22 November 1964 . . . . . . . . . . . . . . . .Thor. elta/OSO- 2 Lift-offOSO-2 Pitch Angle Drift . . . . . . . . . . . . .. . . . . . . . . . . .OSO-2 Spin Rate Decay . . . . . . . . . . . . . .OSO-2 Spin Rate . . . . . . . . . . . . . . . . .Pointing Control Readout . . . . . . . . . . . . .Long Term Battery Charge Rate and Voltage

Battery Charge Rate and Voltage . rbit 3637OSO-2 Hub Temperature . . . . . . . . . . . . .OSO-2 Rim Skin Temperature . . . . . . . . . . .

. . . .Battery Charge Rate and Voltage . rbit 2435 . . . .. . . .

Page...1

1-31-41-51-61-71-81-101-111-171-181-213-93-11

3-134-25-25-55-65-85-135-145-155-175-18

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LIST OF ILLUSTRATIONS (Continued)

Figure5-105-11

5-125-135-145-155-165-17

OSO-2 Bottom Skin Temperature . rbit 56Azimuth Power T ran sis tor Tem perature -Orbit 2611 . . . . . . . . . . . . . . . . . . .Solar Cell Panel Temperature . rbit 3769 . . . .Solar Plot From OSO-2 Data During Orbit 194 . . .Sun Photo Taken from Earth on 16 February 1965. .Analog-to-Digital Processin g Line . . . . . . . .Quality Control and Edit Program . . . . . . . .

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Quick-Look System, Fort Myers to GSFC . . . . .LIST O F TABLES

Table5- 15-27-17-27-37-47-5

Page5-19

5-205-215-225-235-355-375-39

Successful "Playback-ON" Attempts . . . . . . . 5-11Analog Tape Track Assignments . . . . . . . . . 5-33OSO-2 Temperatures - Orbit 4120 . . . . . . . . 7-3Transmitter RF O N Attempts . . . . . . . . . . 7-5Spacecraft Battery Voltage . . . . . . . . . . . . 7-5OSO-2 Temperatures (Degrees Centigrade) -March 1966 . . . . . . . . . . . . . . . . . . 7-7

OSO-2 Status - Orbit 4120 . . . . . . . . . . . . 7-2

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w w n w neORBITING SOLAR OBSERVATORYoso-2I I I U IV I \ I VI

SECTION 1INTRODUCTION

1 .1 INTRODUCTIONThe sun is a seething sphe re of atomic particl es with temperatures and

forces so great that they a re almost incomprehensible to man. The surface ofthe sun is punctuated by sun spots , solar fla res and prominences and massiveemiss ions of radiation are associated with these so la r phenomena.

Solar radiation and events have a definite effect on communications and cli-mate on ear th and are the determining factor in sustaining human life. If solarradiation decreased by as much as 596, the ea rt h would become a frigid ball andlife, as we know it, would cease.

It is believed that the sun is responsible for magnetic-field and charged-parti cle phenomena beyond the ea rth 's atmosphere. Therefore , if man is toventure into the reaches of space to explore our s is te r planets and beyond, in-formation relative to the intensities of these magnetic fields and charged parti -cles must be obtained to protect him adequately.

Man has been observing the sun and solar phenomena and gathering dataabout them for hundreds of yea rs. However, his effor ts have been hindered bybeing earth-boyd and capable only of observing a small fraction of the totalsol ar spectrum due to the earth's atmosphere. Iman could take his instru-ments beyond the ea rth's atmosphere, he would be capable of studying the entireso la r spectrum. Great str ides toward this end have been made during recentyear s with the advent of balloon and rocket astronomy and man has increased hisknowledge about his nearest star. However, only tiny bit s and pieces of the so-la r puzzle could be fitted into place because of the limited equipment that couldbe ca rr ied aloft and the sh ort durations they remained above the atmosphere.

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With the development of NASA's Orbiting Solar Observatories, so la r scien-tis ts , for the first time in his tory, have been afforded the opportunity to observethe sun, unobstructed by the eart h's filter ing atmos phere , for extended periodsof time , Because of this, la rg er pieces of the sol ar puzzle are now fall ing intoplace and man is hopeful of being capable of accu rate long-range weather-pre-diction , development of more efficient communications s ys tems , and gainingmore information about magnetic-field and charged-particle phenomena.

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1 . 2 GENERAL DESCRIPTION OF O S 0

The OS0 spacecraft, designed by Ball Brothers Research Corporation undercontract to NASA, ar e earth-orbiting satellites used as stabilized platforms forsolar-oriented scientific instru ments .John F. Kennedy Space Center atop a modified Douglas Thor-Delta launch vehi-cle into a circular orbit 300 nautical miles above the ea rth . The inclination ofthe orbit to the equator is 33 degre es. The orbital period is approximately 95minutes.

They are presently launched fr om the

Observatory gro ss weight is approximately 600 pounds. Instrument weightcapable of being ca rr ied aboard the spacecraft is approximately 40% of the grossweight of the observatory.

The total expected useful life of the observatory w a s six months and waslimited only by the amount of high pres su re nitrogen gas capable of being ca rr iedaboard to maintain pointing control of the instruments . However, on futurespacecra ft, a magnetic bias coil has been provided to reduce the effect of theearth's magnetic field on the spacecraft's pitch and roll attitudes.great ly increase the overa ll operational life of the O S 0 spacecraft.

This should

1 . 2 . 1 SPACECRAFTThe main body of the spacecraf t cons is ts of a wheel of aluminum alloy sep-

ara ted into nine wedge-shaped compartments as illustrated in Figure 1-1. Fiveof these compartments contain apparatus for five experiments not requi ring so-la r orientation. The remaining four compartments house the electronic control s,batteries , telemetry equipment, and radio-command equipment. Each wheelcompartment provides approximately 1000 cubic inches of volume, and eachwheel experiment can weigh up to approximately 45 pounds. Two compar tmentscan be used by one experiment, and in this ca se , the maximum experimentweight is 90 pounds, The wheel also contains three extendable ar m s which a r e

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-WHEEL STRUCTURE

P THERMAL SHIELD

COMPARTMENTRIMPANEL

Figure 1 - 1 . Major Features of Wheel Structurefolded down around the th ird stage of the launch vehicle during the launch phase.When the spacecraft is injected into orbit, the arms are extended which in-cr ea se s the effective diameter of the wheel fro m 44 nches to 96 inches.

A fan-shaped sail is mounted on top of the wheel by means of a rotating alu-minum shaft and carries a sol ar a rray and pointed experiments.ar ra y supplies the electr ica l power required for daylight operation, and itchar ges storag e batte ries fo r night-time operation. Space is provided in thesail section fo r two pointed experiments, each 4 inches wide by 8 inches high.The length w a s limited to 36 inches by the curv atu re of the Delta shroud , andthe thickness of the eyes mounted on the pointed exper iments. The length cannow be expanded to 57 inches because of improvements to the Delta vehicle. Atotal instrument weight of approximately 88 pounds can be accommodated in thesail section.due to Delta improvements.sail structure.

The solar

The radius of the sail is 22 inches which can also be increasedFigure 1-2 illust rates the major fea tures of the

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Figure 1-2. Major Features of Sail StructureThe two main structures are connected by an aluminum shaft which runs

from the base of the sail, through the cen ter of the wheel and terminates in asupport ring stru cture on the underside of the wheel. This shaft is held in placeby two bearings, one on top and one on the bottom of the wheel. Mounted on theshaft between the two bearings is a high pres su re nitrogen gas tank for pitchprecession jets located on the sail structure. A torque motor, mounted on thetop of the shaft, con trols the azimuth position of the sail while the spacecraft isin daylight. On the base of the shaft is a slip-ring assembly which allows trans-mission of power, t elemetry sign als , and control signals between the sail sec-tion and the wheel. Figure 1-3 illustrates the azimuth shaft assembly.

The inherent stability of the spacecraft is due to the gyroscopic spinning ofthe ent ire wheel struct ure . Figure 1-4 illu strates the sequence of events duringlaunch. After second stage burnout, the enti re third stage and spacecraft com-bination is spun up to approximately 120rpm, activating an acceleration switchwhich st ar ts the Launch Sequencer Tim er . After third-stage burnout and 90seconds after the spacecra ft time r is sta rted, squibs ar e fired in the spacecraftto release the three a rms . Centrifugal force extends them. Extending the armsinc rea ses the gyroscopic stability of the wheel and decr eas es the spin rat e toapproximately 96rpm. A t the end of the a r m is a high pressure nitrogen gas

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Figure 1-4. Launch Sequence Eventstank which supplies nitrogen gas to de-spin jets o r spin-up jet s, also located onthe three arms. The wheel rotation is reduced by expelling nitrogen gas fromthe de-spin jets , thus slowing the wheel to 30 f 5% rpm. Photoelectric so la rsensors ( so l a r eyes) mounted on the r im of the wheel sense the sun once eachrevolution and electronic ci rcui try measures the period of rotation . When themeasured rotational period drops below 28.5 rpm, nitrogen gas is expelledfr om the spin-up gas jets and the spin rat e is increased to within 5% of nominal.Figure 1-5 illustrates the gas control systems.

The spacecraft is maintained in a position so that the spin axis is normal tothe so la r direction within * 3 degrees. This is accomplished by expelling nitro-gen gas from one of two precessio n jets mounted on the sa i l to gyroscopicallypre ces s the spacecraft about its pitch axis. A set of pitch control ey es , mountedon the sail, senses when the spin axis has drifted thr ee degrees f rom normal tothe so lar direction, and electronic circuitry activates the proper precession jetto prec es s the spacecr aft to one degree past normal in the opposite direction.Figures 1-5 and 1-6 llus tra te the locations of the pitch gas cont rol componentsand control system components, respectively.

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Figure 1-6. Control SystemThe OS0 spacecraft rolls about the roll axis at approximately one degreeper te rr es tr ia l day because of the precess ion of the observatory orbit and the

ea rth orbit. The Aspect Measuring System provides a method of determiningthree-axis aspect with respect to the celestial sphere fo r the OSO. Aspect canbe determined within three degrees. The measurement requires a magnetom-eter mounted on one ar m of the spacecraft, with its sensitive axis in the planeof rotation. The Aspect Measurement System provides two s et s of measureddata which a re read out directly by the OS0 telemetry system. These two read-outs are: (1) a measurement of the time inte rval which re la tes the wheel posi-tion in time to the spacecraft data word pulses for spin rat e and spin angle de-termination, and (2) a measurement of the time interval between the magneticfield position and the solar direction for spacecraft roll angle determination.The magnetometer produces a sinusoidal output as the wheel rotates, exceptwhen the spin axis is parallel to the earth's magnetic field. The magnetometeroutput reaches its maximum positive value when the angle between the earth'smagnetic field and the sensor axis is a minimum. Zer o output occurs when the

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.sensitive axis becomes Perpendicular to the magnetic field. When the wheel hasrotated 180 degrees and the sensor's sensitive axis is opposed to the field butfo rm s the minimum angle between the sensitive axis and the lines of force, themagnetometer produces its maximum negative value. The zero output of themagnetometer is distinguished with electronic signal conditioning; therefore ,the instantaneous time at which the magnetometer's axis is perpendicular to themagnetic field can be determined. This event is recorded from a pulse outputwhich may be sent both to an experiment and to an on-board logic circuit. Asecond pulse! to loca te the spin vector in the spin plane re lative to the magne-tometer sensitive axis, is generated from a sun senso r. The angle in the spinplane between the normal to the magnetic field and the sun vector is determined,for a known spin rate, by measuring the time interval between these two pulseswith a counter circuit which counts the spacecraft 400 cps clock pulse s. Theroll angle of the spin axis , with resp ect to the ec liptic plane , is calculated byusing this angle, the magnetic field characteristics and the earth-sun line atthat point in space and time.

Azimuth control of the sail and pointed experiments is accomplished by asys tem of co ar se and fine azimuth eyes and ser vo control. Each satell ite night,with no sun as a reference, the torque motor is stopped and the sail structurerotates synchronously with the wheel. During sate llite dawn, the azimuth coar seand fine eyes acquire the sun, and a servo control sys tem and the azimuth torquemotor driv e the sail in a direc tion equal and opposite to the spinning wheel.When the pointed instruments are such that they are within 5 degrees of the cen-te r of the sol ar disc, the coa rse eyes are switched out of the circuit, and onlyfine eyes a re used to move the sa il through the remaining 5 degre es. The finecontrol eyes then maintain the sail and pointed instruments to within f 1 arc-minute of the center of the so la r disc.

The elevation fine eyes develop e rr or signals fo r control of the elevationservo in a manner similar to the azimuth control system. The pointed instru-ments a re mounted on an elevation frame which is controlled by the elevationtorque motor. The elevation se rv o is capable of moving the elevation fram eand pointed instruments through an angle of * 5 degrees in a plane containing theso la r vector and the spin axis. It will maintain the pointed instruments in ele-vation to within i 1 arc-minute of the cente r of the sol ar disc.

The azimuth and elevation systems of the OS0 spacecra ft can also be oper-ated in a ras ter scan mode. Figure 1-7 illustrates the ra st er scan. This scanmode sweeps the pointed ins truments in azimuth and elevation in such a way thatthe entire solar disc and part of its corona can be mapped by the experimentalinstruments. The scan consist s of a square pattern 40 arc-minutes on a side,centered about the cen ter of the maximum intensity of the solar disc. The scan

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STARTOFRASTER\

TIME256 SECSI \ II 1 fI I\ I I

ELEVATION4 ZIMUTH

Figure 1-7. Raster Scanstarts at a point 20 arc-minutes in azimuth and elevation from the center andsweeps in azimuth. The azimuth sweep is accomplished in 6 . 4 seconds, and attheend of the sweep, the elevation step s down 1 arc-minute and the azimuthsweep is reversed. A s seen in F igure 1-7, forty elevation steps a re accom-plished, presenting a complete picture of the sun every 256 seconds.

When the spacecraft is separated from the third stage rocket, it has a nu-tation due to wobble of the unbalanced burned-out third stage, the unsymmetrical

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Telemetry is accomplished by a PCM/FM system containing two independentand parallel multiplexing systems. Each system has a multiplexer-encoder,digital tape re cor der and a tra nsm itt er. Experiment measurements and space-cr af t operating data (housekeeping data) a r e processed by the multiplexer-e ~ c n d e rnt,o eight-bit words and 32-word main frames on a time shared basis.The multiplexed exper iment and housekeeping data is simultaneously recordedby the tape reco rder and transmit ted to ground stations on a rea l time basis bythe transmi tter . Once during each orbi t the tape re corder playback commandis sent m d the tape record er is connected to the tr ansmi tter to transm it the datait recorded during the orb it. Playback of the orb ita l data is accomplished inapproximately 5 minutes, and at the end of the playback period, the tape re-cor der automatically re tu rn s to the recor d mode and the transmit ter again be-gins transmitting re al time data.

The command system is a PCM/AM/AM tone-digital type capable of execu-ting 7 0 commands. Two command rece ivers operate continuously fo r protectionagainst a single rec eiver failure. The rece iver outputs a r e a pulse-durationmodulated audio tone and a r e fed to decoders which in tur n actuate latch-typere lays fo r command execution. The OS0 command sys tem enables ground con-troll ers to: (1) control power (on or off) to experiments and various spacecraftsys tem components; (2) select various space craft components by signal or powercontrol; (3) operate the spacecraf t control system by switching from automaticto manual control or vice vers a; (4) control the real time and playback modes ofoperation; ( 5 ) control the pointing or ras ter modes of operation; (6) bypass theday-night control power system; and (7) bypass the automatic undervoltage andR F time r power cir cui ts. The sys tem uses digital tone techniques, and thecommands a r e transmitted from STADAN (Satellite Tracking and Data Acquisi-tion) stations. The spacecraf t receives the coded R F command signal and de-codes the digital information according to three possible ad dre sses. Two ad-dr es se s activa te two redundant-output decoders in the wheel and the thi rdactivates a decoder in the sai l. The command fr am e consist s of two repea tedadd ress words and three repeated o r different command words. An alternatecommand fra me of one addr ess word and one command word sent twice in suc-cession can also be used. The command execution is performed if one of thesecommand words activates the addressed decoder circu it. Commands are pre-programmed and automatically fed into the ground transmitting equipment bymeans of a punched paper tape. Emergency commands can be transmitted bymeans of a manual control panel in the ground equipment.

The spacecraft electr ical power system consists of the so lar a rr ay , themain batte ry pack, the squib batte ry pack, the undervoltage switch , and theturn-on circuitry. The solar array is mounted on the front face of the sailstr uct ure . There a re approximately 1872 silicon sol ar cells arranged in 36

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thrust of the separa tion spring and the unsym metr ical motion of the a rm s asthey ar e extended. A fur the r cause of nutation is unbalance of pointed instru-ments which have mechanical scanning devices. No nutation occurs when theinstruments are pointed, but when the pointing control is turned off at night andthe oriented section starts spinning, a dynamic unbalance occu rs that causes awobble of approximately 10 minutes. When the pointing control is turned on thefollowing day, the wobble shows up as a nutation. A nutation damper is providedto remove these undesirable motions. It works on the principle that, if energyis removed fro m a free ly rotating body, the body ro ta tes about the ax is of max-imum moment of inert ia. This axis is made to coincide with the azimuth shaftaxis by careful balance before launch. The Nutation Damper is shown in Figure1-8. It is mounted on the oriented section and cons ist s of a tuned pendulumwhich moves in a silicone oil bath. The nutation energy is transmitted to thependulum of the nutation damper and is dissipated by the oil. The damping con-stant is small so the bob moves through large amplitudes for small nutations.

s IL

rEXPANSION BELLOFRAM

Figure 1-8. N u ation Damper

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para llel strin gs of 52 each. The solar ar ra y converts solar energy into electri-ca l power to fu rnish power for daytime operation and to charge the main batter-ies fo r Eight operation. The design power available from the s ola r a rr ay isapproximately 27 watts. The main battery pack is made up of 42 nickle-cadmiumtype F cells. The 42 cells are distributed in six packs of seven ce ll s each. Toproduce an 18.9 volt dc supply, 14 cells a re connected in se ri es , and the thre epacks of 14 cell s each a re connected in para llel to produce the nec essa ry powercapacity. The pack voltage range is 16 volts to 22 volts fro m a nearly dis-charged state to a fully charged state. To isolate the squib firing circuit fromthe main spacecra ft power supply, firing energy for the various squibs is takenfro m separate batteries. The squib batte ries a r e nickle-cadmium type C cells ,and the re a re two battery packs in both the sa il and the wheel st ructures . Thesa il packs supply energy to fi re the azimuth, nutation damper and elevationsquibs. The wheel batte ry packs fire the a rm and compartment door squibs.

The spacecraft batte ries have a rat her flat disch arge curve until they be-come almost fully discharged. Near the fully discharged s tat e the voltage fallsabruptly and the impedance rises s o that very little power can be drawn fromthem. If the batteries were to become fully discharged, they would not be ableto deliver the cur rent nece ssary to allow the se rvo to catch and orient the sa iltoward the sun. This would cause the spacecra ft to become disabled. Sectionsof the spacec raf t can be turned off by the command sys tem to conserve power.I he command system is not operative and the batteries discharged to the pointwhere the voltage drops abruptly, an undervoltage switch is provided whichshuts down the entire spacecraft and permits the batteries to charge from therotating solar ar ray . When the batteries a re about 10 percent charged, theundervoltage switch turns on the spacecraft again.

During the dark portion of the orbit , ce rtain of the components in the space-cr af t do not need to operate. To conserve power, the pointing control, thepointed experiments and the spin-up sys tem a re turned off. A redundant pai r ofso la r senso rs located on the r im of the spacecraft wheel feed a circuit whichactuates a rel ay. This relay turns the power on when the spacecraf t comes intothe sun and off when it goes into the shadow. Thes e turn-on eyes a re simi la r tofine eyes and also serve as sen sor s for determining wheel spin rate.

Thermal control of the OS0 spacecraft is accomplished passively by use ofsur fac es with well controlled optical prope rties . All wheel compartments arethermally interconnected by the main frame and central casting, and heat gen-erated and absorbed by the wheel is well distributed. Experiments a re secureddirec tly to the compartment deck and receive heat from and supply heat to thespacecra ft struct ure . Instrument temp eratu res within the wheel vary fro m 7degrees C at night to 13 degrees C during the day. The pointed instruments

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receive radiated heat fro m the elevation and azimuth fr am es , the front and r e a rof the s olar ar ray, the wheel top skin, the e ar th, dire ct sunlight on the frontface , and internal heat sources. The external su rfa ces of the pointed instru-ments are highly polished to keep the heat inside during the observatory night.Temperatures in the pointed inst ruments range f ro m 0 degrees C to 10 degrees C .1 . 2 . 2 EXPERIMENTS

Experiments ca rri ed on board OS0 spacecraft vary with each spacecraftand are determined by the objectives of each part icula r mission. They are alldesigned, however, to observe the sun and solar phenomena and to measureelectromagnetic radiat ion produced by these phenomena in the ultraviolet, X-rayand gamma ray region of the spectrum which is absorbed or reflected by theear th's upper atmosphere. Experiments car ried aboard OS0 spacecraft alsodetect and measure celes tial phenomena from other a reas in space such a s c e-lestial ultraviolet energy ; proton-electron energy ; cel est ial neutron flux char-acteristics ; zodiacal light intensity and direction; and interplanetary dust par-ticle number, momentum and kinetic energy.

1 . 3 MISSION OBJECTIVESWith the OS0 eries of spacecraft, NASA has embarked upon a comprehen-

sive space and astronomy program to study so la r behavior affecting the s truc -ture and circulation of ea rth 's atmosphere, the physical nature and origin ofsolar-surface phenomena, and the relation of thi s so lar activity to te rr es tr ia levents. This is being accomplished by a unique combination of government andprivate solar research organizations over an eight year period from 1962through 1970.

Solar activity follows a cycle averaging 11 ye ar s. The next period of max-imum activity will occur in 1969 and 1970. Many solar physicists believe thatif a complete study is made of radiation from the quiet sun, the incidence oftransient solar phenomena should be predic table. The re su lt s of OSO-1,launched on 7 March 1962, give some support to this belief. When all the datahas been reduced from OSO-2, launched 3 February 1965, it is believed thatthey too will support th is theory. A total of eight OS0 spacecraft ar e presentlyauthorized and will be placed into earth orbit during the eight year period.1 . 3 . 1 OSO-2 MISSION OBJECTIVES

OSO-2 was the second in this series of eight spacecraft which NASA willlaunch for the purpose of observing and investigating sol ar phenomena andeffects. Its mission objectives were:

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a. To study electromagnetic radiations fr om the sun in the ultraviolet andX-ray regions of the spectrum.To map the so la r disc in ultraviolet light and X-ray emi ssions .T n may, the intensity of the white-light corona of the sun.To monitor bursts of so la r X-ray emiss ions.To determine the origin of the polarization of zodiacal light.To measure the direction of a r r i v a l and energies of primary cosmicgamma-ray radiation in the energy spectrum from 100Mev to 1Bev.To detect gamma ra ys fro m the sun and from other sour ces in space inthe energy spectrum fro m 0 . 1 to 0 . 7 MeV.To measu re ultraviolet radiation from nebular and stellar sources.

b.c ;d.e.f ,

g.

h.i. To measur e the emissivity stability of spacecraft temperature controlmaterials.

1.3.2 OSO-2 EXPERIMENTSThe OSO-2 mission objectives were accomplished by a complement of eight

scientific instruments divided into tw o categories - solar oriented and non-oriented instruments. The oriented instruments were mounted in the sa il sec-tion and required constant pointing in the solar direction. The non-orientedinstruments were mounted in the wheel section and only required a quick-lookat the sun once every two seconds or did not require a look at the sun at all.

Three of the eight inst ruments on board OSO-2 were provided by privateso la r re sear ch organizations and the remaining five were provided by variousgovernment r ese arc h facilities.

1 . 3 . 2 . 1 Pointed Experiments1.3.2.1.1 Naval Research Laboratory Ultraviolet Telescope and Corona-

graph-Dr. R. Tousey, Investigator.

This portion of the NRL experiment package consisted of two experiments:a coronagraph experiment to map the intensity of the white light corona of theartificially eclipsed sun, operated during the pointing mode, and an ultravioletscan experiment in which the sun was successively mapped in three ultravioletwave-lengths during the spacec raf t ra st er mode. (Lyman-Alpha, 1216 ang-stroms; 388 angst roms; and H e I I , 304 angstroms.)

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The coronagraph experiment used an occulting disc to artificially eclipsethe sun and a mechanical scanner to scan around the s ol ar corona in a spiralmotion. The coronal light was detected by an end-window type photomultiplier.

The ultraviolet s can experiment consisted of a dispe rsion grating followedby a s e t of th ree ultraviolet photomultipliers, each located at the focus of thethree ultraviolet wave lengths. The high voltages requ ired to operate the photo-multiplier s were obtained from sepa rat e high voltage power supplies, only oneof which provided voltage at a given time. A voltage-selector energized each ofthe photomultipliers in sequence and in coincidence with signals generated atthe end of each sol ar scan.

1 .3 .2 .1 .2 Naval Research Laboratory X-Ray Telescope-Dr. T . A. Chubb,Investigator.

This portion of the NRL experiment package was grouped into ras te r modeand point mode experiments. The purpose of the pointed experiments was tomonitor burs ts of so la r X-ray emission in th ree wavelength bands and to se ar chfor X-ray emission from prominences at high altitudes above the solar disc .The purpose of the ra st er experiment was to repetitively map X-ray sourc es onthe sun in tw o wavelength bands.

The pointed mode experiment utilized four Geiger counters: A 2 to 8 ang-str om burst detector, an 8 to 20 angstrom background detec tor, a 44 to 60 ang-st ro m burs t detector, a background detector and a prominence detec tor. Theburst de tectors looked at the sun and recorded so la r X-rays continuously duringthe point mode of operation. The background detector looked away from the sunand provided a basis fo r correcting the data for counts due to V a n Allen or otherparticle radiation. The so la r prominence detector looked at the region aroundthe sun. During this observation, the sun was a rtific ially eclipsed by an oc-culting disc which was extended approximately 24 inches in fron t of the X-raycounter.

Mapping by the raster experiment was accomplished by recording as a func-tion of position responses of two very tightly collimated Geiger counter X-raydetec tors. The pulse generated in these detectors was alternately switched tothe data storage s ystem during each sun ra st er , thereby permitting ra st er pat-terns at two wavelengths simultaneously.

1 .3 .2 .1 .3 Harvard College Observatory Ultraviolet Spectrometer andSpectroheliograph-Dr . L. Goldberg, Investigator.

This experiment consisted of a spectrome ter which scanned over the longwavelength region of the ultraviolet spec trum between 300 and 1400 angstroms.

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The spectrometer used a windowless photomultiplier with a tungsten photo cath-ode as a detector which has a z er o response to wavelengths longer than 1400angstroms and a relatively flat response t o sho rte r wavelengths.

---*..--- c ~ & Q I-canning---- the i~itrsviolet pectrum: a second type of observationalprocedure was available. Upon command from the ground, the spectrometermoved to any desi red wavelength and a raster-type motion of the pointed experi-ments caused a monochromatic image of the sun to be constructed. The spe ctr alrange was covered with an approxirr-ate resolg tion of one aogstrom. The spec-trom et er was designed to have an acceptable angle of about 1.8 minutes of arcin both the horizontal and verti cal directions.

1 . 3 . 2 . 2 Wheel ExDeriments1 . 3 . 2 . 2 . 1 Goddard Space Flight Center Ultraviolet Spectrophotometer-

D r . K . L. Hallam, Investigator.Ultraviolet radiation from nebular and ste ll ar sources were plotted in a

wide-sky coverage of this light sou rce in both the northern and southern hemi-spheres. The overa ll spec tral coverage was from 1300 to 2600 angstroms.This experiment was expected to extend the knowledge of st el la r atmospheres,interstellar atmospheres , interstellar gas , inters tella r dust, and was to assistand complement the concurrent rocket astronomy program s by pre limina ry butcomprehensive data about the brighter ultravio let sources. The spectrophotom-eter had a field of view of 1/2 degree by 1 degree and was measured to an ac-curacy of 1 minute of arc. Intensities of the brighte r sources were measuredto an accuracy of 1 percent. The dynamic range of the instrument was sixstellar magnitudes.

1 . 3 . 2 . 2 . 2 Goddard Space Flight Center Low Energy Gamma Ray Telescope-M r . K . J . Frost, Investigator.

This experiment was designed to detect gamma-rays from the sun and fro mother sources in space, and to analyze thei r energy spectrum from 0 . 1 to 0.7MeV. Of particular interest in this experiment was the ability to detect the0.501 Mev electron-position annihilation line and to study any possible temporalvariations.

1 . 3 . 2 . 2 . 3 Ames Research Center Emiss ivity Detectors-Mr. C. B. Neel,Investigator.

To support the Apollo project , it was necessary to determine the perform-ance of spacecra ft temperature-control coatings in the space environment.

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Coatings to be tes ted fo r emissivi ty stability were applied to di sc s and exposedto s olar energy and space environment during orbit. Measurements of thermal-radiation charac ter ist ics of the coatings versus tim e were thus obtained.

1.3.2.2.4 University of Minnesota Zodiacal Light Telescopes-Dr. E. P.Ney, Investigator.

The purpose of this exper iment was to determine the orig in of polarizedzodiacal light, a nebulous light seen in the w e s t aft er twilight and in the eastbefore dawn. Experiment apparatus ca rr ied in the rotating wheel section meas-ured the intensity of polarized zodiacal light along the spin axis. To do thi s,photomultipliers covered with shee ts of Polaroid were mounted on the top andbottom of the wheel sec tion. The photomultipliers had a ten deg ree field of viewand measured both visible and infrared light. Rotation of the Polaroid sheet withthe wheel section produced an alternating signal of 1/2 cps which was a directmeasure of the intensity of the polarized light.

1.3.2 .2.5 University of New Mexico High Energy Gamma Ray Telescope-Dr. C. P. Leavitt, Investigator.

This experiment measured the direction of ar riva l and energies of primarycosmic gamma-ray radiation in the energy range fro m approximately 100Mev to1Bev. The primary purpose of this experiment was to locate disc ret e sourcesof radia tion and to dete rmine its energy spectrum in the range mentioned above.Energy resolution was expected to be about 30 percent and directional accuracieswere measured to within A 10 degrees.

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SECTION 2OSO-B DISASTER

2 . 1 EVENTS LEADING UP TO THE DISASTERThe OSO-B pacecraft arrived at Cape Kennedy on 12 March 1964. Routine

checkout and preparation of the spacecra ft and experiments took place until9 April 1964, at which time the payload was covered with a polyethelene bag,purged with dry nitrogen and placed in it s shipping conta iner. The payload wasstored in hangar AE to await the arr ival of the Delta third stage rocket motor.

Because of its heavier weight, the O S 0 spacecraft uses a rocket motor witha thicker wall casing. When the motor ar rived, it w as given a receiving inspec-tion, and it was discovered that there w a s a defect in the rocket motor casing.This motor w a s rejected, and a new X-248 A-6 rocket motor w a s flown downfr om Wallops Island, Virginia on 9 April 1964. The igniter paddle was removedfr om the rejected rocket motor to be installed in the new motor. During theremoval of this paddle it was damaged and a new Delta paddle was built by theNaval Propellant P lan t. The new Delta igniter paddle was flown to Cape Kennedyon 11 April 1964 and installed in the third stage rocket motor. The rocketmotor was transported to the Spin Test Facility on 12 April 1964.

On 13 April the payload was removed f rom the shipping container, placedon a truck and, at approximately 0400 hours, it was moved to the Spin TestFacility.2 . 2 THE DISASTER

Between 0930 and 0939 hours EST on 14 April 1964, the third stage X-248A-6 solid propellant rocket motor inadvertently ignited and burned in the SpinTest Facility at Cape Kennedy. The rocket motor with the spacecraft attachedtor e loose from the alignment fixture in which it was mounted and shot to theceiling of the facil ity. When it hit the ceiling, the spacecraft w a s torn loosefro m the third stage motor and fell to the floor. The rocket motor continuedon to the corner of the building and burned until its fuel w as expended. Elevenmen were burned - three fatally and eight others suffered injuries ranging fromcri tical to minor. The three men who died were not killed immediately butdied as a resul t of their burns within a couple days to a couple weeks after theaccident occurred.

Eye witness interviews after the accident indicated that the Douglas person-nel had just completed their ordnance checks of the third stage/spacecraft

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combination. One of the Ball Bro ther s Re search employees stepped ove r to thespacecraft t o adjust the polyethelene shroud which was placed over the space-cra ft and third stage a s a dust protec tor and to purge them with nitrogen. A she touched the shroud a crackle was heard and the third stage ignited.

2 . 3 ACCIDENT INVESTIGATIONA Fact-Finding Committee w as appointed by the NASA Goddard Space

Flight Center Directo r. The committee was comprised of the following person-nel :

D. G. Mazur, Chairman NASA-GSFCW . D. Baxter, Lt. ColD r . B. Bartocha N P PR . H. Gablehouse Ball Brothers

AFMTC

E . E . Harton NASA Headquar tersE . H. Helton NASA-Wallops IslandL. T . Hogarth, Secre tary NASA-GSFCR . J . JohnsonJ. J. NielonL. R. PiasechiW. R. SchindlerR . SteinbergerL. Swain

Douglas Aircraft CompanyNASA-GLOBJP LNASA-GSFCABLNASA-LRC

The committee investigated the following items at the East ern Test Range(ETR) immediately afte r the accident to establ ish the circumstances surroundingthe accident:

a.b.c .

Hardware configuration at the time of the accident.Eye witness testimony and reports.Examination of the accident a rea including inspection and test ing ofsignificant items .The time sequence of events leading up to the accident involving therocket motor and the spacecraft.

d.

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e. Review of procedures.f . Determination of possible causes.g. Plans f or tes ts to verify the possible cause s.

The committee divided the investigation into four areas: (1)heat, (2)R Fsignal, (3) electricity and (4)mechanical shock and/or vibration. The commit-tee al so undertook the investigation of a simi la r accident which took place a fewmonths earlier at the GEaIloma Ordnance Eeijot zt Pr yo r, Oklahoma, io deter-mine if the two accidents were related. In the Oklahoma accident, the DouglasAi rcr aft Company was preparing a destruct tes t of the X-248 ocket motor inorder to test a new Delta third stage destruct system. No one was killed in thisaccident, but test equipment and a crane used f or moving the rocket motor wereconsiderably damaged. One person received minor injuries.

The first of the investigative courses of action taken was to investigateelec trost atic discharge. This investigation was to determine possible modes bywhich an electrost atic charge could have caused ignition of the X-248 A-6 motor.It w a s divided into five tasks which were directed toward developing a compre-hensive picu tre of the electros tatic charac ter ist ics of the spacecraft/motorconfiguration and the motor/igniter assembly. The first of these tasks was todetermine the elect rosta tic sensitivity of the X-248 squib. The second tas k wasto determine the electrostatic sensitivity of all bulk explosives in the X-248motor. Another task was to determine total and inter-element e lec tri ca l char -acte ristic s (resistan ce, capacitance , and charge storage) of the spacecraf t/motor and the motor/igniter under the application of both static and transientelec trost atic voltages. This task w a s also to establish the critical inter-elemental breakdown voltages and paths. Task number four was to determinethe e lectrosta tic potentials and energies that could have been pres ent under thecircumstances prevailing at the time . The final tas k was to t ry to duplicate theX-248 nadvertant ignition in both the Eastern Test Range and Oklahoma acci-dent configurations.

Electrostatic sensitivity tests of the Delta X-248 squib were conducted bythe Franklin Institute. These tes ts consisted of discharging, through the squib,incremental voltages of a 500 pico-farad capacitor through a 5000ohm resi stanceto simulate the capacity and resi stance prope rtie s of a human being.

Measurements of the electrostatic sensitivity of the X-248 bulk explosiveswere conducted by the Naval Propellant Plan t. It was determined from thesemeasurements that the X-248 bulk explosives could not have been a factor inthe accident.

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2 . 4 CONCLUSIONS A N D RECOMMENDATIONSThe committee concluded that the cause of the ignition of the th ird stage

rocket motor at Cape Kennedy was a n elec trosta tic discharge through the ignitersquib. Cornel1 Aeronautical Laboratory recommended the following changes tothe X-248 rocket motor: (1) use of a squib insensitive to electrosta tic energiesup to 25 kilovolts and 500 pico-farad, (2 ) use of a res ist ive plug between thesquib case and the bridgewire, (3) use of a Faraday cage covering all sensitiveparts of the squib assembly and (4) use of a conductive s pray on elect rosta ticsensitive portions of the paddle. Cornel1 Aeronautical Laboratory demonstra tedthat the l as t three of the forementioned changes a re adequate to insure againstaccidental squib initiation due to elect rosta tic discharges up to 100 kilovolts.In all subsequent Delta launches, an X-258 rocket motor was used instead of anX-248 because it used an igniter assembly le ss sensitive to electrostatic dis-cha rges, and the squib can be inse rted on the launch tower.

Precautionary measures were also suggested which would apply to any solidpropellant rocket motor: (1)to avoid the use of non-conductive ma te ri al s, espe-cially plastics, (2) to use squib arrangements which would permit installation aslate in the operation a s possible, (3) to check the conductivity of each ignite r-motor sys tem planned for usage to verify a low resis tive path between all con-ductive components, (4) to s tr ic tly adhere to proper grounding procedures when-ever a solid motor is to be handled with or without an igni ter insta lled. It wasalso recommended that procedures f o r grounding personnel, spacecraft, motorand associated systems and components should be carefully considered for futurerocket motor handling.

2 . 5 SPIN BALANCE FACILITY REWORKThe possibili ty of a new Spin Test Facil ity was investigated; however, a

new facil ity could not be made ready until November 1965. It was decided torework the damaged facility fo r use in the Delta program. During the reworkof the Spin Test Facility, the following additional safety features w e r e added tothe building:

a. The pit in the southwest cor ner was floored over with portable deck-ing which can be removed if the pit is required for future operations.A new personnel door was placed in the center of the north wall.Roll up doors on the ea st end of the building were replaced with two6 by 10 feet swing-type doors. The remainder of the original open-ing w as replaced with blast panels.

b.c .

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Panic hardware was improved on all personnel eg re ss doors. Blanklatch facings were installed on the door fra me s.The interior of the west wall of the building was covered with gyp-sum wallboard to provide sealing and to r et ar d fires.The protruding tra ck s on the exter ior of the west end of the buildingwere removed.An emergency audible warning syste m was installed.A "Cone of P rotect ion Lightning System" was installed around thefacility.The existing communication syst em was removed and replaced byan explosion proof intercom system.Conductive plastic mats were supplied for use in ar ea s where ord-nance is handled.A sprink ler sy stem was installed in the high pay are a.Placards denoting explosive mate rial s, cl as se s, personnel lim its,e t c .were installed.Personnel safety showers were installed at all personnel egres sdoors.A n additional closed circu it TV system was installed with cam era sin the high bay area and monitors in the office trailer and controlroom. Personnel can now witness opera tions without being physi-cally present in the bay. Procedure s were changed so that allspacecraft testing necessa ry in the pit area is performed remotely.The personnel tr ai le r located a t the west end of the building was re-moved to a more remote location.The guard shack for the a re a was removed to approximately 350 feetfrom the facility.The relative humidity inside the building was increased from 50% to60%.

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c

SECTION 3DEVELOPMENT OF OSO-2

After the OSO-B disaster OS0 Projec t Management decided to rebuild theOSO-B spacec raf t using prototype par ts , OSO-B spa re parts and new parts.The OSO-B pratotype and flight Spacecraft were re turned to Ball Brothers Re-se ar ch Corporation in Boulder , Colorado. The exper iments were removed andreturned to the ir respective investigators for examination and testing to dete r-mine the extent of damage. When the damage to the spacecraft was surveyed, aschedule w as se t up to rework the OSO-B prototype spacecra ft from existingflight s pa re components, updated prototype components, undamaged flight com-ponents and new procurement. A ll new procurement and OSO-B flight subas-semblies were given a complete acceptance test. OSO-B flight sp ar e units wereonly given thermal tests , whereas the prototype components were given func-tional and thermal tests. The go-ahead for OSO-2 was given on 17 June 1964.

The spacecraft was completely stripped of its components and the wheel andsail structu res were separated. The flight subassemblies that were removedfrom the OSO-B spacecraf t were inspected and tested to determine th eir usabil-ity. 35%of these subassemblies were determined to be suitable for use asOSO-2 flight subassemblies, 16%were suitable for use a s flight spares and theremaining 49%were considered surplus and not suitable for OSO-2 usage.

When the OSO-2 spacecraft w a s delivered to Cape Kennedy, it was composedof 19% OSO-B flight subassemblies , 29%OSO-B flight spares , 22%0SO-B pro-totype subassemblies and 30%new procurement.3 . 1 . 1 W H E E L STRUCTURE

The OSO-B prototype wheel structure was completely stripped of compo-nents, cleaned and used as the OSO-2 wheel structure. All the paint was re-moved fro m the interio r of the compartments, and only the GSFC Ultravioletand University of New Mexico experiment compartments were repainted blackfor flight. The paint was stripped from the ri m panels, and the exte rio r of thespacecraft was left unpainted until after the testing program was complete.Epoxy radia l st rips painted with aluminum paint were instal led on the wheel toreplace those covered with aluminum foil. New upper and lower wheel covershad to be procured for the OSO-2 spacecraft.

OSO-B prototype arms , a rm damper assemblies , azimuth shaft and spingas bottles were installed on the wheel for OSO-2. A back-up slip ring assembly

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+

w as assembled from OSO-B parts, and a new slip ring assembly w as procuredfor the flight installation. The wiring of the wheel had to be updated to theOSO-2 flight configuration.3 . 1 . 2 SAIL STRUCTURE

The OSO-B prototype sail st ru ctur e was cleaned, painted and used a s thestructure for the OSO-2 spacecraft. A compatibility f i t of the solar cell sub-strate indicated that relocation of s ev er al sail mounting holes w as necessary.One of the mounting holes was damaged when the old dwnmy a r ray was removedat Cape Kennedy pr ior to the accident. This was corrected by rive ting a back-up piece of metal behind the tor n hole.

The sail w a s mounted to the wheel st ru ctur e on 3 September 1964. Inte-gra ted comprehensive testing of the spacecraf t and experiments was conducteduntil 24 September at which time the observatory was installed on the balancemachine and the observatory balance operation was begun.3 . 1 . 3 BEARINGS

The top azimuth bearing, lower azimuth bearing and the elevation bearingsthat were installed in the OSO-2 spacecra ft were the OSO-B flight spare units.The OSO-B prototype top azimuth bear ing housing was also ins talled on OSO-2.

The OSO-B bearings experienced a grinding of su rfaces because of a loss oflubricant. Par t of the lubrication process was to place the bearings in a centr i-fuge, but this caused some of the lubrication to be forced from the bearings.Centrifuge operations were reduced, and as a result, the bearings used in OSO-2 were properly treated by a revised process. The quantity of lubricant impreg-nated into the retainer is now determined by weight.3 .1 .4 ATTITUDE CONTROL SYSTEM

Fifty percent of the OSO-B2 spacecraft attitude control system was com-posed of OSO-B flight and prototype subassemblies. The remaining fifty percentconsisted of OSO-B flight spa re s and new procurement.

The fine eyes used on OSO-1 contained a deposited knife edge. The vendorwent out of business, and as a resul t , OSO-B fine eyes w e r e provided by a filmemulsion process. This was a better process, however, prolonged exposure tothe sun caused a change and degraded the eyes. A deposited knife edge usingaluminum was developed which proved to be three t o four tim es smoother thanthe knife edge used on OSO-1.

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New solenoid valves were procured for OSO-2 which incorporated in-linefilters in addition to the regular filt ers . Leaks were encountered with the newvalves, and a s a resu lt the specifications were tightened on the finishes of theseating surfaces . Acceptance tes ts were changed to as su re that failur es wouldnot occur again. The Ball Brothers Research Corporation bench te st fixturewas a ls o cleaned, and the specification was tightened on the allowable par tic lesize.

The acceleration switches passed the acceptance tests, but l at er they werefound to have a problem due to an increased bearing friction. It was determinedthat the switch problem was due to excessive bearing preload. The switcheswe re modified by using a preload washer ins tead of the old method of adjus tingthe bearing spacing.3 .1 .5 COMMAND SYSTEM

Practical ly the en tir e OSO-B command system w a s reusable for OSO-2.The only subassemblies fr om OSO-B that were not used were the hybrid circu-lator assembly and the V H F diplexer assembly. These two units were takenfrom OSO-B flight spares.

The OSO-B flight decoders were returned to the vendor fo r inspection andtesting. The prototype decoders were installed in the spacecraft to check outthe command system while the flight units were at the vendor being test ed. Theflight decoders were returned to Ball Brothers , given a receiving inspectionand installed in the spacecraft.

A vendor representative conducted checks on the OSO-B flight and prototypecommand receivers. After the completion of the se checks , it w as found that allfour of the receivers were suitable f o r flight usage. The OSO-B flight unitswere selected and installed in the spacecraft.3 .1 .6 TELEMETRY SYSTEM

The telemetry system installed on OSO-2 consisted largel y of OSO-B lightspare units. The OSO-2 telem etry system consisted of 57%OSO-B flight spare s ,28% OSO-B flight components and 15% OSO-B prototype components. N o newsubassemblies had to be procured for OSO-2.

One of the OSO-B flight multiplexer-encoders had the epoxy cracked beyondrepa ir and this unit was scrapped. The other OSO-B flight unit had an encodinge r r o r in one channel and a bad solder joint. This unit was later reworked anddesignated for use on OSO-C. The OSO-B flight spare and one prototype multi-plexer-encoder were installed as OSO-2 flight components.

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Five tape recorders were returned to the vendor with instructions to dis-assemble all mechanical modules. A ll modules w e r e to be examined prior toreplacement of bearings. The vendor was also instructed to pick th ree tape re -co rder s that could be rebuilt - two for flight units and one fo r a flight spare . Acomplete check was made of all electrical components. A special lubricatedtape was required for these tape reco rd er s, and this was orde red from Minne-sota Mining and Manufacturing Company on the same specification as earl iertapes. The tape that was de livered was not the sarne as e ar lie r tapes andprovedto be unsatisfactory. As a result, two year old tape was used to make up a reelfo r the tape reco rders. One reel of tape had an ex tra drop-out which was notpresen t in the other ree ls . End of tape and splice drop-outs w e r e expected, sothis ree l was installed in the tape r ecorde r originally designated as the alternateflight unit. During the testing program, this tape rec ord er had signal drop-outsat two tape positions. Pre liminary investigation showed that the drop-out a reain addition to the sp lice was approximately 20 bits long compared to about 7 bitsfor a splice. This amounted to a loss of le ss than 3 words per orb it. This unitwas changed-out. However, la te r in the testing program, the primary tape re -corder developed er ra ti c drop-outs, and the tape reco rder with the ex tra drop-out was insta lled in the primary position because of its predictable drop-outarea.

During checkout, when the University of Minnesota experiment was turnedon or off, the tape recorder went into a playback cycle. It was discovered thattransient s from the Minnesota experiment and other sou rce s caused the TimeMarker Generator to tur n on. It normally took 0.75 volt to tri gger the timema rker . Modifications were made to the Time Marker Generator to increas ethe threshold voltage to 8 volts to trigger the time mar ker . A switching relaywas also placed in the PCM Time Marker power line.

The transmitte rs installed on OSO-2 were originally an OSO-B flight and aflight spa re unit. Trans mitter cur ren t measurements were made after the ob-ser vatory acceptance test s were completed. During these measurements it wasnoted that the transmitters operated in two disc ret e cur rent modes that differedby 60 ma. The normal current with the arms down was 7 0 ma. The two curren tsmeasured were 70 ma and 130 ma. This was determined to be a normal condi-tion for a bad mismatch such as occurs with the a rms down. When the trans -mit ter s ar e operated with the a rm s up the normal curr ent is 100ma.3.1.7 ELECTRICAL POWER SYSTEM

A new set of slip rings to the same specification as OSO-1 were procuredfor use in the OSO-2 spacecraft. The OSO-1 slip ring assembly exhibited auseful life span of two ye ar s. Its total life span included OSO-1 assembly, te stand launch a n d orbital operation.

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A new s ol ar a rr ay had to be procured for the OSO-2 spacecraft. Duringvibration testing, an open cir cui t was detected in the ce nte r panel and a poorsolder connection w as found in one of the side panels . Vibration tests werecontinued with the open ci rcui t still present. The ar ra y was returned to the

Y h i l P 2t the vendnr'sfacility, a solde r ball was als o removed. Furth er tests revealed no more dis-crepa ncie s, and the a rr ay was installed on the spacec raft.----J-- -- A +h- n-n~\nnnv .r w n - n c l i n l o n t i n n v x r n c tq lmnV G l l U U I CUlU b l I b I lbbbUUUSJ S b I I I U U I L u L L Y U L V l l 1.U" "-*&"I*.

Because of the loilg lezd time in procuring new s q k b batteries, it was de-cided to use OSO-B flight bat te rie s on the OSO-2 spacecraft. A set of squibbatteries was assembled from cells on hand that were less than four monthsold. One pack of wheel squib batter ies developed a shorted cell during vibrationtesting at Ball Brother s Research Corporation and was replaced with a new pack.

Ball Brothers Research Corporation originally proposed to use 45 main bat-tery cells which they received for OSO-C. They a lso had 34 cells on hand thatwere seve ral months old. Because of excessive loss of F cells in incoming in-spection and acceptance te st s, fabrication of flight batte ries from new cellscould not be accomplished consisten t with the proposed launch schedule. It wasnecessary to use the reconditioned OSO-B main batte ries. The pack wiring wasupdated and a deep discharge cycling of the ba tte rie s was accomplished. Afterdeep cycling of the bat ter ies it was found that the ba ttery capacity was nearlythe same as it was a year and a half before when they were originally acceptancetested.

3 .2 EXPERIMENTSAl l the experimental instruments w e r e removed fro m the OSO-B spacecra ft

and returned to the ir re spec tive investigators to determine the extent of damageas a result of the accident. A ll the experiments were physically and functionallychecked and the necessary repairs were accomplished. Only the Harvard Col-lege and Ames Research Center experiments w e r e replaced with the OSO-Bflight spar es . This was done because the damage incur red by the se instrument scould not be repaired in the time limit set by the established schedule. TheOSO-B flight units were rew orked, however, and were used as OSO-2 flightsp ar es . The University of Minnesota flight experiment was initially insta lledin the OSO-2 spacecraft, but it w a s later replaced by the OSO-B flight sparewhich had been reworked to increase its sensitivity. A ll the experiment instru-ments were returned to Ball Brothers Research Corporation during August 1964for incorporation into the OSO-2 spacecraft.

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3 . 2 . 1 POINTED EXPERIMENTS3 . 2 . 1 . 1 Naval Research Laboratory InstrumentThe instrument installed on the OSO-2 spacecraft was basically the original

OSO-B flight model. Only minor repa ir s had to be made to the ultraviolet tele-scope and coronagraph section such as straightening the bulkhead, replacementof a collar in the occulting support and replacement of the heat shields. Theinstrument was delivered to Ball Brothers on 31 August 1964 for incorporationinto the OSO-2 spacecraft.

During the thermal-vacuum tests conducted on the observatory, it w as notedthat when the NRL nstrument was turned on, it interfe red excess ively with theHarvard College data count. This condition was co rrec ted by placing RF chokesin series with the common power line between the Naval Research Laboratoryand Harvard instruments.

3 . 2 . 1 . 2 Harvard College Observatory InstrumentHarvard's flight spa re unit w as successfully qualified and delivered to Ball

Brothe rs Research Corporation on 24 August 1964. It was essentially the s ameas the original flight model except for an additional command for wavelengthselection.

Rework and test of th is instrument went well up to the observatory accept-ance te st when a noise problem w as detected during the thermal-vacuum testing.The noise was present in the Harvard data count whenever the spacecraft con-trol system w as operating o r when the Naval Research Laboratory instrumentwas turned on. The noise problem w as solved by installing RF chokes in se r ie swith the flexprints between the Harvard instrument and the spacecraf t. Newflexprints were provided which were made of mylar and not coated as usual, butwere painted with aluminum paint instead. The noise generated as a result ofthe Naval Research Laboratory instrument operation was eliminated by placingRF chokes in series with the common power lines between the Naval ResearchLaboratory and Harvard instruments.3 . 2 . 2 WHEEL EXPERIMENTS

3 . 2 . 2 . 1 GSFC Ultraviolet SpectrophotometerThe experiment installed on OSO-2 w as the original OSO-B flight instru-

ment. A checkout of the unit was performed and relatively lit tle damage wasfound. Repairs consisted mainly of clean-up, readjustment and checkout ofoptics and electronics.

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The instrument was delivered t o Ball Brothers on 25 August 1964 fo r in-stallation into the OSO-2 spacecraft. It perfo rmed well during integration andacceptance t es ts of the observatory with the exception that the Elgin GratingStepper was off sequence. The grating continued to cycle, however, and wouldE& czcse loss cf data- &xr-jng data rp&irtt.ion.

3 . 2 . 2 . 2 GSFC Low Energy Gamma Ray TelescopeThe experimental imtnmefit f l m n on OSO-2 WES the original OSO-B flight

model. Only minor repa ir s had to be made such as addition of magnetic shieldsand the installation of a new optical coupler. The instrument was delivered toBall Brothers on 31 August 1964, and it performed well during integration andobservatory acceptance testing. There was an indication of a loss of backgrounddata during thermal-vacuum tes ting , but it w as found to be caused by a poorlycollimated light source in the test setup. The experiment performed as ex-pected when exposed to the sun.

3 . 2 . 2 . 3 Ames Research Center Emissivity DetectorsThe Ames experiment installed on OSO-2 was the original OSO-B flight

spa re unit. A check of the OSO-B flight unit indicated the possibil ity of damageas a resu lt of the accident. The OSO-2 experiment was delivered to Ball Broth-ers on 24 August 1964. No problems were encountered during the preparationof the instrument o r during integration and acceptance testing.

3 . 2 . 2 . 4 University of Minnesota Zodiacal Light TelescopesA check of the OSO-B flight instrument indicated no apparent defects as a

re su lt of the accident. This unit was delivered to Ball B rothers on 17 August1964 and initially installed on the OSO-2 spacec raf t. The OSO-B sp ar e unitwas reworked and updated to inc rease its sensitivity by replacing photomulti-plier tubes and providing a bet ter telescope protective device. This updatedinstrument w a s delivered to Ball Brothers on 12 October 1964 and was installedon the OSO-2 spacecraft. There were no significant problems with this experi-ment, and it performed well during the integration and acceptance tests.

3 . 2 . 2 . 5 University of N ew Mexico High Energy Gamma Ray TelescopeThe University of New Mexico instrument flown on OSO-2 was the original

OSO-B flight instrument. A complete check of the ins trument was made and noserious difficulties were found as a result of the accident. The OSO-2 flightunit was delivered to Ball Brothers on 31 July 1964 and performed well duringthe integration and acceptance tests .

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After arrival at Ball Brothers, the New Mexico experiment indicated a sec-to r readout anomaly. This w as known before the accident because th eir counter ,when exceeding the maximum count rate of 31, did not r es et t o ze ro but consist-ently read out the sixteenth bit during ground test. Since the nominal sp in rateis subject to only 5% variation and New Mexico can tolerate a 15%change with-out exceeding thei r maximum count ra te , this situation was not considered aproblem. If the spin rate should vary as much as 20% during flight, it wouldstill be possible t o determine the sector by looking at the spin rate housekeepingdata. This anomaly could have always existed since the experiment was neverexposed to more than a 15%spin ra te change until just before the OSO-B accident.

3 . 3 OBSERVATORY3 . 3 . 1 WEIGHT A N D BALANCE

The OSO-2 spacecraft was buttoned up fo r the balance operation and mountedon the balance machine on 24 September 1964. A dummy solar array, dummy so-lenoid valves and a tes t se t main battery were installed for the balance operationbecause the flight units were not yet ready for installation. The observatorypre-shake balance operation was completed the following day. The observatorywas weighed after the balance operation, and the launch weight was computed tobe 549.7 pounds. This included six pounds of weight fo r the flyable lifting lugs.Figure 3-1 shows the OS0 spacecraft in the Ball Brothers Balance Machine.3 . 3 . 2 ACCEPTANCE TESTS

Acceptance testing of the OSO-2 spacecra ft consisted of vibration testingand thermal-vacuum testing. The observatory was prepared for the vibrationtesting and the sine sweep was completed on 5 October 1964. A brief mechani-cal inspection indicated no failures, but the solar array continuity check indi-cated an intermittent circuit. A poor solder joint was discovered and repai redwhich appeared to fix the discontinuity. The Z axis sinusoidal vibration wasperformed, and all spacecraft systems and experiments checked out satisfac-torily. The random Z axis vibration tes t was completed on 6 October, and noproblems were observed during a quick visual inspection of the observatory .The X axis vibration testing was performed and completed on 7 October. Dur-ing these tests, the right hand panel of the solar a r ray indicated discontinuitiesin both the sine and random vibration operations. All three solar a rra y panelswere returned to the vendor for failure analysis and rep air . The sola r arr aywas returned from the vendor on 9 October, and they indicated that they couldfind nothing wrong with it . A receiving inspection was performed on the solararray, and by illuminating individual shingles, it was found that one shingle was

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open in the center panel. The center panel was returned to the vendor fo r re-pa ir. The cause of the intermittent in the right hand panel was discovered t o bea shorted solar cell caused by a minute solder ball. Both the right and left pan-els were returned to the vendor for comprehensive inspection. The solar arra yw as returned fro m the vendor and installed on the sp acecra ft on 15 October 1964.The Y axis vibration was completed on 16 October , and no mechanical failu reswere found during a brie f visual inspection. Figure 3-2 illustrat es the vibrationtest set up.

Thermal-Vacuum testing consisted of placing the observa tory in a vacuumchamber and pumping it down to evacuate the air. The observatory was thenoperated, as it would be while in orbit, for a period of 7 2 hours at a tempera-ture of 25 degrees Cent igrade . After the "Hot Cycle" was completed, the ob-servatory was operated for 72 hours in the "Cold Cycle" at a temperature ofzero degrees Centigrade.

The observatory was installed in the vacuum chamber on 22 October 1964,and the Thermal-Vacuum testing was begun. During the "Hot Cycle" tes ts , theHarvard instrument recorded spurious noise at the serv o transition points ofthe scan pattern. This noise was caused by the driving of the torque motors,and was a known phenomena. Fi lte rs were placed in se ri es with the Harvardflexprint to reduce the noise. It was found that this would reduce the noise onlyif the metallic plating was not pre sent over the flexpr int. Because of this , theOSO-2 had to use aluminum painted flexprints instead of plated flexpr ints . Sup-plementary high temperature vacuum te sts were run with the stripped flexprints ,and no discernible noise effects were noted from the control system operation.However , operation of the Naval Research Laboratory experiment caused noisehigher than des ired by Harvard, and fil ter chokes had to be placed in the com-mon power leads of the two experim ents. The cold thermal-vacuum tests werebegun on the morning of 31 October 1964. No seri ous problems were encount-ered with the %old cycle, I' and the observatory was removed from the vacuumchamber on 3 November 1964.

3 .4 SHIPMENT TO CAPE KENNEDYThe spacecraft w a s buttoned up for shipment to Cape Kennedy on 16 Novem-

ber 1964 and placed on the balance machine to check the alignment of the align-ment nubbin. The alignment was not within the Douglas Aircraft Companyspecification and had to be co rre cted . On 17 November 1964, another balanceoperation was completed, and the alignment of the spin axis perpendicular tothe separation plane was checked. The back of the so lar a rr ay substra te wasgiven a second coat of white paint which was nec essary a s indicated by emis-sivity measurements which were made.

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The payload was installed in its shipping canis ter on 20 November 1964 andleft by truck for Lowry A i r Force Base together with its support equipment.The shipment w a s loaded aboard a Military Air Transp ort Service airc raf t on21 November 1964, but it had to be unloaded for transfer to another aircraft be-cause of mechanical problems with the first airplane. Figure 3-3 shows theloading of the Military A i r Trans port Service airc raft . The second airc raftarrived at Lowry Air Force Base on 22 November 1964, but the shipment wasreloaded on the first ai rcraft due to a development which made the second air-plane unavailable. The airplane waited for departure on 23 November 1964 sothat it would arrive a t Cape Kennedy during the daylight hours. This was nec-es sa ry because the landing s tr ip at Cape Kennedy is not equipped with landinglights.

3.5 PRE-LAUNCH PREPARATIONSThe spacecraft arrived at Cape Kennedy at 0730 EST on 23 November 1964,

only five months after the official go-ahead for OSO-2 development was given.A check of the external radioactive sourc es indicated that the University of Min-nesota's Carbon 14 sources had leaked and were emitting beta rays. After de-contamination activi ties had been accomplished, everything seemed all right.Some of the accele ration switches mounted on the spacec raft and container werefound to be tripped when the spacecraft was uncanned at Cape Kennedy. It w a sdetermined that the switches were probably tripped due to handling during theloading operations at Lowry AFB.

.

3 . 5 . 1 HANGAR CHECKSThe spacecraft was removed from its shipping cani ste r and placed in the

clean tent in hangar AE on 25 November 1964. No visible damage was incurredby the spacecraft. Radio frequency checks were completed on the spacecraftexcept fzr xdtage standing wave ratio measurements which could not be per-formed due to faulty test equipment. A l l eight experiments were given compre-hensive checks and were found to be all right. The main batteri es were placedin a deep discharge cycle to check the ir capacity. The battery discharge cyclingtests were completed on 1 December 1964, and the measured capacity of 18ampere-hours was far in excess of requ iremen ts. On 4 December 1964, thespacecraft was placed in its shipping canister and readied for extended storagein hangar AE to await the a rr iva l of the Delta third stage. At this point, launchoperations were suspended.

The Ball Brothers Research Corporation crew arrived again on 6 January1965 and the spacecraft was again uncanned and installed in the clean tent.

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Comprehensive and sun pointing tests were conducted on the spacecraft and ex-periments. On 8 January 1965 the fa iring compatibility f i t checks were succes-fully completed and the radio frequency in terference te st s w ere begun. R Fspectru m measurements on trans mitt er number 1 with the arms down indicatedmultiple sidebands at approximately 2 Mc inte rva ls. This phenomena had beenobserved during bench checks when a transmitter w a s connected to a severelymismatched reactive load.

The prime tr ansmitt er was being coupled into the command rece ivers , andthis w a s discovered to be caused by a shif t of the plastic s leeves on the ends ofthe antenna stub s. This prevented the stubs from touching the a rm s when thearms were down. When the sleeves were relocated to the co rr ec t position, thestubs touched the a rm s and tran smi tter operation w a s normal with no couplinginto the command receivers.

It was determined that the compatibility between the transmitters and thearms down antennae was improved if the arms down antenna impedance wasmade slightly reactive. This was accomplished, without changing the a rm s upimpedance, by changing the protective tip sleeving on the driven elements sothat elect rica l contact was made with the spin bottle bracke t. With this modifi-cation, the command and telemetry systems functioned normally with the armsdown.

Later, it was discovered that spurious signals from the transmitters werestill getting into the command system. Arms down impedance adjustments weremade by changing the length of the R F cables. An attempt w a s made to adjustthe length of the diplexer to "T" Power Divider coaxial cable to eliminate thespurious signals with the a rm s down and to match both transmi tte rs with thearm s up fo r optimum performance. The antenna stubs were insulated from thea rm structure with heat-shrinkable tubing. Tes ts af ter changing the cablelength indicated that neither transmitter could be made to generate spurioussignals with the ar ms down. Tra nsm itter number 2 was not delivering adequatepower with the a rm s up, s o the output coaxial relay cable was replaced with anew length cable. After this replacemen t, c ur ren t into both tra nsmi tte rs w asnormal under all conditions, power into the matched load was normal and thestanding wave ratio was norm al a t both frequencies.

On 13 January 1965, emissivity and absorptivity measurements on the ther-mal surfaces of the spacecraft were made. The wheel experiments were givena pre-comprehensive check-out on 14 January 1965. A l l but the GSFC Ultra-violet and Gamma-Ray exper iments functioned normally. During the wheel pre -comprehensive checks , the GSFC Ultraviolet instrument was commanded "ON",but no data were observed at the telemet ry ground station . However, the wave-length selector stepping was audible which indicated that the instrument was

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operating. An oscilloscope w as connected to the ri m panel tes t connector, anddata were immediately received at the ground station. The data amplitude w as4.4 volts, fur ther investigation disclosed no abnormalities. The instrumentwas turned on again at 16.2 volts to determine any possible voltage discrepancy,but normal operation was observed. On 20 January 1965, the instrument w asturned on simultaneously with the spacecraft and data were received immedi-ately. On 24 January 1965, the instrument buffer module w as installed and acheck revealed proper operation. Proper operation was again verified duringchecks made on 26 January 196.5.

The GSFC Gamma-Ray experiment appeared to malfunction during pre-comprehensive checks on 14 January 1965. A t that time, however, there w asa data indication from the dead-time analog monitor. A two (2) microcurieRadium-226 source was substituted f o r the two Sodium-22 so urces normallyused, and a digital count was observed. A check of the digital numbers versusthe source distance w as made, and normal operation was observed. La te r,checks were made using the two Sodium-22 sources once again, the the properdigital counts were observed.

The spacecraft was prepared for the balance operation on the Ball BrothersBalance Machine. The balance operation indicated a slight unbalance which wascorrected. Perpendicularity measurement between the bearing axis and theplane of the attach fitting indicated a misalignment of approximately th ree quar-te r s of an a rc minute. This amount of misalignment would not affect the space-craft performance, but it would result in the alignment nubbin having a run out-side the Douglas Aircraft Company tolerance. The run out was not corrected,but Ball Brothers notified Douglas of the offset of the nubbin so that the effectcould be taken into account upon mating.3.5.2 SPIN A N D BALANCE CHECKS

During the evening of 29 January 1965, al l non-flight handling covers wereremoved from the OSO-2 observatory in hangar AE. The protective capsulewas installed around the observatory and purged with nitrogen gas to provide anine rt atmosphere. The observatory and the capsule were placed in the shippingcontainer and transported to the Spin Test Area on a flat-bed truck at 2130hours.Upon arrival at the Spin Test Area, the observatory and capsule were removedfrom the shipping container and weighed. The nitrogen purge line w a s connectedto the capsule purge port. The payload was mated with the third s tage andplaced on the alignment fixture to check the alignment of the alignment nubbin.Upon completion of the nubbin alignment, the thi rd s tage and observatory weremoved to the spin balance fixture, a nd the spin balance operation was performedby the Douglas Aircraft Corporation personnel.

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3 .5 .3 LAUNCH TOWER CHECKSWhen the spin balance operation was completed, the spacecraft was again

installed in the shipping container and transported to the launch complex at 1800hours on 31 January 1965. High winds delayed the mating of the second andthird stages until approximately 1900 hours.

3 . 5 . 3 . 1 T-3 Day ChecksT-3 day checks were begun at 2200 hours on 31 January 1965 and were suc-

cessfu lly completed approximately 0300 hours on 1 February 1965. After ap-proximately 25 minutes of eqiupment set-up and observa tory and equipment turn-on, rem ote communications checks were performed. Remote experiment statuschecks and control system checks were performed. The spin and pitch gas sys -tem s were tested for leaks and then each was pres sur ized to 21OOpsi. The taperecorders were loaded with tape and operated in the playback mode to time theplayback period. A t the end of the tape recorder operation, all power was re-moved from the observatory and the T-3 day checks were sec ure d.

3 . 5 . 3 . 2 T-1 Day ChecksT-1 day checks were initiated late in the evening of 2 February 1965 and

were completed in the ea rl y morning of 3 February 1965. Countdown tasks per-formed on T-1 day were to verify the flight readiness of the launch vehicle,space craft and facilities equipment; to a ssure compatibility with Range Systemsfor launch; to install spacecra ft fairing s; and to install and electrically connectordnance. Clearance to start the count was obtained from all systems, a com-munications check was performed and the necessary power was turned on. Acheck was made to verify that the engine sequencing and cut-off circuitry wereperforming properly. The vehicle tanks and bottles were pr essuri zed , and thelubrication oil tank w a s filled. The second stage re tr o sys tem was pressurizedand topped.

Comprehensive performance tests by R F link were performed on the flightspacecraf t and the third stage telemetry system . Performance measurementswere made of the spacecraf t systems and the third stage telemetry while oper-ating. The vehicle electrical sy stems were checked for prope r performance.The RF systems received an open loop composite t es t , first on external powerand then on internal power, using the applicable Range ground transmitting andreceiving systems. The flight batte ries were sec ured and connected prio r tothe internal power checks.

Preparations were made for prope llant flow and loading of the second st age .Samples were taken from the propellant tr ai le rs and tested . After a leak check

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History of Orbiting Solar Observatory OSO-2

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