Meteorological Satellite Data Systems

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I T H R U I

X-650-64-241

T M X-55093

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METEOROLOGICALI S AT E L L I T EDATA SYSTEMS

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DARD SPACEFLIGHTCENTER -

GREENBELT,MD.



METEOROLOGICAL SATELLITE DATA SYSTEMS

Herbert ButlerAeronomy and Meteorology Division, GSFC

September 1964

Goddard Space Flight CenterGreenbelt, Maryland



CONTENTS

Page

INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

OBJECTIVES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

ORBITS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3

EQUATORIAL ORBITS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

POLAR ORBITS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

SENSOR PLATFORMS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7

Stabilization and Attitude Control . . . . . . . . . . . . . . . . . . . . . . . . 7Three-Axis Stabilization . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7Spin-Stabilization . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11Gyromagnetic Stabilization . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 2

Gravity-Gradient Stabilization . . . . . . . . . . . . . . . . . . . . . . . . . . 13

SENSORS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

T V Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 5Automatic Picture Transmission (APT) . . . . . . . . . . . . . . . . . . . . 1 6Inf ra red Sensors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 9

GROUND STATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20

TIROS I . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 0

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ILLUSTRATIONS

Figure Page

1 Goddard Missions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2

2 Meteorological Satellite Development . . . . . . . . . . . . . . . . . . . 3

3 CoverageGeometry . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

4 Synchronous Meteorological Satellite . . . . . . . . . . . . . . . . . . . 5

5 Nimbus Television Coverage . . . . . . . . . . . . . . . . . . . . . . . . G

6 Nimbus Stabilization . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

7 TIROS Picture Coverage Comparison . . . . . . . . . . . . . . . . . . 1 0

8 Gyromagnetic Stabilization . . . . . . . . . . . . . . . . . . . . . . . . . . 1 2

9 Gravity-Gradient Configuration . . . . . . . . . . . . . . . . . . . . . . . 14

1 0 TV Picture System. Block Diagram . . . . . . . . . . . . . . . . . . . . 1 7

11 Orbital Coverage fro m Fairbanks. Alaska . . . . . . . . . . . . . . . . 2 3

1 2 TIROS I Ca mer a Field of View for 380-Nautical-M i l e o r b i t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 5

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METEOROLOGICAL SATELLITE DATA SYSTEMS

.INTRODUCTION

There a re many pra ctic al fa ct or s affecting the design of da ta sys tems fo rmeteorological sate llit es which must be taken into consideration. It is the pur-pose of th is discus sion to identify and examine som e of th e key el eme nts o r var-iabl es which a re a pa rt of the complex matrix that mu st be solved to produce aneffective sys tem design, and to illu stra te how suc h a matrix w as solved fo r oneof the for thc omi ng TIROS satellites.

Figure 1 illu stra tes the wide variety of sha pes and sizes of s pa ce cra ft which

we re used to me et an equally wide variety of functional require ments. F or me-teorological purposes, these shapes and sizes have been narrowed down to twobasic types: one in which the spacec raft o r sensor platform is stabiliz ed and con-trolled by means of an attitude-control system to keep the sen so rs accu rately ori-ented toward the earth, and the second in which the spa cecraft spins at a moder-ately rapid rate (- 1 0 rpm) to provide the stability of a gyroscopic body (Figure 2).

Underlying the design of both type s a r e many common requirements too de-tailed to be covered in this paper. They include such practica l fac tor s as power,data transmission, size and weight, ruggedness and reliability to withstand theri go ur s of launch and environm ent of space, and the need for the design to allow

adequate ground testing before launch.

OBJEC TIVEP

The o bjectiv e of a meteorological satellite syste m i s to provide observa-tions ove r the entire globe on a daily basis. These observations should includecloud-picture covera ge and cloudtop measurem ents on both the su nlit portionand the nighttime portion of th e eart h, at le ast once every day. Becaus e the datais perishable, it should be collected from the satellite and relayed to a cen t ra lprocessing point with no more than a two to th re e hour delay. In addition toproces sing, preparing, and disseminating global information fro m a cent ral point,it i s desi rabl e to provide direct readout from the satellite to local stations. Withsuch a capability, a us er anywhere in the world could receive data from the sate l-lite, when it was i n his vicinity, which would c ove r a significant area surroundingthe user.

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Lnc0.-VI

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Figure 2 -Meteoro log ica l Sa te l l i t e Deve lopment

The resolution of the cloud pic ture s should be in the o r d e r of one to two

miles at the subsatellite point (Figure 3) and should not exceed four to five milesat the point of contiguity (the point at which picture s fro m adjacent orb its ov er-lap). At the point of contiguity, the angle between the local ver tic al and the se n-sor in the satellite should not exceed 65 deg rees (zenith angle).

ORBITS

Sev era l fa ct or s must be taken into consideration when selecting the type oforbit that would m ost closely satisfy all requi remen ts for a meteorological satel-lite pro gra m. Some of the maj or factors ar e cost, reliability, time and ar e a re -

quired for proper data coverage, power requirements , equipment par ame ters ,and operating requirements , weight of spacecraf t, and location of da ta-r eadoutstations . The principal advantages and disadvantages of so me of the mo re com-mon orbits a re discussed in the following par agr aph s.

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1st O R B I T S U CC E E DING OR BIT

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C E N T E R OF

y E A RT H

Figu re 3-Coverage Geometry

EQUATORIAL ORBITS

Because of launch-vehicle range safety requirements, all U.S. satellites a r elaunched over water. Thus, launches fr om Cape Kennedy have been re str ict ed toorb its which a r e equatorial but a r e inclined to the equator at angles ranging fro m-28 degrees to +58 degr ees. The 58-degree o rbi t has been quite useful for TIROS.

It has provided coverage o ver the mo st densely populated areas of th e ear th , butit cannot provide the much-needed coverage over the polar regions. Fu rt he r, theorb it plane of this type orbi t pre ces ses about the ear th at a fa i r ly rap id rate(about 4 degrees per day) so that the relations hip between the sun and the orbitplane varies, resulting in a varying daily time of observation and peri ods of in-adequate illumination.

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A particularly interesting variation of the nominal equatorial orb it is thepu re equatorial, i.e., wherein the satellite is ste er ed into an orb it plane coincid-ing with the equa torial plane of the earth. At the nominal alti tudes used f orTIROS (approximately 400 nautical miles) , th is would provide only a narrow belt

of coverage approximately 500 nm above and below the equa tor. However, if thealtitude could be raised to about 2500 nm, the a r e a of observ ation could be ex-tended to a very useful *30 degrees of latitude (*1800 nm). The sa tel lit e wouldcover the same area every three and one-half hours and could provide an excel-lent hurricane 'lwatch" a s well as frequent observ ation of o ther me teorolo gicalphenomena origin ating in the tr opica l zone.

A fu rth er variation of the pure equatorial or bi t is the earth-synchronous or"stationary" orbit. A t approximately 22,300 miles, a satellite, like Syncom, willmove about the ear th at the s ame angular velocity a s the ea rth and therefor e willnot move relative to the earth. It can be made to hover above any given point on

the equator (Figure 4) nd could observe the sa me ar ea continuously for hou rs,days, o r months. This orbit would be particularly valuable, of cou rse , for thestudy of the development of a storm area.

Figur e 4-Synchronous Meteorological Satel i te



POLAR ORBITS

The minimum orbit al requiremen t fo r meteorological pu rposes is globalcoverage of the ent ir e sunlit portion of the ea rt h at lea st once every day. A near-po lar o rbit having an inclination of app roxim ately 80 degrees to the equator (re t-

rogr ade) permits the plane of the orbit to pre ce ss at a ra te equivalent to that ofthe rotation of the ear th around its axis so that it i s synchronized with the sun,thus maintaining the sat ellite orbital plane on the earth-sun line. The best con-ditions for ea rth viewing ar e obtained by launching the sp ace cra ft at local noono r local midnight. This launch time yields a n orbital plane that contains theearth-sun line. Consequently, the spa cec raf t will always view the earth at nea rlocal noon on the sunlit side and near midnight on the dark side (Figure 5). Thesun-synchronous orb it also enables a spacecraft such as Nimbus to require onlyon e axis of paddle rotation for direct sun-pointing.

F i g u r e 5-N im b us Te l e v i s i o n C o v e r a g e

SUN RAYSFIRST ORBIT

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Two pra cti cal variatio ns a re possible with the sun-synchronous orbit: one,to launch the spacecraft at 9:00 a.m. o r 9:00 p.m. to provide observ ations ea rl yin lhe day; o r two, to launch the spac ecraft at 3 : O O p.m. o r 3 : O O a.m. to providelate-day observations.

An interesting variation of polar orbit is the pure-polar, o r 90 degree in-clination, orbit. This orbi t would permit complete ear th coverage and completedata transmis sion by using a single station located nea r one of the polar extrem-it ie s of the earth. However, th er e are several distinct disadvantages in the ap-plication of a pure-polar orbit. The space craft 's orbital plane would reg re ss atthe rate of 1 degree per day, resulting i n a daily change in illumination level.Approximately 30 da ys of each 180 would be sp ent in the twilight zone, thus p ro -ducing se ve re proble ms in maintaining adequate therm al contro l and power.Another d isadvantage would be that instead of on e axis of paddle rotation re-quired fo r a near- polar orb it, two such axes would be re quire d for di re ct sun-

pointing.

SENSOR PLATFORMS

Stabilization and Attitude Control

-

There ar e sev eral established active and passive methods fo r stabilizingand controlling the attitude of sat elli tes f o r meteorological applications. Theprincipal determinants to be considered for selecting the best system dependupon the a ccurac y nece ss ary fo r optimum perfo rmanc e of th e meteorological

sen sors to be carr ied on the satel l i tes . Other fac tors to be considered ar e al t i -tude, accuracy, cost, complexity, reliability, control lifetime, redundancy, weight,and power requi remen ts. Another outstanding requir ement is that the spacecraftbe designed so that the cent er of gravity is a cce ssi ble to per mit dynamic testing.There a r e seve ral advantages and disadvantages inherent in each system. Someof the major sys tem s a r e discussed in the following paragraphs.

Three -Axis Stabilization

The three-axis stabilization control system (Figure 6) has been successfullydernonstrated i n orb it to be an effect ive method fo r controlling the attitude of asatel l i te . For example, the Nimbus spacecraft functions as a s tabilized platformwhich carr ies a variety of sensors to view the earth continuously both day andnight.

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SOLAR PADDLE

PITCH

E A RT H

DIRECTIONIN ORBIT

F i g u r e 6-N imbus Sta bi ization

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-

-

- -

- *

The Nimbus attitude con trol subsystem employs two in fra red h orizon scan-n e r s , tw o coar se sun sens ors and a rat e gyro acting as a yaw sensor. Threemotor-driven flywheels and eight gas nozzles act as torque generato rs to pro-vide attitude control. A pneumatic tank contains the gas supply.

The spa cecr aft coordinate sys tem defined by the location of the op tical sc an-ne rs and inert ia l devices used for axis orientation is:

0 The yaw axis points toward the center of the ear th

0 The roll axis, perpendi cular to the yaw axis, i s para llel to the orbitalplane

0 The pitch axis is perpendicular to both the yaw and roll axes

After s epar atio a fr om the launch vehicle, the horizon sc ann ers located onthe roll axis sen se the spa cec raf t's attitude with res pec t to the eart h. The scan-n e r s , one looking forw ard and the othe r looking to the r e a r , ge nera te a sky-earthsignal which, when applied to t he attitude compute r logic ci rcu its , produces bothpitch and roll e r r o r signals, thus operating the pitch and roll flywheels and noz-zl es to sta bili ze the spac ecra ft in pitch and r d .

The position with respect to the sun of the co ar se sun s ens or provides yawattitude control during initial stabilization. E rr o r signals fro m the sens ors a reamplified to drive a yaw flywheel and gas nozzles to reduce the negative rollaxis-sun angle to a minimum.

When the sp acec raft pa ss es into the ea rt h' s shadow on the first orbi t, yawcontrol is switched to the ra te gyro. The gyro produces an er r o r signal propor-tional to any divergen ce of the spac ecra ft roll axis fro m the spacecraf t velocityvector in yaw. This e r r o r signal is applied to the yaw reaction wheel to correctthe yaw error.

The ga s nozzles reduce the larg e stabilization e r r o r s , while the flywheelscompensa te fo r smal l e r r ors . A tachom eter m onit ors the speed of the flywheelsto preven t flywheel satura tion . When the saturat ion point is approached, the gasnoz zles a r e activated to reduce the speed of the flywhe els back to zero.

Under ide al conditions, i t i s believed that pointing acc ura cie s in the range of*l to *2 degrees will be achieved. However, depending upon atm os ph er ic condi-tions in th e field of view of the horizon scan ners , er r o r s of approximately twicethi s m agnitude could be introduced.

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.D

?I+(A ) CO NVENTIONAL TIRO S ORIENTATION

(B) TIROS WHEEL' ORIENTATION

Figure 7-TIROS Picture Coverage Comparison

1 0



Spin-Stabilization

A spinning satellite (Figure 7A) normally will not maintain a simple orienta-tion reference with respect to the ear th s ince it s spin axis is fixed with res pec t

tospace. A reference is derived if th e attitude of th e spin

axiswith resp ect to

the earth is known, as well as the angle between the camera and the sun (a "fixed"refere nce) for each picture taken. Spin-stabilization ori ent s the spacec raft withrefe ren ce to the e arth without wobble o r prece ssion about the spin axis by meansof a proper selection of the ratio of principal moments of in er tia , toget her withan internal energy absorber. For instance, in the standa rd TIROS configuration,the spin and ca me ra axes are paral lel and s pac e stabilized in the orbi t plane.Ther efore, the sen sor s view the ear th only during a sm al l portion of each orbit .

The TIROS wheel configuration (Figure 7B) is mor e efficient in that it pro-vides the advantage of an ear th oriented ca me ra system in a spin-stabilized

spacecraft so that the cameras look down at the ear th from any point in the or bi t ,thus gre atly increasing data coverage capabilities.

The spacecraft spin rate is controlled by thre e devices af ter final separation.Thes e spin-control dev ices consist of the despin mechanism (yo-yo), the spinuprockets, and a magnetic spin-control device.

The yo-yo provides fo r the initial spindown fr om the boost-stage stabil iza-tion rate (-120 rpm) to its operational spin-rate (-10 rpm). This i s accomplishedby releasing a s e t of two des pin we ights which are caged during boost and attachedto cab les that a r e wrapped around and hooked to the sa tel lite housing. When re-lea sed , centrifugal forc e ca use s the weights to unwind the cab les fr om the hous-ing. The weights thereby acqu ire angular momentum fr om the satell ite, which i s

slowed accordingly.

The spinup rockets (10) are mounted in diametrically opposite p a i r s underthe periphery of the baseplate to provide for coars e increases in the sp in ratewhen commanded to op er at e by groun d command.

The spinup rockets a re being replaced by me ans of a technique wherein thespin rate can be controlled by a magnetic spin-control device sim ila r to a d.c.

motor. The magnetic spin-control device per mits close , ground-command con-t rol of the sa tel lite spin-rate and consists pri ma ril y of a coil of wire woundaround the peripher y of the sat elli te and a pulse-controlled, solenoid-operatedstepping switch. The stepping switch, operating in respo nse to ground-stationcommands, in cre ase s o r dec rea ses the amount of cur ren t flowing in the coil.Each change in curr ent effects a corresponding change in the mag netic field ofthe s at el lit e, providing (through the interac tion of the magnetic fie ld s of the

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spacecraft and th e ear th) the mechanica l force necessary fo r increas ing o r de-cre asi ng the satellit e spin ra te required f or attitude control.

F o r TIROS, it has been possible to control the attitude of th e sp ace cra ft towithin a few deg ree s and, to determi ne the attitude, aft er the fact by analy sis of

the da ta, to approximate ly one degr ee. This is achieved by using both attitudetelemetry data from sun and horizon sen sor s and by geographical referencingfro m the pictures.

Gyromagnetic Stabilization

Gyromagnetic stabilization (Figure 8) is a s imple, thr ee axis , a t t itude-cont rol system utilizing the principle of magnetic torquing fo r controlling eit herspinning o r non-spinning earth -orbiting sa te ll ite s by means of ground-stationcommands. This system has many desirable feature s : simplicity , reliability,a potential fo r long operational life, low weight and power r equ ire men ts, andadaptability to sat ell ite sy ste ms of any si ze o r weight.

YAW

STA 6 LIZEDPLATFORM

MAGNETIC

TORQUING COI

(NORMAL TOORBIT PLANE)

ENSOR TO DETECTLOCAL VERTICAL

ORBIT^DIRECT ION

AREAITO EARTH'S CENTE R

F i g u r e 8-Gyromagne t ic S t a b i l i z a t i o n

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Fig ure 9-Gravi ty-Gradient Configurat ion

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mech anical dis plac em ent, and rotation of sat elli te pa rt s.provided by using a lossy sp ring o r hyst eresis ba rs .

Damping control is

The best pointing a ccurac y achieved to date h as been in the o rd er of *5 to

1 0 deg ree s with no moving par ts (not even came ra shu tter s or rela ys) on board.However, becau se of the at tra cti ve ne ss of t hi s technique fr om the viewpoint oflong life and reliability, considerable effort is being expended to improv e thestability.

SENSORS

The most useful meteorological function performed by satellites to date hasbeen to provide observations of the cloud cov er of the e ar th , and to de mo ns tra tethat re gul ar and continuous observ ations of the cloud cove r can be effectively in-

tegrated into a meteorological system . Both TV-picture and infr are d-ra dio met ricsensin g sys tem s have been employed, and both techniques have demo nstrate d thatthe data can be used to gain a be tte r understanding of the ea rt h' s atmo sphere .The TV pictur es, used operationally f o r more than two ye ars , have been limitedto the sunlit portion of the e ar th bec aus e of inheren t sensitivity limitations in thesim ple vidicon televisim system presently employed. The infrare d sy ste ms havedemo nstrate d the ability to provide observations of the nighttime cloud cov er andinformation leading to the determination of cloud heights.

TV Systems

The pr im ary ch ar act eri sti c of the vidicon ca me ra tube which ha s made itexceptionally well-suited to satellite application is the image-storage capabilityof the vidicon targ et surf ace. The photosensitive t arge t mate rial can st or e animage projected on i ts surface fo r several seconds after an exp os ur e of only afew milliseconds. This allows time f o r a slow-scan readou t of the ima ge andre su lt s in a frequency-bandwidth c omp res sio n of the video signal. Thus, a nar-row er rad io freq uen cy bandwidth (compared to conventional television band widths)can be employed to tra ns mi t the data to ground o r to accum ulate the d ata on arelatively s imple tape re cord er before i ts t r ansmis sion to ground.

In the TIROS sa tel lit es , a half-inch dia met er vidicon tube has been employed,providing a 500 scan-line image. The exposure tim e is 1.5 milliseco nds with areadout t im e of 2.0 seconds. Thirty-two p ictu res can be sto red on the tape re -co rd er and played back to the ground station i n l es s than two minu tes o ve r achannel less than 1/2-m egacy cle wide. When the sa tel lit e is in range of theground station, the pictures can be transmitted directly, bypassing the tape

1 5



reco rde r. The camera s have been used in pa ir s fo r redundancy and ar e ope r-ated as shown in the functional block diagram (Figure 10).

In the advanced vidicon camera system designed for Nimbus, three identicalcam era s ar e mounted in a trimetre gon arrange ment with overlapping fields of

view. The th re e adjacent pic tur es are aligned ac ro ss the directio n of the o rbitand provide television covera ge of a 420 by 1400 nautical-mile a re a fr om an alti-tude of 500 nautical miles. Each cam era employs a one-inch vidicon cameratube providing an 800-scan-line image; each ca me ra is equipped with a variableiris which per mit s the ex posure to be optimized for so lar elevation angles rang-ing f rom 90 deg rees down to approximately 4 degrees . The three cam eras a reexposed simultaneously fo r 40 milliseconds, which is practic al only on a thre e-axis stabilized platform, and a re readout in parall el in 6.25 seconds resu lting ina bandwidth compression simi lar to that describ ed ear lie r. The th ree pictur escan be simultaneously transm itted o r recorded on a four- t rack tape re cord ertoget her with refe ren ce tones and timing signa ls. Each of the video trac ks can

store 5 7 pictures. Each cam era is equipped with built-in calibra tion faci lit iesfo r both sweep and gra y-s cale lin eari ty.

Table 1 shows the relation ship of o rb it altitude to ar ea cov erage and re so -lution for thr ee different len s/ca mer a combinations and typical coverage fo r theNimbus trimetregon arrangement.

a4utomatic Picture Transmission (APT)

The APT system is designed to provide wide-angle cloud-cover pictures in

rea l t ime to local ground stations fo r immediate use. The equipment employsthe sam e basic principles a s the TV cam era s describ ed above and is capable oftaking pictu res continuously throughout the sunl it portion of each orb it.

The APT sy ste m co ns is ts of an 800-scan-line vidicon, an F M t r a n s m i t t e r,and associa ted electro nics. The vidicon is s i m i l a r to oth er 800-scan-line vidi-cons except for the addition of a polystyrene lay er to provide extended image-sto rag e capability. The tube is operated through the pre par e, expose, and read-out phases by varying the mesh potential with respect to the target potential; i.e.,the image is projected on a prepared photoconductive lay er, then tra ns fe rr ed(developed) by potential change to the sto rage la ye r f o r readout. The prep are ,expo se, and develop ope ration s ar e accomp lished during the f i r s t eight secondsof ea ch 208-second pi ctu re cycle. During th e next 200 seco nds , the pi ctu re in-format ion is read out line by line at a scanning ra te of fou r l ine s pe r second andtransmitted to the APT ground stations. Sta rt and phasing signal s a r e transm ittedto the ground stat ion s at the beginning of each pic tu re to syn chr oni ze the ground

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Table 1Picture Coverage Resolution and Viewing Angle forDifferent Sensor Lenses and Spacecraft Altitudes

1.15

Altitude

Lens

Picture size-diameter innm inscribed in circ le

1.8esolution in nm/TV lineat Z = 0"

Resolution i n nm/TV lineat Z = 65"

IZ at center edge

Z at corner

375 nm

440 I 720

0.9 1.4t.1 3.2

34" 1 4 8 "

770-0.9-2.1

50 "

46 " I61" 164"

A = HA = 39", EHA = 30° , W/500 line vidiconB = HA = 52*, EHA = 42", W/500 line vidiconC = HA = 54", EHA = 44", W/800 line vidiconnm = nautical mile s

E D GE H A L F A N G L E

( E H A )

I H A L F A N G L E

500 nm

T00 960

2.7 4.3

z

C

1050

1.2

2.9

53 "

46" 164" 68 "

750 nm 1

900 1500

4-.7 12.7t.1 6.4

38" 155O

50" 174"

N IM B U S AV C S A L T 7 5 0 n m

C

1600

1.8

4.3

5 8 "

8 0 "

a/

II

II O bI I

I I

C O M P O S I T E P I C T U R E D IM E N S I O NS ( oc x d e ) 5 1 0 x 3 1 7 0 nrn

AV E R A G E R E S O L U T I O N 0 6 - 1.3 nrn / T V L i n e

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station facs imil e re co rd er s with the vidicon scanning beam. The syst em hasbeen so designed that only rela tive ly simple and inexpensive ground stationequipment is required.

Infrared Sensors

A number of inf rar ed radi ome tric experim ents have been flown on TIROSsa te ll it es which have util ized th e combination of t he spin ning motion of the sa te l-lite and its motion along the o rbi t to provide a means of scanning the earth's sur-face. These measur eme nts have provided some very useful information on theearth's albedo (0.2 to 6 micron), on the water vapor emission band (5 .9 to 7 . 0micron), on the thermal emission from the ear th (7 .5 to 3 5 micron), and on thenear-visi ble portion of the spect rum ( 0 . 5 to 0 . 7 5 micron) for refer ence and com-parison with the TV system s. The most promising observat ions have been madein th e 8- to 12- mic ron "window." In th is window, the radi atio n pa tt er ns have pr o-vided a rema rkab ly good desc ript ion of the ear th 's cloud cove r in both daylightand nighttime areas , thereby indicating a potential for full global coverage witha single sensing system.

-4 ypical rad iom eter cons ists of an optical sy ste m, a photoconductive detec-tor , associated electronics, and a mechanical drive-all enclosed in a suitablehousing. In co nt ra st to television the radiometer forms no image, but insteadinte grat es the energy received from the target. Composition of a picture i s

achieved by a scanning m ir ro r technique. The m ir ro r, located in the radio met er,scans the eart h a s the satellit e advances in it s orbi t. The mi rr or ref lects thereceived energy and focuses it on a mechanical chopper, which provides the nec-e ss ar y modulation of the energy signal. The modulated sign al act ua tes the dete c-tor which produces an el ectr ical output signal correspondin g to the energy-signalintensity . The output is recorded on a tape reco rde r designed to st ore the datataken fo r the full duration of an orbit o r fo r multiples thereof.

The data a r e rec orded at the ground station and converted fro m analog todigital for m to perm it computer processing. Data a r e then re-reco rded onmultichannel tape which c ar ri es the data as well a s all pertinent reference datasuch as time and location coordinates; this multichannel tape is ready for machineprocessin g to produce meteorological information.

Plans for two improved infrared radiometers for the TIROS wheel satellitea r e presently being investigated. One of these is a modified Nimbus five-channelmedium-resolution radiometer (MRIR) to yield a one o r two channel instrumen toperating in the 10-11 mic ron window. The pr es en t design concept of t he NimbusMRIR would be re tain ed; however, the size of the r adio mete r cast ing, scanning

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mirror, and electronics module would be reduced to reflect the reduction in thenumber of channels. Also, the nomin al scann ing ra te , field of view, and info rma -tion bandwidth would be changed to be com patible with the sa tel lit e' s o rbi tal pa-ram ete rs and the data gathering requirements of the mission. The rota tion of th escanning m ir ro r would be phase-synchronized with the spin of the wheel sa tel lit e,resulting in a sca n pattern having a diagonal orientation with respect to the satel-lit e velocity vector and pas sing through the sub sate llite point.

The other infrare d sys tem which is being actively investigated i s a multi-se ns or medium-resolution instru ment. This design con sist s of 1 6 individualsmall reflective optics and thermopile detectors mounted in a stepped fashionaround the cyli ndri cal surf ac e of the spinning wheel. The opt ica l axes of thesens ors are inclined at different angles to a plane normal to the spin axis. Th esen sors are sequentia lly sampled as the axis of each instantaneously lies in aplane passing through the spin axis and subsatellite point, resulting in a t r ans -verse row of sca n spots which make up a swath extending about 1,000 miles ove rthe sur face of the e art h centere d about the subsate ll i te point.

GROUND STATIONS

The location of ground statio ns to re ad out da ta and to contro l the op era tionof the satellite is a complex proble m of geography, lo gis tic s, and comm unica-tions, and is a maj or element in the design of a satellite system. Locating sta-t ions is particularly difficult for a polar orbitin g satell i te because the groundstation for this sys tem should be located eith er at the North o r South Pole s inor de r to "see" and cominunicate with the sat elli te at l eas t once pe r o rb i t o r elsethe data will be delayed by some multiple of an orbi tal period. However, the re -

quirement fo r continuous wideband communications (approximately 50 to 1 0 0 kc )to transmit the data to a central pro cessin g point makes it almost impossible,o r at least impractic al, to locate fulltime operational facili t ies within the A rcticZones. This i n turn makes it necess ary to use at least two ground stationswidely separated i n longitude to provide dat a recor ding facilit ies capable ofsto ring at leas t two orb its of d ata on board the sate llite .

TIROS I

Th e TIROS I sate l l i te sys tem is an experime nt to obtain cloud-picture cov-

era ge of the ent ire sunlit part of the ea rth on a daily bas is using a spin-stabilizedsate l l i te .

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To satisfy the global coverage requirement, a polar o r near-polar orbi t wa sindicated; beca use of its special advantages for a meteorological satellite, a sun-synchronous orbit was se le ct ed . Advantages of a sun-synchronous orbit are con-stant ground illumination and relatively constant satellite/sun angle which facili-tat es pic ture inter preta tion and simplifies the the rma l and power supply designsof the spacecraft. Also considered to be an advantage by many meteor olog ists is

the fact that all cloud pictures will be taken at the same local t ime; i .e ., if th esatellite is launched at 1 O : O O a.m., then the cloud pictures fo r eve ry orbit will betaken at 1 O : O O a.m. The tim e recommended fo r thi s expe rimen t is 3:OO p.m. inor de r to provide the meteorologists with late-day o bservations.

In the standard TIROS configuration, the spin axis and the camera axis a r eparal lel. Since the spin axis is fixed in space and is stabilized in the orb it plane,the cam era s view the earth for only a small pa rt of the orb it as shown in Figu re 7A.If a sa te lli te of th is configuration w a s to be launched i n a sun-synchronous orbit

and the spin axis steered to a position in the orbit such that it was vertical mid-way in it s usefu l ar ea , then it could provide about 50 de gr ee s of latitudinal cove rage.

However, as indicated earlier, TIROS has a wheel configuration in which theca me ra s a re positioned radially instead of paral lel to the spin axis, which i s muchmor e efficient. The cam era s can observe the ent ire sm!it pa rt of the or bi t, nom-inally about 120 degrees. A fur the r advantage in the sy stem as designed fo rTIROS I is that, by me ans of a fair ly s imple horizon sen sor t r iggering ar range -ment for the camer a shutter, all of the pi ct ur es are taken st ra ig ht down. Takingthe pictures straig ht down will provide a much m or e uniform product fo r themeteorologist to analyze and w i l l require a minim um amount of re cti fic ati on tocor rec t fo r p ic ture d i s tor t ion .

TIROS I will us e two ground stations which togethe r w i l l intercept nearlyall of the orbit al pa ss es ( 8 0 percent). The polar station, located nea r Fairba nks,Alaska, is not quite f a r enough north to se e all of the pass es (Figu re 11). Eventhe addition of an Ea st Coa st st ation at Wallops Island, V irgin ia, did not quitesolve the problem. A compromise was made by providing tape- record er sto rag ecapacity on board the spacecraft to accommodate at l e a s t four or bi ts of pict uredata. The probl em of inter cepti ng all the passes i s of co nsid erabl e conc ernsince the orb its that cannot be seen from eithe r station generally occur out ove r

ocean a re as w here information on the weather is a t a premium.

'

The most s erio us problem was encountered in the design of the ca me ra sy s-tem to provide continuous coverage with no gaps at the nominal TIROS orb it alti-tude of 3 8 0 nautical miles. With the standa rd TIROS one-half inch vidicon andthe widest angle lens (104 degrees) pract ical fo r thes e applications, a gap of ap-proximately 900 nm would ex ist at the equator between succeeding orb its ; this

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gap would n ar ro w down away f r o m th e equator until contiguity is achieved at alatitude of approximately 68 degrees. I t is obvious that even the use of one-inchvidicon cameras would not close the gap.

The solution to the problem was to cant the cameras. The optic axes, in-ste ad of being par al le l to the sp in axis as in previous TIROS satellites or atright angles to the spin axis as originally conceived fo r the wheel configuration,will be canted approximately 26.5 degr ees to each sid e of t he plane of rotation.Figure 1 2 shows the a re a covered by a single f r a m e of o ne of the c am er as . When

30°

FIELD OF VIEW FOR52 DEGREE HA LF ANGLE LENS26.6 DEGREE CANTING ANGLESCALE : 1 " ~00 NMALTITUDE= 380 NAUTICAL MILES

150"

200

IO"

0

350°

340°

330"

160"

I 70°

180°

I 90°

2000

2100

Figure 12-TIROS I Camera F ie l d of Vie w for 380-Na ut ica l -Mi le -Orb i t

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combined with t he mi r ro r image provided by the second camera, a "butterfly"patte rn is created that extends approximately 1000 nm to eith er sid e of the s at el -lite tra ck and provides adequate overlap at the equator.

Some comp romi se was ne ces sar y in zenith angle which exceeded design ob-

ject ives by a few de gr ee s. Spec ifica lly, the zenith angle at the point of contiguityat the equator is 68 degrees although 65 degrees is the pre fer red maximum value.One other limitation of the canted c am era design is that som e of th e redundancyafforded by the use of two came ra sys tem s i s lost s ince both cam era syste msare required to provide full global coverage.

Regretably, i t was not practical to include i n this spacec raft e i ther an infra-red sensor system o r a fu rth er tes t of the APT cam era system. However, it is

believed that the experiment to obtain global picture data will provide valuableinformation leading toward the development of an operational meteorologicalsatellite system.

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Meteorological Satellite Data Systems

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