ariel i, the first international satellite project summary

8/8/2019 Ariel I, The First International Satellite Project Summary

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NASA SP-43

ARIEL I

The FirstInternationalSatellite

N A T I O N A L A E R O N A U T I C S A N D S P A C E A D M I N I S T R A T I O N


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A R I E Y I O N A L S AT E L L I T E


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Ariel I, t he International Ionosphere Satellite, in orbit attit ude.


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NASA SP-43

ARIEL I

The H i r s t

InternationalSatellite

The Project Summary

Prepared by

GODDARDSPACEFLIGHT CENTER

Greenbelt, Maryland

O f i c e of Scienti f ic an d Technical Inform at ion

N AT I O N A L A E R O N A U T IC S A N D S PA C E A D M I N I S T R AT IO N

1963

Wushing ton , D. C .


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F O R E W O R DOn April 26, 1962, the first international satellite,

Ariel I, was successfully launched from Cape Ca-naveral, Florida. This satellite, designed for iono-spheric research, represents two years of cooperativeeffort between the IJnited Kingdom and the UnitedStates. Th e history of Ariel I and its experimentshas been compiled in brief form by members ofNASAs Goddard Space Flight Center. This his-tory, consisting of reports and descriptions con-

tributed by the U.K. and U S . scientists assigned tothe Ariel I project, is presented herein as a projectsummary.

ROBERT . B A U M A N N ,S. riel I Project ManagerGoddurd Space F1igh.t Center

Greenbelt, M ary lan d


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PA G E

CONTENTSV

11113.5

77

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1 11 11 11314

151517181921222727

292933

3546

47474748495 5


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viii


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CHAPTER 1

IntroductionPURPOSE OF DOCUMENT

The Ariel I (1962 01) project summary docu-ment has been compiled to record the historyof the first internat ional satellite. Herein aredescribed the events leading up t o the concep-tion of Ariel I, the technical and managementplans used in the project development, the sci-entific and engineering considerations employedin the design of the satell ite and support equip-men t, and the design detai ls of these compo-nents. A list of related documentation con-taining design information and experimentalresults is also included.

LIST OF CONTRIBUTORS

This document is a collection of descriptionsand reports authored prior to and after thelaunch of Ariel I (April 26, 1962). These con-tributions have been edited, revised, and up-dated to produce continuity. The followinglist includes the authors who have providedmajor contributions to this summary.

United Kingdom Contributors

Dr. P. J. Bowen, University College LondonProf. R . L . F . Boyd, University College LondonMr. A. C. Durny, Imperial College LondonProf. H. Elliot , Imperial College LondonDr. K. A . Pounds, University of LeicesterDr. J. J. Quenby, Imperial College LondonProf. J. Sayers, Univers ity of BirminghamMr. J. Wager, University of BirminghamDr. A. P. Willmore, University College London

United States Contributors

(National Aeronautics and Space Administra-tion, Goddard Space Flight Center)

R. C. BaumannR. E. BourdeauA. BuigeP. T. ColeC. F. FuechselW. H. Hord, Jr .R . E. KidwellV. L. KruegerT . J. Lynch

R. G. MartinW. H . MeyerM. SchachJ. C. SchaffertJ. T. SheaJ. M. TurkiewiczC. L. Wagner, Jr .H. D. White, Jr .F . C. Yagerhofer

GENERAL INTRODUCTION

Ariel I , the first international satellite, wasdesigned to contribute to mans knowledge ofthe ionosphere and its complex relation to thesun.

This project developed from proposals madein 1959 to NASA by the British National Com-mit tee on Space Research. These proposalswere in response to a United Sta tes offer t o theCommittee on Space Research (COSPAR) ofthe Internat ional Council of Scientific Unions

to launch scientific experiments or completesatellites prepared by scientists of other nat ions.The content of the program and the division ofresponsibility between NASA and the BritishCommit tee were agreed during discussions t ha ttook place in late 1959 and early 1960. Sub-sequently, the NASA Administrator assignedproject responsibility for the United States tothe Goddard Space Flight Center (GSFC) .

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ARIEL I : THE FIRST IN TE RN AT IO NA L SATELLITE

GOD DAR D SPACE UNIVERS ITYFLIGH T CENTER PROJECT SCIENTIST COLLEGE COLLEGE

This assignment included the design, fabrica-tion, integration, and testing of the spacecraftstructure, power supply telemetry, commandreceiver, thermal control, and data storage.GSFC supplied the vehicle, was responsible forlaunch, performs dat a acquisition via the world-wide Minitrack network, and provides dataprocessing. Th e United Kingdom (U.K.) had

U N I V E R S I N O FBIRMINGHAM

the responsibility for the design, fabrication,and testing of all flight sensors and their asso-ciated electronics up to the telemetry encoderinput. The U.K. also is responsible for dataanalysis and interpretation. Th e distributionof responsibilities is outlined in Figure 1-1.

A list of the experiments and electronic sub-systems follows.

ASPECTOSMIC RAY

UNITED KINGDOM

ELECTRONDENSITY

-DATAATELLITE TRACKING SATELLITESTRUCTURE 8 6 DATA INSTRUMEN-

MECHANISMS ACQU ISITION TATION PROCESSING

8 DATA

VEHICLE6LAUNCH

C O O R D I N AT I O N

TELEMETRYI

ELECTRONSADIATIONL

h------

-ALPHADFIGURE -1. Prime responsibilities for Ariel I .

Experiments

Electron Temperatu re and Densit?University College London

This experiment, based on Druyvesteyn'smodification of the Langmuir probe, will deter-mine the value of the electron dens ity andtemperature near the satellite.

I o n Mass Composition and Temperature-University College London

This experiment is, basically, the same as the

TEMP. 6DENSITY

SPECTRUM8 TEMP.

electron temperature experiment. However,the method of measuring the tempera ture isdifferent.

Solar Lyman- Alpha Emission Measurement-

Two par ts of the solar spectrum were selectedfor measurement: the Lyman-a lpha line ofhydrogen, and the rather hard x-ray spectrum.The function of the two solar radiation measure-ments is to enable simultaneous and nearly

University College London

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I N T R O D U C T I O N

continuous observations of t he stat e of theionosphere and of the solar atmosphere.

X - Ra y Emission-University College Lon don

solar conditions.This experiment provides an indication of

Solar Aspect Measurem ent-Univers i tyCollege Lon don

The solar aspect subsystem provides thelati tude and longitude of the sun in the satellitecoordinate system. Also, the subsystem pro-vides the satellite spin rate.

Cosmic R ay Analyzer-Imperia l CollegeLondon

The purpose of this experiment is to makeaccurate measurements of t he primary cosmic

ray energy spectrum and the effects of inter-planetary magnet ic field modulation of thisspectrum.

Ionosphere Electron Den sity Mea surement-Un ivers i ty of B i rmingham

The measurement of electron density is per-formed by this experiment, and by the Lang-muir probes of the University College London.However, the two methods of measurement arequite different and therefore complement oneanother.

Electronic Subsystems

TeZ em et ry -G odda rd Space F l ight Cen te rD at a Encoders-Goddard Space Flight Ce nterTa pe Recorder-Goddard Space Fl ight Cen terPower Sys te m- Go dda rd Space F ligh t Cen te rSpacecraft Parameters (Housekeeping)

System-Goddard Space Flight Ce nter

Spacecraft and orbital characteristics aredetailed in the following list.

Scientific Instrumentation

Electron temperature and density sensors (2)Ion mass sphere (1)Solar radiation detectors, Lyman-alpha at

Solar aspect meter (1)Electron density sensors (1)X-ray counters (1)

1216A (3)

Spacecraft Characteristics

Size, basic structure-23 inches O.D. by 22

Weight-I36 poundsSpin rate -36 to 12 rpm throughout lifeLifetime-1 year

Power-p-on-n solar cells and nickel-cadmium

Da ta Storage-100-minute tape recorderAntenna-modified crossed dipoleTracking and da ta frequency-136.408 Mc

Orbit Parameters

inches high

batteries

nominal

Perigee-390 km (210 nautical miles)Apogee-1212 km (654 nautical miles)Inclination-53.86 degreesPeriod-1 00.9 minutesEccentricity---0.057

Three complete payloads were constructed:one for prototype testing, a flight model, and abackup in case of malfunct ion in the first launchatt emp t. Provision was made for two launch-ing vehicles, including one for backup. Ariel Iwas launched on April 26, 1962. A cutawayview of the satellite is shown in Figure 1-2.

RELATED WORK

Experiments relating to those in the Ariel Isatellite have been performed in soundingrockets and satellites. Fu ture plans includeaddi tional experiments in t he following fields.

Ion and Electron Studies; Electron DensityMeasurements

Experiments similar to the ion mass spec-trometer, the Langmuir probe, and the electrondensity experiments were flown successfully byBourdeau and collaborators on the ExplorerVI11 satellite (1960 E l ) , which was launched onNov. 3, 1960. Results have been published inthe open literature (References 1 to 5). Aspherical ion mass spectrometer was also flownon Sputnik I11 (1958 62) by the USSR (Refer-ence 6).

NASA plans to use similar experiments onthe Atmospheric Structure Satellite and onthe EGO and POGO satellites.

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ARIEL I : THE FIRST INTERNATIONAL SATELLITE

0

FIGURE -2. Cutaway view of Ariel I.

Cosmic Ray

The primary purpose of the Ariel I cosmicray experiment is to carry out a satellite in-vestigation of time var iations of heavy nucleiin the primary cosmic radiation with a thin-walled omnidirectional Cerenkov counter. Anexperiment with a similar objective but usinga completely different approach was carriedout on Explorer VI1 (1959 ( 1 ) . In this casePomerantz and Schwed used a thin-walledomnidirectional ionization detector. Again, theprimary objective was the satellite investiga-tion of time var iat ion of heavy nuclei. Duringthe eight months of operation of this detectoron Explorer VII, significant changes in thespectrum of heavy nuclei were observed.

Detailed accounts of the findings of thissatellite can be found in References 7 , 8, and 9.

Solar Radiation

The purpose of the x-ray and ultravioletmeasurements is to increase knowledge of th esuns ionizing radiations throughout the solar

cycle. The measurements of solar radiationare similar to experiments previously performedin space vehicles.

Solar x rays were first quantitively measuredin 1949, utilizing thin beryllium-window photo-counters in a V-2 rocket flight (Reference 10).The Lyman-alpha line of hydrogen was firstdetected by Rense in 1952 (Reference 11 ) .Since that time, there have been several dozenmeasurements of the solar x-ray and hydrogenLyman-alpha intensities utilizing variousrockets: Aerobee-Hi (with solar pointing con-trol s), Nike-Deacon, and Nike-Asps. I n addi -

tion, experiments have been flown on ExplorerVII, the Naval Research Laboratory Grebsatellites, and on the Orbi ting Solar Observatory(1962 f l ) these include radiation measure-ments of both the quiet and active sun.A survey of the results of measurements of thesuns ionizing radiations is given by Friedmanin Reference 12.

The Orbiting Solar Observatory I1 (OS0 11)satellite scheduled for 1963 will continue

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I N T R O D U C T I O N

measurements of solar x-ray and ultravioletradiations. This O S 0 series of satellites willmake the most comprehensive coverage of th esolar electromagnetic radiations ever att emp ted .For example, experiments will be flown tomeasure the 2500 to 3000A band, hydrogen

Lyman-alpha line, 400 to 10A band (with 1Aresolution), 1 to 8A region, 100 kev to 3 Mevgamma rays (100 kev to 1.5 Mev with 100kev resolution), and 50 to 500 Mev gammarays. *

RELATED DOCUMENTS

In addition to the specific reports referencedin the text, there are many reports relating tothe design, development, and experimentalmeasurements of Ariel I. The reference andbibliography list a t the end of this summaryreport gives the specific reports referencedherein as well as an additional bibliographyof related documents.

*For additional information, refer to the GSFC O S 0I1 Project ,Development Plan .

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CHAPTER 2

Management PlanAPPROACH

Th e approach of the Ariel I project manage-ment utilized a concept of a joint UnitedStates-United Kingdom working group withvarious ad hoc committees named as required.The basic Ariel I working group membership

was as follows:United Kingdom

University of Bir-mingham

Professor J. SayersUniver sity Colhge

LondonDr. H. ElliotDr. J. J. Quenby

Imperial CollegeLondon

Dr. R . L. F. BoydDr. A. P. WillmoreMr. M . 0 . Robins

Dr. E. B. Dorling

EZectron Density

Project Scientist

Cosmic Ray

Project ScientistAlternate

All Other Experiments

Project ScientistAlternateU.K. Project Man-

agerU.K. Coordinator

United States

National Aeronautics an d SpaceA d m i n i s t r a t i o n

NASA Headquarters:Dr. J. E. NaugleMr. M. J. Aucre-

manneGoddard Space

Flight Center:Mr. R. C. BaumannMr . R . E. BourdeauMr . J. T. SheaMr. H. J. Peake

U S . Project OfficerU.S. Project Chief

U .S Pro ect ManagerProject ScientistU.S. CoordinatorTelemetry R F

Dr . R . W. RochelleMr . J. C. Schaffert

Mr. P. T. ColeM r. C. L. WagnerMr. F. C. Yager-

Mr. M. SchachM r. W. H. HordMr . J. M. Turkie-

Mr. C. H. LooneyMr. H. E. Carpen-

Mr. C. J. CrevelingMr. A. Buige

hofer

wicz

ter

Mr. C. P. Smith

Mr. R . H. Gray

Telemetry codingSequence program-

Data storageMechanical designPower supply

ming

Thermal designEnvironmental testingElectrical systems in-

Tracking systemsTracking operations

tegration

Data reductionOperations control

Atlantic MissileRange, VehicleCoordination

Atlantic MissileRange, Operationsand Launch Di-rector

ASSIGNMENTS

Management responsibility for the Ariel Tproject was assigned as follows.

Project ManagementNASAs Goddard Space Flight Center was

assigned project management responsibilitiesfor the Ariel I project.

Experiments Sys tem Management

The United Kingdom Scientific Mission ac-cepted system management responsibilities forthe Ariel I experiments system.

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ARIEL I : T H E FIRST INT ERN ATI ONA L SATELLITE

Spacecraft Syste m Management SCHEDULE

Goddard Space Flight Center was assigned The major phases planned and accomplishedsystem management responsibilities for the throughout the program are shown in FiguresAriel I spacecraft system. Key milestones in the Ariel I-1 an d 2-2.

project are reported in the Program Manage-ment Plan, described in the next section.racking and Data System Management

Goddard Space Flight Center was assignedsystem management responsibilities for theAriel I tracking and da ta system. Acquireddata are sent to the United Kingdom afterprocessing (digitizing), and the U.K. is respon-sible for data reduction and analysis.

NASA Headquarters Direction

The Director, Office of Space Sciences. NASAHeadquarters, is responsible for overall direc-tion and evaluation of the performance of theGoddard Space Flight (enter as the Ariel IProject Management Center and as the Sys-tems hfanagement (enter for the spacecraftand the tracking and dat a systems.

The Director, Office of International Pro-grams, NASA Headquarters, is responsible fordefining and interpreting international agree-ments relating to the project and for providingthe United States (NASA) coordination withthe United Kingdom.

Each country has an Ariel I Project Manager,Project Coordinator, and Project Scientist.

All working decisions are subject to the approvalof the Project Managers. Responsibility fo rthe coordination of the many aspects of theoverall program is vested in the Project Co-ordinators. The Project Scientists are re-sponsible for the ultimate compatibility andintegration of the various experiments. Over-all policy matters are decided by the NASAAdministrator for the United States and by theChairman of the British National Committeefor Space Flight for the United Kingdom.

In the case of subsystems, the work (andcontract monitoring) was the responsibility ofthe individual in charge of the subsystem.

REPORTING PROCEDURES

Reports indicating progress in accomplishingthe scheduled milestones and summarizing theproject status were furnished by GSFC toNASA Headquarters on a biweekly basis asrequired by NASA Management InstructionG-2-3, Program Management Plans. TheU.S. Project Manager submitted to the Dir-ector, GSFC, each week a written report thatdescribed significant events occurring on theproject, highlighted problem areas, and indicatedany assistance that was required.

Periodic presentations on th e Ariel I projectare made to the GSFC Executive Council.This group, chaired by the Director of GSFC,is composed of top management officials. Thepresentations cover all significant aspects, suchas funding , procurement, etc. Emphasis isplaced on defining problem areas and applyingnecessary measures to resolve them.

The U.S. Project Manager makes monthlysubmissions of data for the NASA Adminis-

trators progress report. Such da ta includethe progress made on Ariel I project objectivesduring the previous month, progress made dur-ing current month, and plans for the comingmonth.

The U.S. Project Manager is responsible forthe preparation of semiannual budget reportson his project.

PROCUREMENT

All procurement in the U.S. was handledaccording to normal GSFC procedures. Asinglepoint procurement contract system wasused.

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M A N A G E M E N T P L A N

Legend

STRUCTURAL TEST U N IT

PROTOTYPE= LIGHT UNITNO. 1SCHEDULE Dote

INTER NAT lO N A LIONOS PHERE SATELLITE Oct. 1960

A R I E LI

FLIGHT UNIT N 0 .2

q r - f

5

6

7

TASSEMBLY-STRUCTURALASSEMBLY A INTEGRATIONSENSORS A ELECTRONICS

SYSTEMS TESTING

PRELIMINARY BALANCE

VIBRATION

8 ACCELERATIONI9 TEMP. A HUMIDITY

IO THERMAL VA CUU M

13

FINAL CALIBRATIONA TEST

FINAL BALANCE

SHIP TO FIELD

1961

Ell

' 1962-

FIGURE -1. Planned program schedule.

9600439 0 4 6 - 2


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~~

ARIEL I : THE FIRST INT ER NA TI ON AL SATELLITE

ILegend

STRUCTURAL TEST UN ITPROTOTYPE

FLIGHT UNITNO.lEzp FLIGHT UNIT N 0. 27

ESIGN

FABRICATION

ASSEMBLY-STRUCTURAL IASSEMBLY 8 NTEGRATIONSENSORS 8 ELECTRONICS

SYSTEMS TESTING IPRELIMINARY BALANCE 1

IIBRATIONACCELERATION

TEMP. 8 HUMIDITY

THERMAL VACUU M

FINA L CALIBRATION8 TEST

FINA L BALANCE

SHIP TO FIELD

SCHEDULEINTERNATION AL IONOSPHERE SATELLITE

A R I E L 1

DateSept. 1962

1962

F I G U R E-2. Actual program schedule.

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CHAPTER 3

m 1 1

General 1 echnicalBACKGROUND AND PURPOSE

In determining the experiments to be carriedout by Ariel I, every effort was made not toduplicate experiments already performed orunderway by the US. and Russia. Simultane-ously, experiments were selected to (1) take

advantage of techniques developed in the U.K.as par t of the Skylark sounding rocket program,and (2) provide an integrated assault on un-knowns connected with the sun-ionosphererelation.

Of the s b specific experiments carried onboard, five are closely interrelated and provideconcurrent mepsurement of two important typesof solar emission and the resulting changingstates of the ionosphere.

Of these Jive, the three ionospheric experi-

ments measured electron density and tempera-tur e as well as the composition and tempera tureof positive ions. Th e remaining two experi-ments monitored the intensity of radiation fromthe sun in the ultraviolet (Lyman-alpha) andx-ray bands of the solar radiation spectrum.Lyman-alpha radiations originate in the sunschromosphere (solar surface), while x-rays origi-nate farther out in the area around the sunknown as the corona. Previous work has shownthat the Lyman-alpha radiation is relativelyconstant but khat x-ray emissions are quitevariable with solar conditions.

Th e sixth experiment measured primarycosmic rays by means of a Cerenkov detectorcarried on board the Ariel I . Simultaneously,measurements of secondary cosmic rays weremade by means of ground observation and bymeans of airc raft , balloons, etc.

PlanARIEL I CONFIGURATION

Figures 3-1, 3-2, and 3-3 provide essentialmajor component outlines and relative positionsfor Ariel I . I ts basic configuration, which mustfit the envelope of the Delta vehicles payloadcompartment, is tha t of a cylinder IO%, incheslong and 23 inches in diameter. The mainelement of each closure is a spherical sectionwhose large (inboard) terminator circle is S X ,inches in diameter. Each spherical section is5)5 inches long and has an outer surface radiusof cu rva ture of 13% inches. To this basic22-inch-high configuration are atta ched th evarious elements necessary for the support andconduct of the several experiments incorporatedin the system.

Th e spin axis of the satellite is the centra l

axis of the cylinder, which-for purposes of thisdescription-is also considered as the verticalaxis. At the bottom of the satellite is a Winch -diameter fourth-stage separation flange, withinwhich is mounted t he electron temperature gageand the tape recorder.

Extending out horizontally from about mid-way up the lower spherical section, a t 90 degreeintervals around the circumference of the satel-lite, are four solar paddles.

Offset 45 degrees circumferentially from thesolar paddles, opposite mkl exactly counter-balancing one another, and extending radiallyin the same horizontal plane are two boomswith nominal lengths of 4 feet. Th e end ofone boom accommodates the two circular capac-itor plates of the electron density sensor.Electronics associated with this experiment a rehoused in a 4%-inch-diameter by 6)i-inch-longcylinder mounted on the boom close to themain body of t he satellite. At the end of the

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ARIEL I : TH E FIRST INTERNATIONAL SATELLITE

SENSOR BOOM, ELECTRON TEMPERATUREVIBRATIONNTENNAXPERIMENT

SOLAR PADDLE NO. 4E-SPIN MECHANISM /

STANDARD DELPAYLOAD ENVEL

SENSOR BOOM, ELECTRON DENSITY

FIGURE -1. Ariel I satellite-Delta vehicle compatibility, view 1 .

DE-SPIN TERMINAL BOARD( 2 )

X - 248 - A5ASPECT SENSOR X-RAY GAG E SOLAR PADDLE MOTOR

m0 2 4 6 8 I O 12 I4 16 18 20

INCHES

FIGURE -2. Ariel I satellite-Delta vehicle compa tibility , view 2.

other boom is a second electron temperaturegage, whose electronics are located inside thesatellite.

With the exception of a 3%-inch-diameterhemispherical solar aspect sensor, the centralcylindr ical section is free of protuberances.

On top of the sate llite, in line with the spinaxis, is a 5-inch-diameter cylinder containingthe cosmic ray Cerenkov detector. Above this,

12

on an 8-inch-long conical and cylindrical sectiontapering from a 3-inch to 1-inch diameter, is a4-inch-diameter ion mass sphere whose center ofmass is 1oca.ted 14 inches above the forward faceof the main body.

Four turnstile antennas , spaced circumferen-tially a t 90 degrees and angling up a t 45 degrees,are mounted on the top spherical section and areperpendicular to the sphere a t the interfaces.


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GENERAL TECHNICAL PLAN

GSFC ELECTRONIC

HARNESS SUPPORYMAN-ALPHA DETECTOR

I-YEAR CLOCKELECTRONICS

TURN-ON PLUG

BATTERY PAC KGSFC ELECTRONIC PACK

UNIVERSITYOF B I R M I N G H A MELECTRONICS PACKAGE

CONNECTOR BRACKET

TAPE RECORDER PLUG

FIGURE -3. Ariel I satellite main shell, plan view.

There are three flush-mounted solar radiation(Lyman-alpha) detectors on the satellite skin :

two at 60 degrees (one up a nd one down) fromthe equator, and one on the equator. All threeare in the same vertical plane and in the same180-degree sector. There are two proportionalx-ray counters located 45 degrees up and downfrom the equator and directly opposite theLyman-alpha gages.

ARIEL I CIRCUITRY

The electronic ci rcuitry is, in general, coveredby the functional diagram of Figure 3-4 andproduces the signals described in Figure 3-5.The United Kingdom supplied all equipmentlisted under the probes and conditioning cir-cuits of Figure 3-4, while GSFC provided allother subsystems.

CO ND ITION ING CIRCUITS AN D POWER SYSTEMS-ENCODER-PROBES SUBCARRIER OSCILLATORS CONTROL CIRCUITS-TAPE RECORDER

~ COUNTERSRAy121 XRAYSis.

,

1

I I - IGNALJ

-OWERAPERECORDER' T R O 1-'

FIGURE -4. Ariel I functional diagram.

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ARIEL I : THE FIRST I NT ERN AT IO NA L SATELLITE

DATACHANNEL

67

OUTPUTPERIOD(rrrccl

SAMPLE OF HIGH SPEED ENCODER OUTPUT

~ 9 ~ i I l i s e c ~ - 100 mil l i r rc

93 .8 TO 312 cpr-

BINARY BITS INDIGITAL OSCILLATOR

~~

SAMPLE OF LOW SPEED ENCODER OUTPUT TO TAPE RECOR DER

2 min TAPE RECORDER PLAYBAC K

O U T P U T- AT 4 8 TIMES RECORD SPEEDOMMAND SIGNAL

u u

I- 90 mi n - f-, r 0 m in -1 P E R I O D( l l s a c )

2 0 0HIGH SPEED ENCODER HIGH SPEED ENCODERJ

TELEMETRY PROGRAM0 - 5

( v o l t s )

VOLTS INANALOG OSCILLATOR

F ~ G U R E-5. Ariel I telemetry program.

ARIEL I LAUNCH SEQUENCE

The sequence of events from nose cone ejec-tion to satellite separation is:

1. Nose cone ejection2. Second-stage burnout3 . Second and third-stage coast and yaw

until peak of ascent path is reached and vehicleis aligned with its programmed attitude

4 . Third-stage spin up to approximately160 rprn and ignition

5. Second-stage separation and retro-rockets fired

6. Burnout of third stage

7. Coast to allow outgassing ( thrust ) of thethird s tage to cease

8. De-spin of the third-stage-payloadcombination (first de-spin) to 76.5 rpm, byreleasing ('yo-yo" de-spin device

9. Release and erection of experimentbooms (second de-spin) and de-spin to 52.4rPm

10. Release and erection of the iner tiabooms and paddles, de-spin (third de-spin) to36.6 rpm

11. Separation of the payload from thethird stage at a differential velocity of 7 ft/sec

12. Final rate of spin at the end of 1 yearshould not be less than 12 rpm

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CHAPTER 4

United Kingdom ExperimentsI

LANGMUIR PROBE FOR MEASURE-MENT OF ELECTRON TEMPERATUREAND DENSITY

Project Scientists

Dr. R . L. F. Boyd, University College LondonDr. A. P. Willmore, University College London

Project Engineers

Dr. P. J . Bowen, University College LondonM r. J. Blades, Pye Ltd ., CambridgeM r. R . Nettalship, Pye Ltd ., Cambridge

The Experiment

This experiment, based on Druyvesteyns

modification of the Langmuir probe, deter-mines the local value of the electron density andtemperature near the satellite. A t heights wellabove the F-2 maximum, the ionization is prob-ably in diffusive equilibrium, so tha t the varia-tion of density wi th height can be related to thetemperature an d composition of the plasma.Moreover, it is very probable that temperatureequilibrium is closely approximated betweenthe electrons and th e positive ions. Thus, thetwo measurements made in this experiment arerelated. When they are combined with thosefrom the positive ion mass spectrometerprobe-which enables the positive ion tempera-ture and composition to be obtained, it shouldbe possible to study in a comprehensive way thedepartures from thermal equilibrium and diff u-sive equilibrium in the atmosphere.

Since the probe is negative with respect to i tslocal ambient atmosphere (potential of which,relative to the vehicle, will be called space po-

tent ial), the density of the electrons is obtainedby Boltzmanns relation as:

where nm s the density far from the vehicle, V,is the magnitude of the probe potential withrespect to space, and T, is the electron tem-perature.

Thus the current collected by the probe is:

where i o = ( 2 n k / m ) T,neoA A being the probearea).

In order to measure T , and &, from which ne,,

can also be calculated, a composite sweep waveform consisting of (1) a slow sawtooth wave,(2 ) a 500-cps sine wave whose amplitude isabout kT, le , and (3) a 3.5-kc sine wave also ofamplitude kTJe is applied to the probe.

Because the characteristic is exponential andnonlinear, the probe current contains not only500-cps and 3.5-kc components, but also har-monics and cross-modulation terms. Th e cross-modulation component, which is a 3.5-kc wavemodulated by 500 cps, is extracted by means ofa tuned amplifier; and the carrier and the mod-ulation are separated and measured by phase-sensitive rectifiers. The carrier component isused to derive an AGC feedback voltage thatmaintains the carrier amplitude a t the detectorconstant . Then the control voltage is a nearlylogarithmic function of the carrier amplitudein the probe current. Now i t is easy to see tha tthe carrier amplitude is a function of di p/dVp ndthat the rectified 500-cps voltage is R function

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of d2i,/d2Vp divided by di,/dV,. Th e la tt erquantity depends on eV,fkT,, where V , is theamplitude of the 500-cps modulat ing voltage,while the former gives i, (once T, is known).

The function of the slow sweep, which hasan amplitude of several volts, is simply to in-

sure that the probe is sometimes a volt or sonegative with respect to space. Space poten-tial can be recognized as being very nearly thepotential a t which d2i,/dVpf=O; and thus thevalue of V , corresponding to the measurementof i, is known. This gives io and hence ne,.

I 1I I !

I I

III

I I DENGYI! II S I G N A L

II

I L.1.--

1 SAWTOOTHMONITOR

FIGURE-1 . Electron density and temperature probe, block diagram.

A block diagram of the circuit is shown inFigure 4-1. The probe is a small disk, 2 centi-meters in diameter, surrounded by a guard ring.This has the dual purpose of reducing edgeeffects arising from the fact that the satelliteskin is not a t probe potential and of reducing

the st ra y capacitance from the probe to ground.(A similar probe constructed for Skylark isshown in Figure 4-2.) The tuned input trans-former of the amplifier is mounted behind theprobe. The waveform generators and ampli-fiers are mounted on a 5%-inch card in the inte-rior of the vehicle.

When direct transmission of telemetry datais being used, the modulation depth, carrieramplitude, and a sawtooth voltage are all trans-

mitted over the complete probe curve, togetherwith a monitor representing the amplitude VI.In this case, the complete probe curve can bereconstructed, thus enabling a cross-check ofelectron tempera ture to be obtained . The lowda ta ra te of the tape store makes it impossible

to record the probe curves in this way, so thecurves are sampled a t two points whose co-ordinates are recorded. The first chosen is thespace potential point where d2i,/dVP2=O. A trig-ger circuit produces a sampling pulse as the signof the second derivative reverses (and the phaseof the 500-cps wave changes by 180 degrees),and a gate stores the value of the sawtoothand the ca.rrier amplitude in capacitor stores.These values are read out by the telemeter.

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U N I T E D K I N G D O M E X P E R I M E N T S

FIGURE -2. Similar electron density probe con-structed for Skylark rocket.

Then, when di,ld V , reaches a predeterminedamplitude, which is se t a t the lowest valueconsistent with swamping amplifier noise, asecond sampling pulse is generated; and thevalues of di,/dV,, l and the modulationdepth are all stored and in due course readout. Thus, there is recorded the space poten-

tial, modulation depth or electron temperature,and two points on the first derivative curve.As there we in all four ionospheric sensors

being swept in potentittl, t h e effects of mutualinterference are being reduced by operatingthem synchronously. For this purpose, a trig-ger pulse is obtained from the encoder. Thecircuitry operates entirely from f6.5 volt lines,the total power dissipation being 75 milliwatts.

This experiment is duplicated in the satelliteto check the validity of the assumption thatthe ambient electron density is not greatly

affected by the motion of the vehicle. Bothprobes are mounted with the normal to theprobe surface parallel to the spin axis so as tominimize spin modulation of the probe current.One sensor is located inside the separation ring,facing backward from the launch direction, andthe other is on the end of the balance boomopposite to that carrying the dielectric experi-men t, facing forward.

SPHERICAL PROBE FOR MEASURE-MENT OF ION MASS COMPOSITIONAN D TEMPERATURE

Project Scientists

Dr. R . L. F. Boyd, University College London

Dr. A . P. Willmore, University ('ollege LondoiiProject Engineers

Dr. P. J. Bowen, University College LondonMr. J. Blades, Pye Ltd ., CambridgeM r. R . Nettalsliip, Pye Ltd., Cam br id p

The Experiment

The principle of this experiment is somewhatsimilar to the Langmuir probe. It can beshown tha t, if the cur rent to a probe immersed

in an isotropic plasma is i, a t potential V , andthe energy distribution function of the chargedparticles is j ( E ) , hen

Thus, by applying a slow sweep voltage V ,and by applying two sine waves of differingfrequencies in the manner of the electronenergy probe to obtain d'i,/dV,2, the energydistribution function f ( E ) can be constructed.

If V,, is positive, this distribution functionwill be that of the positive ions. However,the positive ion cur rent is, in practice, swampedby the electron current ; and the probe must becovered with a grid a t a negative potential sothat these are biased off. If a spherical probeis used, it is no longer necessary th at the distri-bution function be isotropic. In t his case i t iseasy to show from considerations of the con-servation of energy and angular momentumth at the presence of the grid will not affect themeasurements. Thus, with such a probe the

energy distribution of the positive ions can bedetermined.

If the temperature of the positive ions werevery low, they would appear a t the satell ite asa homogeneous stream moving with the satellitespeed vs . Thus, they would have an energyE= 1 / 2 r n ~ , ~ , hus, the energy distributionfunct ion consists of peaks corresponding to themass spectrum of t he ions. In fact, the thermal

17


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ARIEL I : TH E F I R S T INT ERN ATI ONA L SATELLITE

velocities are not negligibly small; and, inconsequence, the peaks are broadened to aneasily measurable extent. Th e result is todegrade the mass resolution but to enablemaking a measurement of the ion temperature.

Th e basic problem of this experiment, then ,is identical with tha t of the electron tempera-ture measurement, namely, the determinationof d2it/dV$. Th e principal difference is inthe construction of the probe, which is shownin Figure 4-3. Th e inner electrode is the probe

OUTER SPHERICAL GRID INNER SPHERICALP R O B E\

FIGURE -4. Mass spectrometer probe as flown on

Black Knight rocket.

HEAD

- 1 inch LJ

being different. The shape of the energydistribution curve is transmitted by direct,telemetry but is not recorded,

The probe is mounted on the spin axis, toreduce spin modulation of the probe current.I t is carried on a thin tubular support , as farfrom the ve,hicle as possible-namely, some 10centimeters in front of the cosmic ra y packageon which it is supported.

MEASUREMENT OF SOLAR LYMAN-ALPHA EMISSION

FIQURE -3. Sectional view of mass spectrometerprobe.

proper, a sphere 9 centimeters in diameter.This is surrounded by a concentric sphericalgrid of nickel foil 0.1 millimeter in thickness,pierced with 0.5-millimeter holes. The diam-eter of the grid is 10 centimeters. The grid

both serves to repel electrons and acts as anelectrostatic shield for the inner electrode.Figure 4-4 is a photograph of this probe asflown on the Black Knigh t rocket. By makingthe sphere rather large in comparison with theLangmuir probe, it is possible to obtain probecurrents that are not greatly different in thetwo cases; and in consequence the same cir-cuitry serves for both with only slight modifi-cations, the amplitudes of the sine waves andthe inpu t impedance of the tuned amplifier

Project Scientists

Mr. J. A. Bowles, University College LondonDr. A . P. Willmore. University College London

The Experiment

Th e function of t,he two solar radiat ion ex-periments is to enable making simultaneous

and nearly continuous observations of t he st a teof the ionosphere and of the solar atmosphere,so as to investigate in detail this aspect of th esolar-terrestrial relations. T o this end, twoparts of the solar spectrum have been selected:one where the radiation originates at a lowalt itude in the chromosphere, and the otherin the corona. These are, respectively, th eLyman-alpha line of hydrogen and the ratherhard x-ray spectrum. The lat ter radiation isalso emitted from the high-temperature dis-

18


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UNITED KINGDOM EXPERIMENTS

turbed regions of the corona. The densitymeasurement is made by means of the nitricoxide ionization chambers designed by Dr. H.Friedmann of the U.S. Naval Research Labora-tory, Washington, D.C . These chambers aresensitive in t.he wavelength region 1100 to1350A, where nearly all the energy in the solarspectrum is concentrated in the Lyman-alphaline. The chambers are mounted, as shownin Figure 4-5, on a molded nylon support cone

i ! \ ' )L - - . O i

FIGURE -5. Mounting of Lyman-alpha detectors.

t ha t restrict s t he field of view to 60 degrees.Three such chambers are used, mounted atangles of 30, 90, and 150 degrees to the spinaxis so that the array covers the whole sky-.one in each revolution of the satellite, withnearly constant sensitivity.

The ionization current produced by solarradiation is approximately 5X ampere.

This current is amplified by a transistor chopperamplifier, and the peak value of t he current ineach revolution is stored in a capacitor havinga storage charge time constant of 200 seconds.This store is sampled by bot,h low and highspeed encoders. Th e electronic module cardshown in Figure 4-6 operates from -6.5 volt s,1.5 milliamperes, and a bias line of - 4 volts.

FIGURE -6. Electronic card for Lyman-alpha emis-sion experiment.

MEASUREMENT OF THE X-RAY EMIS-SION FROM T HE SUN IN TH E 3 T O 12A

B A N D Project Scientists

Dr. R . L. F. Boyd, University College LondonDr. I(. A. Pounds, University of LeicesterDr. A. P. Willmore, University College London

Project Engineers

Mr. J. Ackroyd, Bristol Aircraft CompanyDr. P. J. Bowen, University College LondonMr. P. Walker, Bristol Aircraft Company

The Experiment

The wavelength region selected for thisobservation is 3 to 12A, in which the solaremission is highly variable and is a sensitiveindication of solar conditions. The circui tryhas been designed, for this reason, to have a widedynamic range. The radiation detectors areproportional counters, filled with argon andmethane as a quenching gas, and using 2 5 ~beryllium windows, which have a high quttn turn

19


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A R I E L 1: T H E F I R S T I N T E R N A T I O N A L S A T E L L I T E

efficiency in the 3 to 12A region. Each counterhas three windows, each 0.5 millimeter indiameter. The windows are spaced at 30degree intervals round the circumference, givingthe counter a field of view of 90 degrees in theplane normal to its axis, without the radiationangle of incidence exceeding 15 degrees. Thiscondition of nearly normal incidence is necessarybecause, near the cutoff a t 13A, the transmissionof the window vanes rapidly with angle of inci-dence. Two such counters are used, withmatched characteristics, mounted at "lati-tudes" (reckoned as though the spin axis of thesatellite were comparable with that of theearth) of 5 4 5 degrees, so that the whole skyis covered once in each revolution. Eachcounter is fitted with a mask that restricts thefield of view in the longitudinal direction, so

that the two counters together have a field ofview which is a sector of 30-degree includedangle.

A block diagram of the circuitry is shown inFigure 4-7. The counter is supplied from an

extra high tension (EHT) generator, which isa dc to dc converter operating from -6.5 voltsand producing 1600 volts; this is stabilized bya corona discharge tube. The counter pulsesare fed by a gain-stabilized linear amplifier to adiscriniinator with a variable bias level. Thodiscriniinator output passes through a gate to a15-stage binary counter capable of countingequally spaced pulses at 1 megacycle.

The gate circuit is operated by a 1-secondpulse from the high speed encoder, so timed tha tthe scaler read-out always occurs in the "off"period of the gate. Closing the gate also oper-ates a staircase generator, which produces awaveform of five equal steps used to obtain thediscriminator bias level. Thus , the discrim-inator automatically moves on at the end ofeach counting period to the next wavelength

interval, until the range from 3 to 12A has beencovered in five steps of equal quantum energyinterval, after which the process is repeated.

In general, the sun is expected to be in thefield of view for a time rather shorter than 1.

( X - R AY BOARD2 ). . .

OUTPUT TO OUTPUTroMIGH SPEED ENCODER HIGH SPEED ENCODER

LOW SPEEDENCODER

PUL SE SEOUENCE DIAGRAM

+ 6 7 rCOYMANO PULSE

ALL 0

0"

( - 2 7 sGENERATED

PULSE- APPROX I+

, 1C I

FIGURE -7. Block diagram of x-ray spectrometer circuit.

20


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second, so that the effective counting time isless than 1 second. The actual time will bedetermined from a spin rate measurement.Moreover, there is only a 1:12 chance tha t thesun will be in the field of view in any given gat eperiod. Thus, the count,ers do not only observethe sun but also the background radiation fromthe rest of the sky. On an average, 5 x 1 2 - o r60-telemeter periods of 5.12 seconds will berequired to obtain a complete solar spectrum.

The high speed encoder is used to sample thebinary out put s of all 15 scalers. For the lowspeed encoder the necessary information capac-it y is not available, so the output s of the centraleight scalers are combined to give an analogvoltage representing the logarithm of the storedcount to base 2 , to one significant digit only,within the range covered by the scalers. I n

addition, t,he discriminator bias level is also

mmm---

FIQURE -8. X-ray linear amplifier.

FIGURE -9. X-ray high speed counter.

FIQURE -10. X-ray low speed counter.

FIQURE -11. X-ray discriminator circuit.

telemetered and , on the high speed encoder only,an EHT monitor.

Figures 4-8 through 4-11 are photographs ofthe main electronic module elements. Theseoperate from +6.5 volt and -6.5 volt lines,with a power consumption of 130 milliwatts.

MEASUREMENT OF SOLAR ASPECTProject Engineers

Mr. J . Alexander, University College LondonDr. P. J . Bowen, University College London

The Experiment

Both radiation experiments require informa-tion on th e solar aspect, as their sensitivity is nfunction of aspect angle and spin r at re ; and it

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ARIEL I : THE FIRST IN TE RN AT IO NAL SATELLITE

was considered that such information wouldalso provide information on the spin rate an dthe stabili ty of the spin axis which would rtlsobe valuable in the interpre tation of the iono-spheric experiments. In all, i t was consideredth at the necessary information consisted of thesolar aspect angle-that is, the lat itude of thesun in the satell ite system-and of the roll posi-tion of t he sun or its longitude measured froma reference plane in the vehicle. In fact, thetime intervals between telemetry samples ofroll position are such as to make possible anambiguity about the number of complete rollsoccurring between measurements: so a spin rateindication has also been included.

The basic method of measurement utilizes asmall pyramid of silicon solar cells a t th e centerof a 9-centimeter-diameter hemisphere. Th e

cells are illuminated by the sun through twoslits whose shape is defined by the intersectionwith t he hemisphere of a circular cylinder 4 %centimeters in diameter, the axis being parallelto the spin axis. The arrangement is shown inFigure 4-12. The cells are covered by an

F I G U R E-12. Satellite aspect slit system.

internal mask so that one cell on each side ofthe satellite equator is illuminated from oneslit, and the outputs from the four cells arecombined so that those from the two slits areof opposite polar ity. Th e slit width and cellsize are on the order of 1 degree, so that in eachrevolution of the vehicle one short positivepulse and one short negative pulse a re produced.The time interval between positive pulses isthen equal to the spin period; the actual time

of their occurrence gives the roll position; andthe ratio of the time interval between positiveand negative pulses to the interval betweenposit ive pulses is a linear function of th e aspectangle.

The positive-going pulses are used to supplya simple pulse-rate circuit covering the range180 rpm to 20 rpm. This will give an indica-tion of the proper operation of the de-spin de-vice and, at the same time, removes the rollposition uncertainty. Th e positive pulses alsooperate a phase-locked sawtooth generator,which produces a triangular wave of constantamplitude whose commencement is always veryclose to th at of the arriva l of the pulse. Thus,this wave represents at any instant the rollposition of the sun and is sampled by the tele-meter. Finally, the positive pulses are used toturn o n , and the negative pulses to turn o$,a bi-stable circuit-thus producing a squarewave of cons tant ampl itude whose mean valueis proportional to the solar aspect angle. Thesquare wave is passed to a simple averagingcircuit whose output is telemetered. Thecircuitry is supplied with 10 milliwatts from+6.5 volt and -6.5 volt lines.

COSMIC RAY ANALYZER

Project Scienti sts

Prof. H. Elliot, Imperial College, LondonDr. J. J. Quenby, Imper ial College, LondonMr . A. C. Durney, Imperial College, London

Project Engineer

Mr. D. W. Mayne, McMichael Radio, Ltd.

T he Experiment

Experimental ObjectivesTh e main purpose of this experiment is to

make accurate measurements of the primarycosmic ray energy spectrum and of the way inwhich this spectrum changes as a result ofmodulation by the interplanetary magneticfield. There are at present several alternativemodels of this field; and, to distinguish betweenthe various possibilities, much more refinedda ta are necessary than are now available fromthe observation of secondary cosmic ra y in-tensity variations deep in the atmosphere andfragmentary information from balloon ascents.

22


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In the present experiment it is proposed toinvestigate the cosmic ray spectrum by using aCerenkov detector to measure the intensity ofheavy nuclei ( 2 2 6 ) as a function of lat itude.The energy spectrum can then be determinedfrom the known values of the minimum energythat a particle must have in order to penetratethe earths magnetic field at a given latitude.The advantages of measurements carried outin this way are: (1) It avoids the invalidation ofcosmic ray measurements by inadvertent de-tection of Van Allen part icles, and (2 ) it avoidsthe int roduction of uncertaint ies by albedoparticles scattered back from the atmosphere.

The detector will sweep through the cosmicray energy spectrum four times on each orbit,giving a virtually continuous check on itsvariation with time. I n addition, airplane

surveys are to be carried out a t the same timeso that the intensity distribution in the atmos-phere can be uniquely related to the primaryspectrum. It is hoped by this means to obtaina sufficiently accurate relation so that theprimary spectrum can be determined at anyfuture time simply by an airplane survey.

A large and a small Geiger counter will beincluded in the inst rument package. On thosep a r k of the orbit where the satellite is outsidethe Van Allen belts the large counter will givea measurement of t he energy spectrum of t he

primary cosmic ray protons, and also burstsof solar protons, in the same way as the Cer-enkov detector measures the heavy particlespectrum. Thus, it is hoped tha t the protonand heavy particle spectra can be compared asa function of time. Inside the Van Allen beltsthe large Geiger counter will saturate, but thesmall Geiger counter will be able to count amuch higher particle intensity. It is hopedto obtain dat a on the time variations of thetrapped radiation from the small counter rate.

Operation of the Sensor

The Cerenkov detector consists of a hollowperspex sphere, 4 inches in diameter, togetherwith an E M 0 6097 photomultiplier 2 inches indiameter, which looks into a hole cut in thesurface of the sphere. Cerenkov light flashesproduced by cosmic rays in the wall of thesphere are detected by the photomultiplier.


23

The pulse output of the tube is fed to a dis-criminator t ha t only accepts pulses correspond-ing to primary nuclei of >> 6 or Z >> 8,rejecting smaller background pulses due tolighter nuclei and the Van Allen particles.Out put from the discriminator is fed to a chainof eight binaries, and the contents of this storeare sampled by both the da ta storage and directtelemetry encoders.

An electronic switch that controls the atten-uation of the photomultiplier ou tput pulses iscoupled to the last binary. There are two levelsof attenuation; and, each time the store is com-pletely Alled, a pulse is fed into the switch fromthe last binary, and the switch changes theattenuation to the other value. I n this way,the two (effective) levels of th e disc riminatorare obtained.

Th e complete instrument package is locatedon the spin axis a t the forward end of thesatellite, with the perspex sphere protrudingoutside the satellite surface and the photo-multiplier and associated electronics just inside.Primary heavy nuclei can arrive at the spherefrom the forward 2r solid angle, where theamount of shielding is in general less tha n 1gm/cm2, but are absorbed by the sate llit estructure in the backward 2 r solid angle. I t isadvantageous to locate the Cerenkov detectoron the spin axis, since the direction in which itsees cosmic rays changes only slowly in timewith respect to the earths surface and the geo-magnetic field. Th e effect on the cosmic ra yintensity of both absorption by the earth anddeflection by the geomagnetic field must betaken into account in the analysis of the results.

Th e Geiger counters are placed in th e instru-ment package with structure shielding on theorder of 2 gm/cm2 from the outside in theforward direction. The effective length anddiameters of the counters are, respectively, 5and 2 centimeters for the larger and 1 and 0.3centimeters for the smaller. An additional 1millimeter of lead is put around the smallcounter.

Both counters feed in parallel a chain of 13binaries, and the contents of this store are readperiodically by both the dat a storage and directtelemetry encoders. If the counting rate ofthe large Geiger tube exceeds capacity of the


30/76

ARIEL I : THE FIRST INTERNATlONAL SATELLITE

binary store, an integrating circuit turns it offwith the result that only the s m d e r counterremains operative-so increasing th e dynamicrange of the system.

S e ns or S t r u c t u r e

A diagram of t he sensor is shown in FigureThe perspex sphere is contained in an-13.

Iir I +I

AXIS ISPIN 5 " d i m .

I%;' diam.1- ' - T I 1

GEIGER COUNTERS ELECTRONICSPOlTED IN ECCOFOAM

FIGURE-13. Cosmic ray analyzer sensor diagram.

aluminium dome, 5 inches in diameter and 4K6inches high, protruding outside the satelliteski n, to which the s talk of the UniversityCollege probe is attached a t the top. Ecco-foam serves as a packing between the sphereand the aluminium container. Th e sphere

temperature must be kept below 70' C, thesoftening point of perspex. A light -tight airleak is provided between the in terior of thesphere and the outside surroundings.

An aluminium cylinder 6>< nches in diameterand 5 inches deep is attached under the domeinside the satellite and contains the photo-multiplier, Geiger counters, and electronics.The photomultiplier is placed along the spin

axis and looks into a 2-inch-diameter hole cu tinto the wall of the sphere. Some of the elec-tronics, together with the two Geiger counters,are located in an annular space around thephotomultiplier; and this whole is potted inEccofoam and placed in a brass shieldingcanister. The canister is floated in a foammaterial to reduce the vibration acting on thephotomultiplier. Th e remainder of the elec-tronics is situated in a further ring surroundingthe floating canister, and this also is potted inEccofoam. Photographs of the sensor tireshown in Figures 4-14 and 4-15.

Th e weight of the potted sensor, includingthe base of the University ('ollege Londonmass spectrometer, is 5.3 pounds. The centerof gravity is on the spin axis, 3 f i nches up fromthe lowest surface on the base of the electronics

cylinder. Th e moment of inertia abo ut thespin axis is 0.2 lb f t2 ; nd, a bou t an axis throughthe center of grav ity an d perpendicular to thespin axis, the mass of inertia is 0.09 lb f t z .

Sensor Electronics

A block diagram of the electronic circuitryis shown in Figure 4-16. Two input powersupply lines are required: 200 milliwatts a t- 9 5 1 volts and 100 milliwatts a t - 6 . 5 5 1volts. The -9 volt line is fed to a dc con verte r,producing a very stable -6 volt line th at is

used to supply the E H T converter, the em itterfollower, the discriminator, and the gatecircuits.

The dc converter incorporating a coronastabilizer supplies the 60-megohm dynode chainof the photomultiplie r with -1000 volts, andthis E H T voltage was designed to have a5 4 volt variation for a f volt variation on t he- 9 volt input supply. The att enu ato r isswitched to permit measurement of ei ther t hecosmic ray rate from ;5>6 or from Z B 8 .

The discriminator output goes via R gatecircuit to the chain of eight binaries. Figure4-17 shows the circuit of a single binary,together with the size of the ou tp ut waveform as a function of supply voltage. Th eoutputs of the first six binaries are fed to theinp ut gates of the high speed encoder, whilethe ou tputs of the las t six binaries are fed tothe input gates of the low speed encoder.

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FIGURE -14. Cosmic ray sensor and electronics.

When a particular information channel is to besampled, the g ate circuits feed three of thebinary outputs to a digital oscillator that takesup one of eight different frequency levels,depending on the s ta te of the three binaries.This frequency output is then put on the taperecorder in the case of the low speed encoder,and goes to modulate the telemetry trans-mit ter in t he case of the high speed encoder.Thus, to read the binary store, two channelsare required on both the data storage and t h edirect telemetry systems.

While the low speed encoder is sampling thebinary store, i t is arranged th at the gate circuit

FIG URE -15. Mass spectrometer probe and cosmicray experiment assembly.

>mc

w

w

FROM ENCODER

-DIGITAL OSCILLATOR

FIGURE -16. Block diagram of cosmic ray analyzer circuit.

25600-439 0-43-3


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BINARY UNIT

- - 6 % O I P

I I P - IU

placed after the discriminator prevents anyout put reaching the first binary. This is donebecause the state of the binaries may changeduring the time taken to sample both binarysets. Th e gate circuit is driven by a pulse fromthe low speed encoder, which has a quiescentlevel of +6.7 vol ts and goes to -2.7 volts dur-ing the period when the st at e of the two set s ofbinaries is read by the digital oscillator. Zerocurrent is drawn from the encoder at $6.7 volts,and lop amp at -2.7 volts.

An electronic switch working from the outputof the last binary will change the value of the

two-position attenuator and successively feedthe two cosmic ray counting rates into thetelemetry systems. When the binary store isfilled, the flipover of the last binary changes thest at e of a switching binary, the output of whichdetermines the value of the attenuat ion. Totell which attenuation value is being used, theswitching binary outpu t is also fed t o both thelow and high speed encoders as one of the threebinary elements in information channel C4, the

other two elements being taken from the Geigercounter store.

Appropriate voltages for the Geiger countersare tapped off the dynode chain. The anode-to-cathode voltage of the larger counter is madeto fall below the working voltage a t counting

rates, in excess of that which the data encodercan handle. Pulse outputs from the twocounters are mixed and fed via a gate circuit toa chain of 13 binaries. Numbers 2 to 12 areread by the high speed encoder, and numbers3 to 13 are read by the low speed encoder. Th edesign of the binaries and the method of sam-pling by the encoders are similar to the case ofthe Cerenkov channel. Thus , four channels ofboth the data storage and the direct telemetrysystems are required. The gate circuit stopsextra pulses being fed to the store, while thelow speed encoder samples the four sets ofbinaries and works from similar pulses to theCerenkov channel gate.

The digital oscillators and the associatedinpu t gat e circuits for both encoders are locatedon a card held below the sensor by extensions tothe bolts that locate the sensor on the satelliteskin.

Plugs dlzd Connec t ions

About twenty-six leads are required betweenthe gate and oscillator circuitry in the sensorand the encoder modules. Two power leads at-6.5 and -9 volts are also required. Abouteight further output leads are necessary if th esensor performance during environmental test-ing is to be monitored independently of th etelemetry system.

Thermal Design

requested:The following temperature restrictions are

1. Perspex sphere, 7OoC (this is the soften-

ing temperature of the plastic);2. Electronics, O o to 4OoC;3. Photomultiplier, 0 to 30C (if possible).

1. Th e outer shell of the unit is made of16 swg aluminium alloy (BS1470) andweights an estimated 16 ounces; it hasa surface finish similar to duralumin.

2. The inner shielding canister is made ofbrass and weights about 16 ounces.

Relevant information about materials:

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3. An estimated 23 ounces of Eccofoam isused in potting.

4 . The remaining 2.3 pounds of sensorweight is made up of electronic com-ponents, circuit boards, photomulti-plier, Geiger counters, etc.

P L A S M A D I E L E C T R I C C O N S TA N TMEASUREMENT OF IONOSPHERICELECTRON DENSITY

Project Scientist

Prof. J. Sayers, Universi ty of BirminghamProject Engineer

Mr. J . Wager, University of BirminghamThe Experiment

The measurement of electron densi ty is per-formed both by this experiment and by the

Langmuir probes of Universi ty College London.However, the two methods are quite differentin principle and, being subject to differenterrors in general, therefore complement oneanother. Two circular disks of wire mesh, 3 .5inches in diameter and separated by 3. 5 inches,form a parallel-plate capacitor (mesh is used inorder to allow free passage of electrons into t hecapacitor). The Capacitor is a t the outer endof a boom 49.5 inches long, so th at the electrondensity will not be greatly affected by thepresence of the satel lite .

The capacitor forms one arm of an R Fcapacitance bridge, operating a t 10 megacycles.From a knowledge of the capacitance at 10megacycles, and of the capacitance in vacuo,the dielectric constant k of the ionosphericplasma can be found; this is related to theelectron density by t he expression

&I-- Nee2

where N e is the electron density, e and m arerespectively the electronic charge and mass,and w is the angular frequency of operation ofthe bridge (here 2 ~ x o- rad/sec). However,N e n the equation will be equal to the ambientelectron densi ty only if the potential of thecapacitor plates is the same as the space po-tential. In general, there will be a potentialdifference between the satellit e and space. Toinsure that at some times the capacitor plates

7rm.02

27

are at space potential, a sawtooth wave withan amplitude of about 4 volts is applied tothem; N e is then taken to be the maximumvalue obtained during the sweep. Immediatelybefore the st ar t of each sweep a rather largenegative potential ( - 5 volts) is applied to the

plates to denude them of electrons and to pro-vide a check on the in-vacuo capacitance.Block diagrams of the circuitry are shown in

Figure 4-18. To avoid a possible interactionwith the University College London probe ex-periments, the experiment is turned on onlyfor 10 seconds in each 61.42-second period.This is controlled by trigger pulses from thelow speed encoder binary sequence. A gatingunit then supplies power t o the VF bridgecircuitry and the bridge potential waveformgenerator-all of which are located a t the hinge

end of the boom-for thi s period. The outputfrom the bridge is amplified, separated intoin-phase and quadrative components (corre-sponding to the real and imaginary par ts of thedielectric constant, respectively), and tele-metered through the high speed encoder. Also,the maximum values in each sweep are selectedby peak reading circuits and are stored incapacitors until they are read out by the lowspeed encoder.

The temperature of the boom electronics,the voltage supply levels, and the variablefrequency (VF) amplitude are also measuredand telemetered by both encoders.

DUT CHM AN EXPERIMENTS

Project Scientist

Dr. A . P. Willmore, University College LondonProject Engineer

C. L. Wagner, Jr., Goddard Space Flight CenterThe Experiment

Th e payload-supporting Dutchman containstwo experiments directly connected with thespacecraft instrumentation. Th e first of theseis designed to measure solar aspect and spin rateduring the period before separation occurs:while the second is designed to detect contami-nation of the spacecraft thermal coatings andsensors as a result , for example, of evaporationfrom the nose fairing following aerodynamicheating. Each of these experiments receives


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E TRIGGER F R O M E N C M E R

GATING UNITIMLIVERSIO sec ELECTRONICS

M G t l SPEED DATA

READINGCIRCUITS

CAPACITORSTORES

LO W SPEED D7

FIGURE -18. Electron density circuit, block diagram.

power from the regulated + 6 volt supply pro-vided for the vibration experiment, and thesignals from the two are combined for trans-mission over a single telemetry channel. I n

addition, an indication of the operation of athird-stage pressure switch is provided on thischannel.

The aspect sensor is of t he same pat tern asthat used in the spacecraft, but the aspectcomputer has been omitted. Th e pulse sequenbefrom the solar cells is telemetered directly. Th eadvantage of th is procedure is tha t the ratherlong time delays associated with the aspectcomputer are eliminated, thus enabling third-stage spin and nutation to be calculated.

A diagram of th e contamination experimentis shown in Figure 4-19. A small, focused lampis used with a filter to produce an app roximatelyparallel beam of suitab ly colored light. A smallsample of this beam is obtained by reflectionfrom an unsilvered glass surface, and its inten-sity is measured by means of a photodiode.

FIGURE -19. Contamination experiment.

Th e remain der of the light is incident on anotherglass slide, the intensity of the reflected lightbeing monitored by means of a second photo-diode. The second glass slide is exposed tocontamination by the fairing, such contamina-tion being revealed by the change in reflectivityof t he glass.

A simple gating circuit is used to switchbetween the reference and sample slides a t ara te of 1 cps. Th e output from the photodiodesis superimposed on the photocell pulses of th easpect sensor.

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CHAPTER 5

United States SatelliteandSubsystems

STRUCTURE AND MECHANICALDESIGN

Introduction

The Ariel I structure and mechanisms designencompasses two separate major areas of en-deavor: (1) The satellite itself requires struc-ture and mechanisms; and (2) just as importantto the tot al payload is the development of aseparation and release system interface struc-ture and associated mechanisms.

The satellite itself relies heavily on the useof epoxy-bonded filaments of fiber glass formuch of its structure, used in combinationwith machined wrought aluminum alloys.The ancillary interface system (Dutchman andseparation structures) is largely composed ofmagnesium thin-wall castings.

Satellite Structure

Genera l

The spacecraft structure should properly bedivided into two main groups: the basic struc-ture, and the appendages. The basic structuremust then be further divided into the followingsubcomponents:

1. Upper dome,

2. Mid-skin,3. Shelf and base assembly,4. Lower dome.

The appendages are as follows:1.2.3.4 .5.

- _

Four solar paddles,Two inert ia booms,Electron density experiment boom,Electron temperature experiment boom,Telemetry antennas.

29

Design Parameters

Several design considerations were paramountin the development of the structure:

The Scout 25.7-inch-diameter heat shieldlimited the payload size to 23 inches in diameterby 2 feet in length, not including certain experi-ments.

Th e af t appendages must be folded i nto thespace described by a hollow cylinder 23 inchO.D. x 18 inch 1.D. x 4 feet 8 inches long.

The structure should be manufactured ofnon-metals and/or nonmagnetic materials, soth at the effects of magnetic spin dampingare reduced to a degree that the half-life ofthe satellites spin ra te be 1 year.

The total satellite weight should not ex-ceed 135 pounds.

The st ructur e must withstand accelerationsand vibrations of the launch vehicle. In thiscase the ABL-X248-A5 motor (last stage)governs.

Ou te r S t ruc tu re

General: As a resul t of the above considera-tions the material chosen for the skin of theAriel 1 was epoxy-bonded fiber glass. Thedomes were constructed from monofilamentglass fibers cross-woven into cloth laminationsthat were molded into a spherical shell of 13%-inch radius by 5% nches high. Shell Epon 828with a CL hardener was used as the bondingagent. The top dome is X s inch thick, andthe bot tom dome-basically used only as athermal shield-is x 2 inch thick. The mid-skin was made from a cylinder of epon-bonded,


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monofilament-wound fiber glass K s inch thick x to the upper structure and the shelf-base23 inches in diameter x 10.7 inches long. assembly. In detail, shear lips prevent radial

movement, pins prevent rotational displace-ment, and machine screws tie the componentstogether.

Lower Dome (Figure 5-la) : The lower domeis segmented and fitted with doublers for instal -lation of sensors and the segments themselves.Gold-plated aluminum machine screws holdthese components in position.

Shelf and Base Assembly (F igure 5-2)

Upper Dome (Figure 5-1): Int o the top of the

General: The instrument shelf and base

l b l

(a) Uncoated upper and lower domes and mid-skin(b) Interior view of upper dome

FIGURE-1. Domes and mid-skin.

upper dome skin was bonded and riveted analuminum disk 8% inches in diameter x 0.2 inchthick. Machined integrally with, and cen-trally located on, the disk is a 7-inch-I.D. thin-walled cylinder extending internally 3.7 inches.This hat-shaped structure supports the cosmicray-ion mass sphere experiments and the eightradial fiber glass ribs, which are also attachedto the dome skin-thus giving stiffness andstrength to the dome. A machined aluminumring was bonded and riveted to the base of th edome to allow the assembly to be bolted to thetop of the 23-inch-diameter mid-skin. Holeswere cut in the proper places of the dome toallow attachment of experiments and antennamounts.

Mid-Skin (Figure 5-la) : The mid-skin fiberglass is bonded and riveted to two end flanges,machined from AISI 6061-T6 aluminum, whichare shaped to provide nonshifting attachment

assembly is made from AISI 6061-T6 aluminumplates, bars, and billets machined into shapeand semipermanent ly affixed in to t he assemblycondition.

Shelf: The shelf is basically a 21-inch-diameter plate 0.08 inch thick with eight inte-gral stiffening ribs leading radially from a7%-inch-diameter integral cylinder to the outerperiphery. This undercarriage of ribbing tapersfrom 0.7 inch a t the cylinder to 0 .3 inch at the21-inch-diameter periphery. The top of theshelf a t the 21-inch diameter is dished upwards1 inch and then again extends radially until theshelf is 23 inches in diameter. This step ismachined to supply room for the de-spin systemand to provide a mating surface for the mid-skin.

Base: The base is a cylinder of 7 inches I.D .and fits into the 7%-inch-diameter short-shelfcylinder, extending aft 6% nches. The aft 2inches is machined to an 8g-inch I.D. to provideproper view angles for the spin-axis-mountedelectron temperature sensor. The forward 7-inch-diameter x 4 %-inch-long space is providedfor containing the above sensor, the satellitestape recorder, and the boom escapement mech-anism. The outer diameter is machined toprovide (1) a mounting surface for the separa-tion adapter ring, (2) a mounting surface for th ebottom dome segments, and (3) a key for sixsupport struts.

Struts: Th e six st ru ts are each machined fromsolid stock and take the shape of a modified I beam. Each strut serves to support theshelf and to supply the mount for a paddle armor experiment boom hinge. The st ru ts arekeyed to the base and, after being fastened intoposition with machine screws, are keyed to theshelf by shrink-fit shear pins.

The separate parts are a s follows.

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UNITED STATES SATELLITE AND SUBSYSTEMS

l b )(a 1

(a) Viewed from above (b ) Viewed from belowFIGURE -2. Electronic shelf and base assembly.

AppendagesGeneral (Figure 5-3): The eight appendages

should be considered in three separate group-

FIGURE-3. Satell ite appendages.

ings: paddle arms; inertia arms; and sensor, orexperiment, booms.

Paddle Arms: The paddle arm and hingedesign was suggested by that used on ExplorerXI1 (1961 V I ) ; however, space considerationsdictated by the Scout heat shield and paddlelocation restrictions required.by t he experimentscomplicated the design considerably. Th e armsthemselves are long slender channels machinedfrom AIS1 7075 aluminum. One pair of armslead directly from their hinges to the paddleinterface; but space and positional requirementsfor the other pair of solar paddles required asecondary folding hinge a t the outboard end ofeach arm. From this secondary hinge extendsthe paddle interface of these two arms.

Inertia Booms: The inertia booms, when ex-tended , provide a proper moment of inertiaratio so that the longitudinal axis remains the

spin axis of the spacecra ft. These booms aremade of thin-walled tubes of epon-bonded fiberglass cloth, rolled into cylindrical shape. Eac hboom is attached to the shelf by a detent-locking, spring-loaded hinge; and at the out-board end of its 30 inches is a 0.7-pound steelweight that is 4 inches long. Both the inertiabooms and th e paddles are erected a t 52.4 rpm,reducing spin to 36.6 rpm.

Experiment Booms: The sensor booms aremade by the experimenters, but the method oferection is part of the structures responsibility.The hinge valves are machined from solid stockand use a double detent lock and a torsion-spring positive force to assure opening in theevent of no payload spin-up. (This consider-ation is true in all appendage extension.)Escapement (Figure 5-4)

Boom erection rotat ion speed of the normalexperiment is 76.5 rpm, which would producelarge shocks to the experiments if these boomswere allowed to open without restraint. AS aresult, an escapement device was designed toreduce these forces by controlling erectionspeed.

The clock escapement principle is employedin the design of the timing mechanism. Sincethe primary interest in the application is toreduce shock-and not the accuracy or timing,the less sophisticated mechanism, called a run-away-escapement, is used. It consists of agear train coupled to an escape wheel, and apallet fitted over the escape wheel to controlits rotation rate. The two booms are required

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FIGURE -4. Escapement mechanism.

to be simultaneously controlled; thus a doublepulley mounted on a common shaft is coupled tothe free end of the gear train. A doubled nyloncord, with one end attached to the boom andthe other end to the pulley, serves as the controllinks. The rota tion ra te of the escapement is a

function of the satell ite acting torque and themoment of inertia of the par t. The palletsmoment of inertia is adjusted to allow the boomsto erect in 2 to 3 seconds. Thi s escapementwill successfully control the boom erection whenthe satellite spin rate is between 60 and 90 rpm.

The booms are also tied together by thismechanism, so that they erect simultaneouslyand thus prevent any unbalance that wouldcause coning of the payload. (The other ap-pendages open too quickly to create thisproblem.

De-Spin Device

The de-spin system is bui lt in a self-containedring fitting in the space provided a t the periph-ery of the shelf. It is of the st ret ch yo-yodesign (Reference 13), which will de-spin a sys-tem having a f 2 0 percent nominal spin rateerror to f 2 percent of t he required final value.The basic components are a pair of equalweights attached to a matched pair of long

tension springs wound one-half turn about thepayload. The weights were released by pyro-technique guillotine cutters, and the weight-spring combinations unhook themselves whenin a radical atti tude from the spacecraft. Bythis system the payload is despun from 160 to

76.5 rpm.Alztenrtds (Figure 5-5)

The four turnstile-type antennas are of thedouble-fold design, so that they can fit into the

FIGURE -5 . Flight antennas.

hea t shield. These antennas are located on thetop dome equispaced 90 degrees apart, and intheir erected position make an angle of 40degrees with the spin axis. Upon ejection of theheat shield, the antennas are erected to theirlength of 21Yi inches.

Battery Containers

The two containers for the spacecraft batterypower supply were designed to withstand theforce of expanding gas generated interna lly inthe sealed batteries, which could result from apartial power system malfunction. The con-tainer can withstand loadings of 50 pounds fromeach of two st,acks of bat ter ies without signifi-can t expansion-thus preventing rup tur e of thebatteries and subsequent contamination of thepayload. These containers are mounted on theinstrument shelf and supply rigidity in the areaof the iner tia boom hinge.

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UNITED STATES SATELLITE AND SUBSYSTEMS

Miscellany

Th e above-mentioned components plus awealth of miscellaneous supports, adapters, andbracketry make up the structures and mecha-nisms of the orbiting spacecraft, accounting for51.6 pounds of it s weight.

Separation and Release Syste ms

Although not a par t of the spacecraft, th eseparation and release system (Figure 5-6) is

l b l

(a) View 1(b ) View 2

FIGURE -6. Separation and release mechanism.

responsible for programming proper de-spin ofthe system, erection of the appendages, andseparation of the payload. Th e system wasdesigned originally to adapt the spacecraftdirectly to the fourth stage of the Scout vehicle.A thin-walled, machined magnesium casting inthe shape of a hollow, truncated cone forms thebasic transition unit from the satellite to thelast stage. A t the satellite interface is theseparation spring, segmented marman clamp,and separation bolts. Here, too, are found the

two 3-pin feed-through connectors that carryinto the satellite the de-spin signal, batterycharge, and turn-on, turn-off circuitry. Thewide base of the cone fits over the forwardshoulder of the vehicle interface and is boltedto studs protruding from that surface. Fas-tened t o the cone itself are two 12-volt batterysupplies and an electronic release sequencer.Bonded to the cylinder of the X-248 motor arebrackets and supports holding the release cordsand pin pullers-which, when retracted, re-lease the cords and allow erection of theappendages. Total balanced weight of thissystem is 18.3 pounds.

Delta Dutchman

When the Ariel I was transferred to theThor-Delta vehicle, an adapter ring was re-

quired to move the folded payload forward some12 inches so that the pedal-leaf second- tothird-stage separation skirt would not damagethe sensor booms during stage separation.This Dutchman (Figure 5-7) is a fabrication ofcast magnesium end rings separated by a 16-inch-diameter cylinder of rolled and welded0.094nch-thick magnesium sheet. The rivetedfabrication is 13.0 inches long.

Since the Thor-Delta can carry a heavierpayload than the Scout, the available space inthe Dutchman cylinder is utilized to containthe electronics and sensors of a vibration andcontamination experiment. The total weightof this composite is 17.0 pounds; the Dutchmanaccounts for 7 .3 pounds of this total.

Detailed Data on Structure and MechanicalDesign

Physical measurements of the structures andmechanisms of the Ariel I payload are presentedas Appendix A.

THERMAL DESIGN AND COATINGS

Thermal DesignFo r a given solar absorptance-thermal emit-

tance (ale) ratio, a variation in mean orbitaltemperature of 35' C can be expected in themain structure over a period of 1 year. Thistotal variation may be subdivided as follows:( 1 ) 0 to 6 watts internal power, 5.' C ; (2) 63 to100 percent sunlight orbits, 20' C ; and (3)variation in sun, spin-axis angle-including

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ARIEL I : THE FIRST~~

INTERNATIONAL SATELLITE

( 0 ) ( b )

(a) View 1 (b) View 2FIGURE -7. Dutchman adapter cylinder and experiments.

shading by paddles and booms, 10' C. A f 0percent tolerance for achieving the design alewould result in a tolerance of about f 7 ' C inmean orbital temperature for a given orbi t, sunangle, and internal power. Skin temperaturegradients can vary from about 20' to 60' C,depending on sun-spin-axis angle. The varia-tion in mean orbital temperatures of the boom-

mounted components due to sun-spin-axisangle changes will be somewhat greater than theccrresponding temperature predictions for themain st ructure because of the nonsphericalgeometry and greater shading effects. Solarpaddle temperatures should remain between+33 ' an d -63' C, the temperatures corre-sponding to the hottest and coldest combina-tions of spin axis, orbital plane, and orbita lposition locations with respect to the sun.

Thermal Coatings

The thermal coating are evaporated gold withnot more than 25 percent of the total surfacearea covered with a combination of black andwhite paints. This represents a compromise ofthe experimenters' requirement for a conductingsurface with a preference for gold or rhodiumover other metals and a thermal requirementfor maximizing the painted areas to minimizeth e tolerances on ale ratio.

The following process has been developed forapplying the thermal coatings so th at mirrorlikegold surfaces which have good adhesion to thesubstrate and which can withstand withoutdamage aerodynamic heating to 25OOF areobtained. First, the surfaces are sanded,cleaned, and baked at 310'F for 1 hour. Thenlayers of varnish, lacquer, paint , and metals

are applied in the following sequence and arebaked at the temperatures and for the timeintervals indicated:

1. Sealing varnish, 300'F for 20 minutes,2. Clear lacquer, 290OF for 30 minutes,3. Conducting silver paint, 28OOF for 18

4. Electroplated copper, 1.5 mil thickness,5. Lacquer, 275'F for 30 minutes,6 . Evaporated gold, 0.00004 nch thick,7. Four coats of black paint in eight

longitudinal stripes, 250F for 30minutes.

Steps 3 and 4 are applied only to the cylindri-cal section and the forward dome to provide aground plane for the antennas. The lacqueris the type used for providing a proper sub-strate for evaporative films. The combinationof the varnish plus the repeated baking atdecreasing temperatures with each successivestep is used to protect the evaporated gold

hours,

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from damage by outgassing of the subst rat e.Final adjustment of a/e was made by theapplication of white pain t over pa rt of the blackpaint areas. The black and white paints arecarbon black and zinc sulfide pigments, re-spectively, in a silicone vehicle.

DESCRIPTION OF ELECTRONIC SYSTEM

Electronic Subsystem

A typical electronic subsystem of t he space-craft is a printed circuit card, which is com-pletely encapsulated in Eccofonm (Figure 5 -8 ) .

Ti *-

Ie.0-

FIGURE -10. Printed circuit card, mounting com-ponent of medium density.

FIGURE -8. Printed circuit card, completely encap-sulated for space flight.

These printed circuit cards are of three types.One type of circuit card provides for mountingindividual electronic components as shown inFigure 5-9. Medium component density is

C L i P YU-mr-C: *f

UNI TED STATES SATELLITE A N D SUBSYSTEMS

35

FIQURE -9. Printed circuit card, mounting individualcomponents.

achieved by mounting the components verticallyon the record type of printed circuit cardshown in Figure 5-10. Maximum component

density is achieved by mounting modules on theprinted circuit cards as shown in Figures 5-11and 5-12.

FIQURE -11. Printed circuit card, mounting unencap-sulated modules.

FIQURE -12. Printed circuit card, mounting encap-sulated modules.

Block Diagram Discussion

A block diagram of the spacecrafts electronicsystem is given in Figure 5-13. Eight probesare shown with their dnta outputs processed


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ARIEL I : THE FIRST INT ER NA TI ON AL SATELLITE

CON DITIONIN G CIRCUITS ANDPROBES SUBCARRIER OSCILLATORS

POWER SYSTEMS-ENCODER-CONTROL CIRCUITS-TAPE RECORDER

I- IGNAL-

OWERAPERECORDER~ .CM(TR0L

FIGURE -13. Ariel I electronic function diagram.

by signal conditioning circuits. The output sof the signal conditioning circuits are connectedto subcarrier oscillators (analog and digital).The subcarrier oscillators convert analog anddigital da ta to representative ac signals. Thefrequencies of these signals are dependent on thevalue of the input analog voltage and/or digitaldata, Figure 5-14. The outputs from thesubcarrier oscillators are sequentially gated as

high speed (HS) or low speed (LS) data outputsby the respective encoder matrices. Theencoder assignments of experimental dat aoutputs are listed in Table 5-1.

The high speed encoder data are transmittedin real time, whereas the low speed encoderdata are recorded on magnetic tape. There isa speed reduction between the two encoders of4 8 : l . By playing back the recorded data 48times faster than it was recorded, the low speed

36

data are transmitted at the same bandwidthas the high speed data.

Playback of the magnetic tape recorder isinitiated by a ground station command to theCommand Receiver in the spacecraft. Th eCommand Receiver in turn commands theProgrammer to switch the Tape Recorder fromRecord to Playback.

The out put format of the high speed and low

speed encoders is shown in Figure 5-14; alsoshown is the telemetry program for almost twocomplete orbits.

The power system for the spacecraft consistsof solar paddles, two battery packs, shuntvoltage limiter, battery chargi

ariel i, the first international satellite project summary

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