TR-59-0000-00604

Copy 0. of 2A5**

ADVANCED WEAPON SYSTEMS STUDY

PART III - ADVANCED MILITARY SATELLITES

VOLUME I

SUMMARY

1 March 1959

DOWNGRADED AT 12 YEAR

:RVALS; NUT AUTOMATICAL6

i..,!,SSIFIED. DOD DIR 5200.10

Work Completed

10 September 1958

SPACE TECHNOLOGY LABORATORIES, INC.

P.O. Box 95001

Los Angeles 45, California

MIS TA4

'ww'ENIAI

this 

U.S.C., 110011111481 coololos bfoisolioo anodise oolloosi linfoopo of the *NW bolos wIlifo smoothy of flo Espionoo Um, Tills 793 sod rm an hossodssloo rovolooke of which IN any n ow to so osoolborkod porno, b orsbliblhof by low.
TR-59-0000-00604

Page I- ii

CONTENTS

Page

INTRODUCTION 1

CHAPTER 1 - COMMUNICATION SATELLITES 2

Section

1.1 LONG-RANGE COMMUNICATION 2

1.2 HIGH-DENSITY SERVICE 3

1.3 LOW-DENSITY SERVICE 5

1.4 NUMBER OF SATELLITES 7

1.5 BANDWIDTH REQUIREMENTS 7

1.6 POWER REQUIREMENTS 8

1.6.1 ZI Relay 9

1.6.2 Intercontinental Relays 10

1.6.3 Low-Density Traffic 10

1.7 "INTERIM" COMMUNICATION SYSTEMS 11

CHAPTER 3 - ASCENT TO 24-HOUR ORBITS 19

3.1 ASCENT TRAJECTORIES 19

3.2 ASCENT GUIDANCE 22

3.3 ORBITAL CORRECTIONS 25

CHAPTER 4 - ATTITUDE CONTROL, ELECTRIC POWER,

AND VEHICLE DESIGN 31

4.1 ATTITUDE CONTROL 31

4.2 ELECTRIC POWER 34

4.3 VEHICLE DESIGN 37

-SECREFIbis iseveseet seolelas lofemailea offeelhe the meNesel 411those of Os Usiled Saba wain N. amplity of Os tapkosin taws. This

U.S.C.. %diem 793 God 754. N. trsominlea or revolaft. it voMkts I. fay welest le ea weeNeelzed ma.. b PreMbilled W Isw.
TR-59-0000-00604

Page-I-iii

Figure

ILLUSTRATIONS

Page

1-la Areas (shaded) Originating Great Majority of

Communications of Military Interest 4

1-lb Earth as Seen from Satellite in 24-Hour

Orbit at Longitude 170•W 

1-1c Earth as Seen from Satellite in 24-Hour

Orbit at Longitude 30•W 

1-2 Continuous Coverage Areas for Low-Density

Polar Relay Satellites 

3-1 Total Trajectory for Satellite Launched on. Equator . . . . . .

3-2

3-3

3-4

3-5

4-la

4-lb

4-2

4-3

4-4

4-5

Mem

This donsiodi aidolos IsionnoNgo died Oho nollsool Odom of Is Wild see Wild Illy nrsoolinp *I Ms bilgoops Mos. Tido

os, indigo 793 ..d 7I. id sr nisololiso d which In oid mow to as modbasIsed roma Is

Flight Path and Attitude of. Vehicle During Ascent Coasting

and Apogee Burning Conditions 21

Position of Vehicle in the Equatorial Plane During Ascent

Relative to the Rotating Earth . . 23

Method for Correcting Eccentricity of Final 24-Hour Orbit. 2.7

Apparent Path of Satellite as Viewed from the

Rotating Earth 28

Coarse Gas Jet System 32

Gas Jet System for Attitude Control and Velocity

Correction in 24-Hour Orbit 32

Schematic of Stabilising Flywheel . . ... • ..... • • • • 33

Orientation of Solar Power Surfaces for Equatorial Orbit . . . 39

24-Hour Satellite Vehicles with Atlas and Titan Boosters . • • 40

Design of the First Added Stage and the Final Stage .of

the 24-Hour Satellite Vehicle with Titan Booster . . 41
TR-59-0000-00604

Page I- iv

Table

1-1

3-i

TABLES

Communication Parameters of Relay Systems

Power Breakdown for Each Phase of Orbit,Establishing

Sequence . . . 30

Page

13

4-i Data on Attitude Control System if Thor or Titan is Used

as Booster 35

4-2 Electric Power Requirements for Interim System (Thor Booster,

Periodically Focused Tubes) 36

4-3 -.Electric Power: Requirements for Initial System (Titan Booster,

Periodically"Focused Tubes) 36

4-4 .Solar Power- System, Interim System 38

4-5 . Solar Power • System, Initial System 38

4-6 . Structural Weight of Titan-Boosted First Added Stage (Using

12,000-lb Thrust F2

 - N2 H4

 Nomad. Engine) . . , • 42

4-7 . Weight Summary for Titan7Booster Vehicle 44

4-8 Payload. Weight Summary (Titan Booster. Equatorial Orbit) • • . 44

111111116 -60fifIDENTIAL

Thsi dearsoat casters Inforrolloo Girdles the adienel delense el the United Stoles irlthla Mr reenlee of the Isolannoo Lows. Nov. 111, m.S.C.. Madsen rn end TIC the renunitalen rendelho el which in any renew to an weeethollerl pones Is erohlbilee by 
TR-59-0000-00604

Page I-1

INTRODUCTION

This volume, the first of four, is a summary of the main results of an

investigation of the feasibility of various applications of high-altitude satellites,

especially satellites in 24-hour orbits. The applications discussed briefly in

this volume and analyzed in detail in the other volumes include:

(1) Communication relay, especially by 24-hour satellites,

ZZZZZ thr ZIWZ/Z/ZAW/17z,

The feasibility of other applications, such as bombing, direct electronic or

photographic reconnaissance, navigation, and guidance, was not investigated

in- this study because other techniques for accomplishing the same objectives

appeared to be more effective or much simpler and cheaper.

Recent work by WAD, the Army Signal Corps, and Space Technology

Laboratories, Inc. in the planning of the ARPA program for the development

of communication satellites was greatly facilitated by the material in

Volumes II, III, and IV, .published in September 1958.

Since the publication of these volumes, the situation concerning the

availability of prospective propulsion stages has changed considerably.

Also, additional studies and analyses have been made in connection with

the ARPA program. As a result, certain detailed portions of this report

are no longer applicable, and therefore this summary volume is confined

for the most part to the more general ideas presented in Volumes II, III,

and IV. Certain corrections developed during subsequent studies are given

in footnotes.

ml, decostoot ossoolos lotormoliso offosSioo Noo Notf000l *Wm of Ow UMW SINN walk Nw 1110•11111. of She isolooloos Lows. MD 111. U.S.C., *mew 793 ossi 794. Nn iroosolssiost osvoiolloo of wIddi to soy stoosor Non ossooliterlasd wow Is osohlhassl frff
TR-59-0000-00604

Page ICHAPTER 1

COMMUNICATION SATELLITES

1.1 LONG-RANGE COMMUNICATION

Present-day systems and techniques for military communications do

not adequately meet military needs for reliability and traffic handling capacity.

The available wire lines have a low capacity and, in any case, they can be

easily put out of action. Low-frequency radio transmission suffers from the

strong atmospheric noise, and the antennas are notoriously inefficient.

Medium frequencies can be used for long-range communication only at night.

High-frequency signals are subject to severe disturbances by various solar

emissions, such as flares; and additional high-frequency absorption, sometimes lasting for several days, occurs in the polar regions. Radio waves in

the lower VHF range can be used for communication, but only at ranges limited

to about 1400 miles. Longer ranges can be attained by series of short-range

links, such as the currently used microwave relay links; but then the multiplicity of equipment decreases reliability, complicates maintenance, and

raises costs.

Placing active, real-time, radio repeater stations in high-altitude

earth satellite orbits will serve to avoid many of these difficulties and consequently will make significant improvements in the reliability and traffic

capacity of military long-range communication circuits. As shown below,

24-hour satellites, moving in circular orbits, are particularly well suited *

for this purpose—especially 24-hour equatorial satellites (which rotate with

the earth and hence retain their positions above preassigned points on the

equator) and 24-hour polar satellites (which move in planes containing the

earth's axis). The altitude of a 24-hour satellite above the earth's surface

is about 19.4 x 10 3

 nautical miles.

* The term "24-hour satellite" is a short name for a satellite having the

pekiod of one sidereal day, which is the period of the earth's rotation

relative to the fixed sitars and is equal to 23 hours, 56 minutes, and

4.09 seconds.

•

This demote, ameba lateneetlee effecting the melee& Mews of Moo Wield Seeks welds I. atoesdne of thebeiseepo Laws, Title 1$. 9.3.C., &Kam 797 end 791. the Inettaisdee or tavola., if wkI My meow te en eseurterivel Nom is bY law.
TR-59-0000-00604

Page I- 3

To determine the best positions of 24-hour commU satellites,

we consider Figure 1-la, in which the shading shows the areas where almost

all of the communication signals of importance to the U.S. military forces

originate and are received. It is apparent that high-density communication

service must be provided (1) between North America-Greenland and Europe,

(2) between North America and Asia, and (3) between the various points in

the ZI. By contrast, the communication service required between the

remaining areas can be a low-density service.

1.2 HIGH-DENSITY SERVICE

Figures 1-Ib and 1-1c show the earth as it would be seen from 24-hour

equatorial satellites stationed above points on the equator at the longitudes

30 •W and 170°W. It is apparent that such a pair of satellites would provide

for the requirements (1) and (2) listed above; moreover, as shown by the

rectangles in these figures, the satellite antennas could have relatively

narrow beams, with a consequent decrease in the required power. A 24-hour

equatorial satellite placed anywhere between the longitudes 45 • W and 130°W

would "cover" all of ZI. Consequently, three equatorial 24-hour satellites

could meet all the high-density traffic requirements.

The main advantages of equatorial 24.hour communication satellites

are: (a) the desired coverage for high-density traffic can be obtained with

only three such satellites, whereas the use of other orbits cannot reduce this

number and will require a substantially greater number if the altitude is

decreased appreciably; and (b) the "stationary" character of an equatorial

24-hour satellite reduces—if it does not completely eliminate—the problem

of antenna tracking, both on the ground and in the satellite; this permits the

use of larger antennas and yields a greater communication capability with

the same transmitted power. (If proper directional antennas are used, the

power required for a satellite depends on the ground area to be covered, but

not on the altitude of the satellite.) As to propulsion, it is shown in Volume II

that even a Thor booster with some modification can put a useful payload into

a 24-hour orbit.

-CONRDENTIAL---

EMI

This usteest sealshts istessustiee Wadies the nolleasi deems* et the %Mei Slain stNitio the staseles el the lispiesese les% Title U.S

a

 .C. ladies 7.3 eed 7114, the heesoiselets es reteluNett et which is any toomer le eu eeeuthestaed peon is seuhibiled W

e
TR-59-0000-00604

Page 1-4

St g 

INC414„,7 pI :44 figgetpiF 4 7 4111M,

Mitit gh;HJI 111U

6

11111*071 40011i111 ,41141

t7.dissaimma, 411111111di r NfrIk

.1123391!ECIFItellElaltdt 13312!!!!!!!V

lima •

1111211111.1111111111111111116:11"`IMMELV.:

ilkb.-AkikiielllItillIC11111111111111111111111F1

M

1111111111111117gftwonite

ossr 1 111EIMIRIMIIII111111116 O 'ValMO I

*NJ AMIN= k

Figure 1-1a. Areas (shaded) Originating Great Majority

of Communications of Military Interest.

Figure 1-. 1b. Longitude I70°W. Figure 1-1c. Longitude 30°W.

Figure. 1-lb and I- 1c. Earth as Seen from Satellite in 24-Hour

Orbit (coverage areas o ed).
TR-59-0000-00604 slam Page 1-5

1.3 LOW-DENSITY SERVICE

The communication equipment of satellites intended for low-density

traffic will be different from that required for high-density traffic. Thus,

a much smaller bandwidth will be adequate, but the power transmitted by

the satellite will be radiated to a substantially larger area. Also, communication with mobile units, such as aircraft, will have to be established, and,

therefore, the dimensions (and gains) of the "ground" antennas will be considerably smaller than those which can be used for the high-density service.

The north polar regions, which are of the utmost importance for the

SAC communication problem, cannot be "covered" by equatorial satellites.

This fact plays a critical part in the choice of orbits for satellites serving

the low-density areas.

• The orbits recommended in this report for polar coverage are 24-hour

polar orbits. The earth track of a satellite moving in such an orbit is a

figure-eight, extending from pole to pole. The northern half of such a track

is shown in Figure 1-2, for the case when the satellite crosses the equatorial

plane at the longitude of 30°W (the satellite crosses the equatorial plane from

South to North and from North to South at the same meridian, because during

the 12 hours which it spends in either the northern or the southern hemisphere

the earth turns through 180°). An important feature of a 24-hour satellite

is that it keeps passing every 24 hours over the same points on the earth,

which reduces the tracking problem considerably. If three such satellites,

crossing the equator at 60°W and spaced 8 hours apart, should be put up,

then at least one of them will be above the 30°N latitude at all times, and the

three of them will provide continuous line-of«sight communication between the

entire United States and certain large areas shown in Figure 1-2 and identified

in the caption of that figure.

The polar orbits described above cover very little of Asia an•Africa,

and miss large portions of the Pacific Ocean. These areas can be covered

by two equatorial 24-hour low-density satellites, one at 170°W longitude,

over the Pacific, the other at 65° E longitude, over the Indian Ocean. However,

the Indian Ocean satellite could not communicate with the United States directly,

but only through a relay station in Japan or Europe.

-CeNFIDERTFAt

This demos' gestates Isfensellos anodise the adleast Mow of Ile Ustled Salo eat& the IIIMINIØe • of the blhigheri lese. Thle IL VAC., Seaga 710 sled 794, tho wessistesies or tovolelles of weigh Is stly ssegrar to se easelhertgod penes Is istoatlatel bs law.
TR-59-0000-00604

Page I-6

Area within continuous line-of-sight of recommended polar satellites

when they are also within line-of-sight of central U.S.A.

so Additional area within continuous line-of-sight of recommended polar

MI satellites.

Figure 1-2. Continuous Coverage Areas for Low-Density

Polar Relay Satellites.

•1111111

decomoof cooforsos lolotosoNoo &foam the Whoa doloalo of Hos UMW Sohn wails dm ~duo of OW Isoioossso Lows, TN% loseetoo M sod TM, She enomotssloso of mobiles of lo any sosooso fo so owoothoslasd mous is ossisIbIled by
TR-59-0000-00604

Page •- 7

1.4 NUMBER OF SATELLITES

Both the high-density and the low-density traffic areas of the earth

can be covered by the following eight

High-Density Traffic

24-hour satellites.

ZI to Europe Equatorial, Longitude = 30°W

ZI to Asia Equatorial, Longitude = 170°W

ZI Internal Equatorial, Longitude 45°W to 130°W

Low-Density Traffic to Aircraft, etc.

Polar, North America, North Atlantic,

Europe, North Africa, and portions of

North and South Pacific Ocean, South

Atlantic Ocean, and South. America

3 Polar Satellites crossing Equator at 60°W

. Pacific Area 1 Equatorial, Longitude 170°W

Asia, Africa and

Indian Ocean 1 Equatorial, Longitude = 65° E

Note that at 170°W both a high-density and a low-density satellite are

required. One double-purpose satellite could be used instead, provided that

an adequate payload could be put into a 24-hour orbit.

1.5 BANDWIDTH REQUIREMENTS

To establish and maintain a real-time satellite relay service is a

complicated undertaking; therefore, it is important to use a conservative

design and, if feasible from the standpoint of weight, to set up an extremely

high-quality service (one with a high signal-to-noise ratio). A high signalto-noise ratio will also reduce the susceptibility to enemy jamming, although

jamproof operation will require, in addition, some technique for spreading

the signal over the total available bandwidth to prevent the enemy from effective spot-jamming. Therefore, in this study the peak signal-to-noise ratio

MIN -6041413ENTAL

deerwest paean Ishwasdies ModShe waisssi dolmas st Nos WWII kakis Wilds Ohs swaskes of Iles Weems laws, This IL I...LC., ladles 793 wool 794. Ss ity mobiles of whisk Is losy *wow is us vsesihssiami mows b wwbbilbsi bY
TR-59-0000-00604

Page 1-8

-CON-9-DENTIAL

was set at 30 db, except for using a 20-db ratio in cases where the power

(and satellite weight) required to provide a 30-db ratio for acceptable bandwidths would be excessive. The 30-db figure permits transmitting high-quality

video and voice information, while providing a negligible error rate for normal

digital signals, such as teletype. For the low-density relays, where the Z0-db

signal-to-noise ratio must be accepted, the limited bandwidth will be inadequate

for transmitting video information, but the relays will still be comparable to

wire-lines for transmitting digital signals.

For an excellent high-density service, a bandwidth of 50 megacycles

(25 megacycles each way) between the terminals of an intercontinental relay

is desirable. This bandwidth would provide, for example, for 5 video channels

and 625 voice channels in each direction for a two-way system. Thirteen or

more teletype or data-link channels could be substituted for a single voice

channel.

For a minimal low-density traffic service the video is unnecessary, and

there is no need for more than 5 to 10 per cent of the capability for voice,

teletype, or data-link signals required for the high-density traffic. Therefore,

the bandwdith for a low-density system needs to be only 100 kilocycles, which

will provide either 5 high-speed data-link channels, or 25 voice channels, or

over 300 teletype channels, or some combination of these channels.

The choice between traveling-wave tubes and klystrons as amplifiers

for the low-density service will be determined by the availability of tubes

and the desire to minimize the total power required for the system.

The bandwidths suggested above are merely design objectives; in fact,

it may prove desirable to reduce the system signal-to-noise ratio to obtain

increased bandwidth, or to increase reliability by providing redundant

equipment.

1.6 POWER REQUIREMENTS

The maximum gain achievable with a satellite antenna is determined

by the area to which it must transmit; and since effective satellite antennas

may be limited to about the size of the satellite itself, it follows that the

IMP° AL

This decsoset coMeles infetweise Wadies Hos Whose /skew of Me Unload lodes wild* M, osseelee .f Ow Ispis•ess Lows, Thie

U.i.C., iodise 793 ..d 714. Ih. kenswissise a ressielion of which Is sow meow ta en onaufberlard ammo Is evabIblIall by law.
TR-59-0000-00604

Page 9

highest possible operating frequency should be used to minimize the antenna

packaging problems. However, because amplifier tubes are not available at

frequencies above 12 kmc, the frequency of 10 kmc was selected in this study

for. determining the power requirements.

Polyrod antenna arrays are the most suitable type of antenna for the

satellite, where limitations upon size prevent the use of the larger, but more

efficiently fed, parabolic dish antenna. The total weight of the largest array

considered (16 rods, X./2 in diameter, and 10 X. in length) should not weigh more

than 10 to 20 pounds at S-band or 1 pound or so at X-band.

The "size" (equivalent area) is the only parameter of the ground antennas

affecting the power requirements for satellite-to-ground transmission; and,

therefore, by the reciprocity theorem, only the size of these ground antennas

affects the power requirements for ground-to-satellite transmission.

1.6.1 Z1 Relay 

The West-to-East angle subtended by the United States at a 24-hour

satellite placed at about 95°W longitude will be about 4 degrees; the North-

.

to-South angle will be about 10 degrees. At 10 kmc this coverage would

require an antenna array of only about 5 inches by 10 inches by 1 foot or so

in depth, and would imply a gain of about 30 db with about a 50 per cent

radiation efficiency. *

 The ground antenna would be a 60-foot dish with a

bearawidth of 0.1 degree, which is compatible with the tracking accuracies

and would provide a gain of about 63 db at 10 kmc. ** The radiation efficiency

of this antenna may be expected to be over 90 per cent.

*

A review of current antenna design practice indicates that, if the assumed

beamwidths are obtained, the gains may fall short of the assumed gains by

as much as 1.7 db. This would imply a decrease in bandwidth capability of

one-third, an increase in satellite t,ansmitted power of 50 per cent, or an

acceptance of a decrease of 1.7 db in signal-to-noise ratio, ** The mechanical tolerances on a dish of this size may prove severe at 10 kmc,

and if the tolerances cannot be met, the assumed gain will not be attained.

However, the conclusions regarding system performance are not* changed if

the frequency is lowered, the decreased gain at the lower frequencies being

exactly made up by the decreased path loss of the propagation path. The

major system change, required by a reduction in frequency would be an increase

in the size of the satellite antenna required to produce the desired beamwidths;

recent studies on antenna structures suitable for satellite vehicles indicate no

severe probleras in the production of lightweight foldable antenna structures

having the required characteristics.

a LI

This decease eardeles infenwellee effectbre fbe aellamel Meese el the Wiled &Nee edible ire meeelsaalp of the Illeriee•pe Lowe. MN 1$. 10.11..C. Soaks 713 sod 794. the tremarlselea M neyelebee whkb le env wwintal he e" easellmodyed amen Is 'rehabbed by law.
R-59-0000-00604

Page I-10

With a 30-db signal-to-noise ratio, the ZI communication relay service

would require 5 watts per megacycle, or a total of 250 watts for a 50-megacycle

bandwidth. Since there is little limitation on power for the ground equipment, it

id advantageous to transmit at higher power levels from the ground, so as to

decrease the amplification required in the satellite.

1.6.2 Intercontinental Relays 

For a relay between eastern North America and Europe, antennas having

beamwidths of about 2 degrees by 9 degrees are required. An antenna of about

6 inches by 2 feet by 1 foot or so would provide a gain of about 33 db. If the

ground antennas are 60-foot dishes, and if the signal-to-noise ratio is to be

30 db, then a transmitter power of about 250 watts is required for two-way

transmission of signals having a 50-megacycle bandwidth.

A relay between Asia and western North America requires similar

antennas, with a third antenna, if desired, to serve Hawaii. A bearawidth

of 1 degree will cover Hawaii; it can be provided by an antenna of about

4 feet by 4 feet by 1 foot without complicating unduly the attitude stabilization

problem of the satellite. With a 30-db signal-to-noise ratio, the necessary

transmitter power is about 10 watts. Hence, the total power required for

this relay is about 260 watts, for a bandwidth of 50 megacycles.

1.6.3 Low-Density Traffic 

For the low-density polar service, shown in Figure 1-2, the satellite

antennas must have 16-degree bearawidths, since the total line-of-sight

region subtends an angle of about 16 degrees at a polar 24-hour satellite. The

necessary parabolic dish has a diameter of only about 5 inches and a high

radiation efficiency. Since the attitude of the satellite will be stabilized,

fixed antennas may be used to obtain the desired coverage. However, the

ground antennas must track the relay satellite. In addition, "ground" stations

may include aircraft, and such mobile antennas will be small. A 15 by 15.by

13 inch aircraft antenna will have a beamwidth of about 3 degrees at 10 lane,

which is wide enough to permit tracking from aircraft and which will provide

an antenna gain of about 33 db.

-CeriFiDENTIAL

111111111111

7W. Iloomoot cootaios btfonoolloo oftedlos tho oollood Mame et the Who/ bass wale tie monies al the fooktooso Lows. This le, U.S.C.. Sodloo 712 owl rm. the troomisslot amebas' at which lo way wow to so ostoothottool woos h orolabliod low low.
-59-0000-00604

Pagel-11

Under these conditions, 2.2 kilowatts are required to provide .a 30-db

signal-to-noise ratio in the 100-kilocycle bandwidth assumed for the lowdensity service. But this power may be reduced to 220. watts if the signalto-noise ratio is reduced to 20 db. If more power is available, a lower

signal-to-noise ratio could also be used to increase the bandwidth.

The power required for a • low-density 24-hour equatorial satellite

placed over the Pacific or over the Indian Ocean proves to be essentially the

same as that computed above for a relay moving in a polar orbit.

1.7 "INTERIM" COMMUNICATION SYSTEMS

The satellite relay system described above is an initial global system,

in the sense that, as the achievable payloads increase, it would be replaced

by a higher-capability system. However, the time required for developing

even the initial system is around four to five years, and hence it is useful to

investigate some interim systems, which would have a lesser capability but

could be developed in half that time by making full use of the presently

available devices and techniques.

Two general classes of "interim" systems have been investigated: (1)

systems providing relay service from point to point or, at most, from a

restricted area to a restricted area, and having a bandwidth of a few megacycles; and (2) systems for relay service between any two points within a

large area with a bandwidth of a few kilocycles.

Three point-to-point systems have been investigated, which provide a

reasonable range of capability and from which the requirements for this

class of systems as a whole may be estimated. They include: (a) point-topoint simplex (one way at a time, but reversible) transmission of FM multiplexed signals with a nominal information bandwidth of 5 megacycles;

(b) point-to-point:duplex (simultaneous two-way transmission) of FM multiplexed signals of the same nominal bandwidth; and (c) simplex transmission

from a restricted area to a restricted area, with a nominal total bandwidth

of 2.5 megacycles.

CONFIDENTIAL

11111111

This fitoomoof asodains imierseetiew Medias the isolishol aims@ of she WWI I* .s whhie the OSSIIMIII. Of M. 1.9144.40 laws. Tiii4 111. lectise 733 sod 794. she humsssississi sr rimmistisio of which la see oweraloroberl rworeme le mo►lhifool he km.
-0000-00604

Page .1-12

One relay system for service between points within a large area was

studied. Such a system should permit communication between mobile stations,

including aircraft. Therefore, voice transmission is desirable, although a

high-speed teletype capability might be acceptable. The system that was

investigated uses the 200 to 400 megacycle UHF communication band, with a

bandwidth of 3 kilocycles. Dipole antennas are used to eliminate the need for

antenna tracking, with the result that the UHF equipment now available for

aircraft communications may be used in its entirety, and new development

is necessary only for the satellite-borne equipment. The UHF voice relay

would be similar in its main features to the ZI relay for the initial systems.

However, the UHF communication band permits the use of more conventional

amplifier tubes and design techniques. The estimated weight and power

requirements for the UHF voice system are about 20 pounds and about Z58 wafts.

Some parameters of the initial system are compared with those of

an interim system in Table 1-1,

CONFIDENTIAL

decemeat seetelee lateasellea sifectlee the eetleeel deisese et Ike Wiled Steen HMIs the weestiee et the Iteleeette lows. Title U.S.C.. Sedge, 713 sod 714. the Ivommeleelaa ar weelmtina al mrldrls 81. nay wow.. Mt am m.....►.4e•A wawa. I. swahilarrl I. lem
Interim, UHF Voice I Initial

To aircraft over Polar Regions, Northern

Oceans, Europe, North America

Continuous Coverage

(3 polar, 1 Pacific)

Bandwidth

Signal-to-Noise Ratio

Satellite Antenna

Required Transmitter Power

Total Satellite Communication

Power

Satellite Communication Weight

Ground (or aircraft) Antenna

dipole

135 watts

258 watts

20 lb

dipole

5-in. dish

220 watts

1725

*

 watts

73

*

 lb

15-in. sq array

3 kc 100 kc

15 db 20 db

— 59-0000-00604

Page •I-13

Table 1-1. Communication Parameters of Relay Systems.

Low-Density Traffic

High-Density Traffic

Interim Initial

Greenland to

• near East

Russia to

Spain

50 mc

30 db

Typical Coverage of One

Antenna

Bandwidth

Signal-to-Noise Ratio

Britain

5 mc

20 db

6-in. by

2-ft array

250 watts

Satellite. Antenna

Required Satellite Transmitter

Power

*

Total Satellite Communication

Power

*

Satellite Communication Weight

Ground Antenna

* Using periodically focused tubes.

1760 watts

rnxri

4 ft sq: array

5 watts

210 watts

86 lb

30-ft dish

100 lb

60-ft dish

11111111P TM. dearoset strieles leformstleo sifinflas Ma *Wood Meow el Ow UMW kehn within She emovele. el No Ilsokirase Lows, Title

1$. u.s.e. two. 7t% mut Ift Naa 411111110.11111eds. venellaNa. .8 Is ow^. abonembsor ....am to ...141.11.4 Iwo Yew
004

TR-59-0000-00604

Page .1-14

V

I

=le

Thu Mareetit sattel's looleneadas offedlove Ilse mullemel deka, the WIWI OWN %Able Nes ...Ma" el ft Itsolater. 111r. !It*
-ergr-milks9-0000-00604

Page I-15

lho doon000f soofolos kloosoffoo Machos floe ..flood I* n••••••• ow. —• ••n•• . • .•..•.. ..flooddolma of Mo UMW *otos WAIN •• • 11•• •

• •• •

inoolos of

•

Iholeitor 

• 

lAnn. 

SO

TIM•
0000-00604 Page

F11/77/1//,"

1031°

,re
zz,zzezziy,

A

11111111116MENfik-59-0000->00604

Page I-17

sisr

-----"

zrzezz—Aw --",rizzz,/7//

f'

,',,,,,#,,,,,,,,,,,,,,A44466466666!4444444464 ",,,,,,,,,A4,444,4,4444644646666666666666666660566664 , ,i0;6/

di

This Meret emetehts Intentetteit effecting the aelleml &teem if the UMW Oulu 

SIM

Willis the ottesiee et the Esttkeer! 4". 11fie
TR- 9-0000-00604

Page 1-18

CONFIDENTIAL

1111111111

decenerd eenteles intenietlen ettecting tite milks& detente et M. United Steles Wilde 1W neettine ei betenette Lone. Me
CHAPTER 3

TR-59-0000-00604

Page 1-19

ASCENT TO 24-HOUR ORBITS

3.1 ASCENT TRAJECTORIES

Equatorial 24-hour orbits. The trajectory recommended for ascent from

a hypothetical equatorial launch site is described in Figure 3-1. The first portion

of powered flight begins on the launch stand and ends when the vehicle is moving

horizontally with the speed of about 25,000 feet per second, at the altitude of

about 140 nautical miles. This is followed by a coast in the 140-nautical mile

circular orbit. When the vehicle is at the proper longitude for transfer to the

desired position in the 24-hour orbit, a "perigee burning" adds about 8000 feet

per second to its speed, and sends the vehicle into the transfer ellipse (more

precisely, into the minimum-energy transfer ellipse, called the Hohmann

ellipse). The missile coasts in this ellipse through 1800 , when the "apogee

burning" adds about 5300 feet per second to its speed, and injects it into the

circular 24-hour orbit. The longitude in which the satellite will hover above

the equator is determined by the length of the coast in the low-altitude orbit.

(The terms "perigee burning" and "apogee burning" are used because they

refer to the burning at the perigee. and apogee of the transfer ellipse.)

The sequence of events recommended for launching an equatorial 24-hour

satellite from Cape Canaveral (latitude 28.5°N), illustrated in Figure 3-2, runs

as follows: The vehicle is launched eastward into a circular 140-nautical mile

orbit (inclined 28.5° to the equator) where it coasts until it crosses the equatorial

plane at a point determined by the desired ultimate position of the satellite

relative to the earth. At this crossing point the perigee burning sends it into

a transfer ellipse, whose plane is inclined 28.5° to the equator.. When the

vehicle reaches the equatorial plane (after coasting 180 o

 in the elliptic orbit),

the apogee burning provides .a thrust of the correct magnitude and direction to

inject the missile into its final circular 24-hour equatorial orbit. Comparison

of Figures 3-1 and 3-2 shows that in the case of launch from Cape Canaveral, the

extra speed to be added at apogee for injection into the equatorial plane is

6013 - 4826 = 1187 feet per second. Since Cape Canaveral turns with the earth

MIS

This d....0 eteteles lehttaietleso Oodles the eetteeet Ibionso it the Wiled Steles vats ammlag ut IM 118.191141.1! TM* II I. Ilaalas. Si ' ...a w.aablaIar adol. Is mew mown.. .....•41•164....1 mamma I. nevalahllowl Imo kw.
Id 4.,

ix+ o.

0

0 V

be 0

V .14 "1: V U

ig 0 0 V C

44 #1 4a a met d i

1 1,1! q

0 d 0 pl a,

4 0 :4-,1 4 " t.- Thl -

u

V 0

.0

0,0

to.0

R

C

•

g .re' xi u

a 7 4 g t

41 4.1 d

ft

V 00

41 4

1

e4

:641 joi 0 r-a Ag

AN

14 CO M 0 Nit ti. 1 o

,c; 2 • 14 N 0

ig 42 0 a

Of e o 6 g

b0•0 là to 0

.. .0 ..•

.0 2.41

44 44

PQ

u

.

a

.1 I.-

VI e ,c;

ei «,0 u its Ch

,

.-4.44 4.0 g As, 0 - .4

4 z. > e g

.:4

.o 441 it:

44

0

•

.0

-n

$4 a

r•a a ia 0

.. I;

ea g I) g > g

14 X!

.21:5°

44 u'l

F4

44 og 11." 0 2 .0

4. o u • I

0

4.4 11

co

UMW

-59-0000-00604

Page 1-20

nIimwseft ceateles Infonseilea aftedine $I *WNW defame of Om WWI Sion withi p lie aseashe 44 N. Espkwie4 THIS
-CONFIDENTIAll

TR- 5 -0000-00604

Page 21

mat

Tle BI• Nomad senhilas blenselist offedine II I* . • Von notional deteer of UT United Steles width' Itte Tenni" of tie lisidener Lees. Itne
-59-0000-00604

Page 1-22

slower than a point on the equator, the total penalty is 1187. + 184 = 1371 feet

per second.

If the perigee burning is scheduled for the time of the first crossing of

the equator, the satellite will take up a station on the longitude of Singapore.

Other final longitudes can be attained by longer coasts in the low-altitude

orbit or by departing.from the strictly eastward launch.

While the vehicle ascends in the transfer ellipse, it moves relative to the

rotating earth as shown in Figure 3-3.

Polar 24-hour orbits can be established by firing to the south (that is,

somewhat west of south relative to the ground) from the Vandenberg Air Force

Base. The method of ascent is similar to that in the equatorial case. However,

in a southward launch one gains no advantage from the earth's rotation.

3.2 ASCENT GUIDANCE

Placing.a satellite into a circular 24-hour orbit involves "ascent guidance"

of the satellite into an approximately correct orbit and then making the necessary

"orbital corrections" with the help of observations taken on the ground.

The apogee of the transfer ellipse will always be within a line-of-sight

of the United States or a friendly land area, because its position is determined

by the communication requirements. On the other hand, the perigee of this

ellipse does not always have this property; for example, if the satellite is

intended for communication between the United States and Asia, the perigee

lies over the • Indian Ocean. Therefore radio guidance can always be used at

apogee, but not always at perigee.

By contrast, certain available all-inertial guidance systems are satisfactory

for the booster-powered flight and at perigee; but the present inertial systems

are not adequate for the apogee burning without external corrections, because

of the cumulative gyro drift for the long flight period. Therefore both radio

and inertial guidance must be provided for 24-hour communication satellites.

Guidance during the booster-powered flight is done with a gyro-stabilized

inertial platform, mounting three accelerometers. This platform works together

CONFIDENTIAL

This dowaseet ante's' loshwatetkra effectles tie ieti geal *Aft' el the UMW &eft within Noe nnalthmi ei Ni. 1101•11“11 UM. This

1S. V.S.C.. ladles% m en4 roc tits *own !ashes we tersoiMiwo wi mrla In awe manor.... ow rwr.s.M.04...+1 oases. I by•

dIMM
TR-59-0000-00604

Page 1-23

11111111111

Tles deceseeet cashew etfeneetlea effecting the aesienel defense et Nee Wiled beta wale lee weenies et the ElvisInt L. Tills Is. U.R.r. Asia4law 701 owl TOA No. ..• .
TR-59-0000-00604

Page I- 24

with a "primary" guidance digital computer. During the first-stage booster

burning period, the vehicle follows a programmed pitch attitude and is stabilized

in roll and yaw using the platform as a reference. The primary guidance computer calculates position and velocity throughout this period. Attitude control

and steering instructions for subsequent stages in the climb to the low-altitude

orbit are provided by the primary computer, which can also issue engine start

. and cutoff signals.

The primary computer is relatively heavy, because the computations to be

done before perigee burning are complex. However, the computations required

during and after the perigee burning are simple enough for a rudimentary

digital computer, here called the "steering and shutoff computer.' Therefore,

the primary guidance computer is housed in the first added stage and is

jettisoned before perigee burning. The steering and shutoff computer is

housed in the final stage.

Low-Altitude Coasting. The inertial guidance system can measure the

velocity and position of the vehicle at burnout more accurately than it can

control these quantities. The resultant control errors in altitude and speed at

the point of injection into the low-altitude orbit affect the low-altitude coast

period. and, together with the downrange error at injection into the low-altitude

orbit, cause an error in the downrange posiition of the satellite at the nominal

time of perigee burning. This can be compensated by adjusting the ignition

time for perigee burning. Velocity and altitude errors at the actual time of

perigee burning can be compensated by adjusting the vector velocity increment

to be added at perigee. The perigee correction biases are computed by the

primary guidance computer, and are transferred to and stored in the smaller

steering and shutoff computer before the first added stage is separated from

the vehicle.

After the separation of the first added stage, the vehicle is reoriented

to the attitude required for perigee burning, and is held in this attitude during

the coast in the circular low-altitude orbit. A simple analog computer commands the attitude control jets by converting platform gimbal angles into on-off

signals for the jets.

—CONFtbItvTtitt

TUB decoom andelas lakeinatise 

.•

oltectIng the Neriesoi 

SIM

defams• of Kw UsIllsol Soles wake the lisepoine as boieme. tows; MN IA 11 r fisebsows WI wool WU Ilia a.
TR-59-0000-00604

Page 1-25

Perigee burning adds about 8000 feet per second to the speed, the actual

amount depending on the perigee impulse correction stored in the steering and

shutoff computer, which provides its commands on the basis of the acceleroXXI eter.readings. Steering commands go to the perigee-apogee autopilot.

. During the final two or three hours of the ascent coasting in the transfer

ellipse the vehicle must be tracked from the ground, to obtain its position and

velocity. The tracking data are fed into a ground digital computer, and apogee

burning corrections are sent to the vehicle as modulations on the tracking waves.

Attitude stabilization during ascent coasting is provided by the platform.

During this coast period the vehicle must be reoriented so as to permit the

motor used for perigee burning to inject it into the 24-hour orbit.

Apogee burning is commanded from the ground when the vehicle reaches

the correct angle relative to the earth (the correct latitude for a polar satellite

and the correct longitude for an equatorial satellite). The accumulated measurement and control errors will prevent it from occurring precisely at apogee, or

at the precisely correct altitude. The procedure that simplifies the problem of

orbital correction is to strive for attaining a horizontal burnout velocity that

would lead to a satellite period as close to one sidereal day as possible.

. After apogee burning the vehicle is reoriented by 90 degrees so as to

point its nose-mounted antennae toward the ground stations. This is done with

gas jets, using data from the inertial platform.. After this rough orientation,

a more precise attitude is achieved and the unwanted angular rates are

eliminated by the attitude control system, which must maintain the correct

orientation of the satellite throughout its life.

3.3 ORBITAL CORRECTIONS

•

If the period of an equatorial satellite is different from one sideral day,

the satellite will keep drifting "off station" to the east or west. To permit

tracking by moving only the feed of the ground. antenna it is probably desirable

to limit the total drift during the operational life of the satellite to about

2 degrees; this corresponds to a total drift of about 800 nautical miles along

the orbit, and to a total shift of the boundaries of the "covered region" amounting

CONFIDENTIAL

11111111111

This decoment cont ales lefornestiso etiectioo the notional &Mom of lb. WNW Saba wale the meenimp ei the Eadenatte Torrs. ml.

le, U.S.C.. Statism 713 ood TIM. Ms transmission Sr revolation of which in any warmest to as vosetherised penes is prohibited by Sow.
TR-59-0000-00604

Page I- 26

-CON-FENTIAL

to about 120 nautical miles on the equator. On the assumption of an operational

life of 9 months, this limitation permits an average drift rate of about 0.2 foot

per second. If the orbit is slightly eccentric but the difference between the

altitudes of the satellite at perigee and apogee is limited to 200 nautical miles,

the satellite will undergo . a daily east-west oscillation limited to only one-half

a degree. Similarly, the error in the inclination of the orbital plane is likely

to be quite small compared to 2 degrees. In any case, such discrepancies can

be reduced within one or two days, as illustrated below for the case when the

eccentricity of the orbit is to be reduced.

Consider a satellite injected into a precisely 24-hour equatorial orbit,

but suppose that this orbit is slightly elliptic and the actual altitude at apogee

burning is greater than the nominal altitude, as indicated in Figure 3-4. Since

the angular eastward motion of the satellite about the earth is slowest when it

is at apogee, and since its average angular rate is the same as that of the

earth, the satellite will first recede to the west of its intended longitude. As

the altitude decreases, the apparent angular regression slows up; and 6 hours

after apogee, when the satellite crosses the altitude of the nominal circular

orbit, the westward regression changes to an eastward progression, whose

rate reaches a maximum when the satellite is at perigee. The path of the

satellite relative to an observer on the earth is, to a first approximation, an

ellipse, as shown in. Figure 3-5. • Detailed first-order analysis shows that the

major axis of this ellipse is horizontal and is twice as long as the minor axis.

In Figure 3-5 the point C lies on the desired meridian at the altitude of a

circular orbit having the period of one sidereal day. The point A is the position

of the satellite at the end of apogee burning. Six hours later the satellite is at

B. At this point an outward radial impulse is applied which halves the inward

radial velocity component.. As a result, the apparent motion of the satellite

during the next 12 hours is along the arc BC rather than along the arc BDE.

When the satellite arrives at C, a second impulse is applied, equal and opposite

to that applied at point B. This impulse stops the relative motion.

An error analysis indicates that the amplitude of the north-south oscillation

caused by having the orbital plane inclined to the equatorial plane may not be

This docassoma soefolos isformaloo offoeflaa lbo nablosol dokoeb of Ike Untied 3101es wifbla Has aosabop of do lookassio Laws. TNN fkallos 713 and 714. lb* froasallssioa or rovolotlos of •IlIcle la soy raosassf to as vaataborliod porno Is prolabnall by low.
. . Figure 3-4. Method for Correcting Eccentricity of Final 24-Hour Orbit.

-59-0000-00604

Page I- 27

WIPP GONFIDENT-bok

Ibis domosot costolsis loloroustion offoctioo Ibs notioul *few if Oho United fists strain IM smotelog of Me Isola** Lows,

foeilae 793 soof M. the troomission M woolen* of wlaids Is aisy nssoot to is vasostbotind *ma is stroliblibmi * low.
Position at the End

of .Apogee Burning I •

OW

Meridian

Desired

111111M *•-0000-00604

Page 148

Figure 3-5. Apparent Path of Satellite as Viewed

from the Rotating Earth.

CONFILLi 

This decennia onnolas Iatornatlea aftonint Na noriesini detente et the United Same withia no neenise ei the Ispieneoe Loos. Title IS, U.S.C., lentos 793 sod 794. the tronsadosien or melonsa of whkb hi say nowt to ea vasetheelted person Is pnbiblind by law.
TR-59-0000-00604

Page I- 29

Lel V

large enough to warrant correction. However, if itt ecessary to reduce

this oscillation, this can easily be done by applying an impulse to the satellite

in the north-south direction as it crosses the equatorial plane.

The discussion given above relates to equatorial orbits. The polar

orbital correction problem is somewhat different, because in that case it is

not necessary to control as closely either the eccentricity or the orbital

inclination.

Orbital correcting impulses can be supplied by gas jets controlled by the

steering and shutoff computer, which receives commands through a decoder

from the ground tracker and computer (such a data link.and decoder, associated

with the transponder beacon, is now used on the Atlas missile). Since the jets

cannot be calibrated precisely, a vernier acceleration sensing system must be

used to control the velocity impulses by a feedback loop through the steering

and shutoff computer. The sensitive accelerometers used during the correction

period may remain caged until the apogee burning is finished.

Estimates of certain weights and power requirements are given in Table 3-1

(the actual power consumption is quite uneven, because of the intermittent use

of the components). The weight of airborne guidance in the last stage, required

to localize the satellite above any point on the equator accurately enough for

any of the applications discussed in Chapter a, proves to be about 370 pounds.

Analysis of the accuracy of the system shows that the suggested combination of inertial and radio measurements at appropriate points on the ascent

trajectory would stabilize for 5 years or more both the position and the altitude

of a 24-hour satellite. The vernier velocity impulses required to stabilize the

orbit probably amount to about 100 feet per second total.

The guidance system recom.mended.for polar satellites differs slightly

from that for equatorial satellites, because only broad-beam coverage is

required; that is, after the initial velocity corrections are made, a polar

satellite need not be stabilized against rotation about the axis directed from

the satellite toward the center of the earth.

CONI i IAL

This deemed contelas lefennelke effecting fie eelfeeel defense of the United Steles within the meek* of the eseleempe Um. This IS. U.S.C.. ismiloo 7113 oast Mt, IM troossission or nivolotiso el width be say =Num to as onoolborlassi Hawn is psoblbiled by low.
....

t i.3

co I 1p

t o

8 :N

0D

°

it 0 

... ! .

in o

*

u

.4

•

.

.•

•

•

ri

421 01

•

.....

•

0

oNZ.

. o in

II

VI • •

-

VI b

3 : 2

01 a gn

0 •

in

I, oft

°

Is I

it it

.....

.

•

o o

a

In

I

.

is

. U

11 4

a m

r

I i

ch 1

•

. 3 3 . 'A

IP

'

I 4

.

•

.3•

it g 0

3 t.) P4

; i

e, 0

ii

0 0

o Al

vi

.

•

•

I

I

44

5

14 w

o .s

•

4 g

141;

e

P i

fr.P '

.1A 7

't]i

o

WI C. VI lin t2

iv 1"1-

.

3 :212

•

. .

.

.

an o in 0 0.4 In 43

m N .4 fq N N 

.

N •

•

/

i

I

•

u

I14

2

I.

.

fri do .

he I

Ea. *v

.

11 dial

os

1o3

°

al • f°

.

-; 4

•

li

• ••.

k o 0 14 .

VI, 1 • '01 i

.1 is 

3

e k

lie 1 i.l.

•

. 1 II I... v 2 . 

A .

ld d •

1 g I • I 111 0 .1 .m= ir 1 I

ov 4,1 •28.0 3,

3 0

gle i8 ii1 4.1 ; : 0

•

..; 4 -..; 4 et: 4 at

"4t

M

0

tAd

N

0,••

• 0,•

•

de5 A

N

•

. •

• 

•

•

o

• r

co le

fa on

10

TR-59-0000-00604

Page I-30

This docesisat awlehts laionwaiss anodise the .shoed Mime d ihe United boss whisk 'Ws sweitiws ai 1W lispissege lows. Title

UMW

IS. U.S.C.. Sookom 70 and 794, the treinaissise er rowslaess el which in any meow teas yotoyaborbsd woe is yrebibbal by 15w.
TR-59-0000-00604

Page 1-31

CHAPTER Ptin"M "

. ATTITUDE CONTROL,. ELECTRIC POWER, AND VEHICLE DESIGN

4.1 ATTITUDE CONTROL

The attitude of the vehicle before reaching the 24-hour orbit may be

controlled by the inertial platform and a "coarse" gas jet system, which can

orient the vehicle properly during low-altitude and ascent coasting, and point

its nose toward the earth after apogee burning. This system can also add the

small velocity increments needed for orbital correction; for this purpose the

jet pairs would be oriented for thrusting rather than for torquing. A simple

jet system is shown in Figure 4-1. Finer attitude control during low•altitude

and.ascent coasting is achieved by using the inertial platform and a "fine" gas

jet system.

After injection into a circular 24-hour orbit the satellite should move

with its nose constantly pointing at the center of the earth. The term roll

then means rotation about the velocity vector, nm means rotation about the

vertical, and pitch means rotation about the axis perpendicular to the plane of

the orbit. • A properly stabilized 24-hour communication satellite will turn in

the pitch plane through 360° each sidereal day.

When in orbit, the satellite is stabilized against roll and yaw by a 10- to

12-pound flywheel, whose axis is perpendicular to the plane of the orbit, as in

Figure 4-2. The angular speed of this wheel is kept constant by an electronic

clock. • A small variable-speed wheel, whose axis is parallel to the axis of the

large flywheel, is used to "absorb" small disturbances in the pitch.angular

momentum caused, say, by meteor impacts. If an accumulation of disturbances

should cause the small wheel to exceed its design speed, it can be stopped

(while the pitch gas jets are used to hold attitude) and left free to absorb future'

disturbances in pitch. If the available information on meteor frequency as a

function of a size is correct, then it is likely that the small pitch wheel alone

will be able to compensate for meteor impact and the fine gas jets will have

to be used only during the orbital correction period (24 to 48 hours), when

relatively large accidental torques may result from an unbalance of the coarse

jets.

MS=

This thenseet cookies lefogutotion eflodise the eolioesi defied of the SWIM Rotes within the n ogrolog of the WINN tern. This Sedges 793 wed 794, the trearmeissios or reveldirre of which M dry •eamter to on esoothedeed penes it prohibited by MIN
TR-59-0000-00604

Page 1-32

(a)

Missile A.•.teas Correction

(d)

U Pressure sensor

. Ceara* Jets for

Attila& Control

for position con.rory

His•pressure

Gas Storage

Pressure Gas Jet

Reduction Activation

Valve Command

/Aonopropellant

Storage

41 Gas Storage

Tank

Valves

re-1 rtae AM for

_46.4 Vernier Control

Decomposition

Chamber

'13

Filter

1:1D

ir

Controlled presses. relief

Wives

Peter

Figure 4-1a. Coarse Gas Jet System.

Note:

I. All valves have electric control

Z. Jets rotate about one axis

Figure 4-lb. Gas Jet System for Attitude Control and Velocity

Correction in 24-Hour Orbit.

This assvaset castales afenneties ettectles the .sliced 4111.111111111 Elf the Minted lass. whine Nis measles of the Wisdom Leos. Me

IS. U.S.C.,./octiee 793 sea 754. M. kensraisstes er reveletiee el which in eel 1111•1111111/ Is is sesetherlind penes is prehlighal br
1111111111-GG"414 r. TR-59-0000-00604

Page 1-33

. •

0

This ilscorsoof ambles inforisolloo albedo. is sootiosol defame of the United *Wes valid* smossios of Oa laoloosso Lows. Tido U.S.C, Soclion 713 fad 714, oho troosasissloo ar Novosksfloo of which is soy sionoof to ea oweelsseispi maw Is wslabllod W low.
TR-59-0000-00604

Page I- 34.

The primary purpose of an attitude control for a communications satellite

is to keep the communications antennas pointed at the ground stations. For the

"stationary" equatorial satellite two small monopolise receiving antennas with

a relatively broad receiving•pattern can be used as finders for the relatively

narrow-beam communications antennas. Error signals derived from the

monopulse receiving system can be used to control the fine gas jet system and

the pitch wheel. For a polar satellite it is only necessary to keep one 16-degree

antenna pointed at the earth, and there is no need for yaw control. Pitch and

roll measurements for a polar satellite can be made by an infrared horizon

scanner and can be used to control the fine gas jet system and the small pitch

wheel. Table 4-1 summarizes certain weights and powers required,for attitude

control if a Thor or a Titan should be used as a booster.

4.2 ELECTRIC POWER

Estimates of the power required for satellite-borne communication

equipment were given in Table 1-1 on page 1-13 for both low-density and highdensity service and for both the interim system (which can be built in 2 to 3

years) and the initial system (which can be built in 4 to 5 years). Power

estimates for other equipment are given in Tables 4-2 and 4-3, for the case of

high-density service. Table 4-2 relates to the interim system and a Thor

booster, Table 4-3 to the initial system and a Titan booster. The use of

periodically focused traveling-wave tubes is presupposed in both tables, because

these tubes can save considerable power.

Photovoltaic cells for converting solar energy directly into electrical

energy appear to be able to supply the necessary power. Other devices involve

large weight penalties or require much additional development—as do lightweight nuclear power systems, which should, however, be considered for more

advanced communication satellites.

. Electric power must be available in the satellite as ac, low-voltage dc,

and high-voltage dc. If the efficiency of the inverter delivering the ac is

80 per cent and the combined efficiency of the transformer and rectifier

delivering the high-voltage dc is 90 per cent, the low-voltage power necessary

-CONFIDEM

11111111113

••Ibis dearmssat Patsies Siforsaceiso sanding Ske effillesusl ask** Id dm WWI SOGO•s Ihe OINPahlf el the el11141/01111141 Lass. T111. . U.S.C, &Who 793 wad MI. tromessioloo asvolotios Wilde J. any meow So es foodborlasd poem, Is wroblIshol bow.
TR-59-0000-00604

Page I- 35

-CONFIDEM

Table 4-1. Data on Attitude Control System if Thor or Titan

is Used as Booster.

Component Weight

(pounds)

._ .

Power

(watts)

Coarse Gas Jets (plumbing and main tank) 15

Gas for Velocity Correction (Thor) 45

**

(Titan) 90**

Gas for Reorientation 6

**

20

*

Fine Gas. Jets (plumbing) 5

Gas for Fine Reorientation 0.6**

Gas for Continuous Attitude Correction

in 24-hour Orbit for Lifetime of

5**

One Year

Monopulse System 20 75*

Attitude Stabilization Computer (analog) 15 20

Large Flywheel 12 20

Small Flywheel 2 10

•

.

TOTALS (Thor) 125.6

145 (Titan) 170.6

* Used very intermittently

**

The weight of this gas is based upon a specific impulse (62 sec/lb) of

catalytically decomposed monopropellant. If a higher specific impulse

is found to be feasible, the weight of the gas will be proportionately

smaller. .

-414FIDENTIALTIth doniastal aistelas lehonwollon ethane IN wolieuel iletwwe of1M MIMI Stelae vt1MM lb..e..1. el 1Iw Esphow Law% Mb

UMW

U.S.C.. Seethe 73n3 end 7f4. lie tronsolyeko w rovelelkw el wbich lo any meow le ea emaillearbad paw is preldbllyt by low.
cotifiBrat TR-59-0000-00604

Page 1-36

Table 4-2. Electric Power Requirements for Interim System

(Thor Booster, Periodically Focused

Power Requirements (watt s)

Low-Altitude

Coast

Ascent

Coast

(5.3 hours)

Apogee

Burning

24-Hour

Orbit

Guidance 360 384. 635 3

*

Attitude Stabilization, etc. 20 20 20 30

Coolant Pump 5 5 5 5

Communication 210*

Solar Controls 10

Total 385 409. 660 258

180 watts during first 24 hours, when cozninunication equipment is not on.

Table 4-3. Electric Power Requirements for Initial System

(Titan Booster, Periodically FocuTerdlubes).

Power Requirements (watt sj

Low-Altitude

Coast

Ascent

Coast

(5.3 hours)

Apogee

Burning

24-Hour

Orbit

Guidance 360 384 635 • 3 *

Attitude Stabilization, etc. 20 20 20 30

Coolant Pump (15) (15) (15) (15)

Communications 1760

Solar Controls (20)*

Total 395 419 670 1828

*

If solar cells are used.

111„ U.S.C., Soefloo 793 end 754, the honsoksion or uo doonnoof coototos Infirmities offondag the oolleoof defame of Ho UMW ifolos .Milo Ike onoaloo of Ike ES$8111111111 Lows. Mb

UMW

yolofies of which In soy inonoto to so onooNombod D.N. is problbIfoil by lea.
TR-59-0000-00604

Page 1-37

to provide the 258 watts listed in Table 4-2 proves to be about 350 watts.

During the eclipses of the satellite this power must be provided by storage

cells, which must be charged by the solar cells while they are illuminated.

This circumstance raises the demand on the solar cells to 378 watts for normal

incidence of sunlight, a figure based on the assumption of a charge/discharge

efficiency of 70 per cent and on a study of the frequenciea and durations of

satellite eclipses at various seasons of the year.

The area of the cell-covered surface required to provide the necessary

power depends on the conversion efficiency of photovoltaic cells (which is a

function of the temperature), on the effect of a covering for protection against

meteoric pitting, and on the fact that the available surface cannot be fully

utilized because of various clamps and connectors. Detailed computations

show that, in the case of silicon cells, the area of about 54 square feet is

required to provide 378 watts at normal incidence. This figure is listed in

Table 4-4, together with other relevant data for the interim system. Table 4-5

gives similar data for a satellite of the initial system.

The orientations of an equatorial 24-hour satellite at different. times of

day and of the solar cells at different times of year are pictured in Figure 4-3.

4.3 VEHICLE DESIGN

This study included a detailed examination of the various propulsion and

vehicle design problems associated with the launching of 24-hour satellites

with .a Thor, a Titan or an Atlas booster. Various added-stage combinations

were considered, to determine the payloads as functions of staging, specific

impulse, and structure. The resulting recommendations are given in detail

in Volume IV of this report. The present summary is confined to some of the

results obtained for the case of an Atlas or Titan booster.

Figure 4-4 is a sketch of a 24-hour satellite vehicle boosted by an Atlas

or an uprated Titan (400,000-pound thrust). The two added stages have .a gross

weight of 20,000 pounds at take-off. Figure 4-5 is a preliminary sketch of the

added stages that could be used with these boosters. Two alternatives for the

12,000-pound thrust engine are shown, a pressure -fed engine and a turbopumpfed engine. Table 4-6 presents a structural weight breakdown for the two

—6-61141F09.7141--

This dscosissi ambibu Nhrw.IN. ~ins Shs aalissel (Mime of Um United bobs wiMN liss wessiag of ilso Embosses lows, Tins

Sectia• 793 sad 794, ihe frossosissien vevelsikii of which Is wsy women to as wwwilmiciesd worsen M swehibillool by law.
M-59-0000-00604

Page I- 38

Table 4-4. Solar Power System., Interim System.

Item Weight fib)

Storage batterlea for eclipses

345 watts for 1.3 hours 32.

Solar cells

Supply 378 watts with normal light zr

Incidence (54 sq ft)

Axle. gear train, motor 10

Servocontrols for surface orientation . 10

Inverter, transformer, rectifier 10

Payload components in 24-hoUr orbit 89

Chemical batteries

For low-altitude coast (jettisoned .before

perigee burning)

475 watts' for 10 hours (maximum) 71 • 14 (equivalent

For transfer ellipse (Jettisoned before payload)

apogee burning)

SOS watts for 5.3 hours 40 17 (equivalent

, Total equivalent payload for entire payload)

. electric power system 120

Table 4-5. Solar Power System, Initial System.

Item Weight (lb) '

Storage batteries for eclipees

2498 watts for 1.3 hours 232*

Solar cells

Supply 2737 watts with normal light .195

Incidence 1391 sq ft)

Axle, gear train, motor 72

Servocontrole for aortaee orientation lb

Inverter, transformer. rectifier ' • 30

Payload components in 24-hour orbit 539

Chemical batteries

For low-altitude coast (jettisoned before

perigee burning)

475 watts for 10 hours (maxbnurn) 71 14 (equivalent

For transfer ellipse (Jettisoned before payload)

apogee burning)

508 watts for 5;3 hours 40 17 (equivalent

Total equivalent payload for entire Poiload)

electric power system 570

•

* Based on 70 per cent discharge/charge efficiency and a weight of

SO pounds per kilowatt hour Of input electric power. • •

1.11111112

-CONRDENTikt_

This declaim carobs islemieties Mediae the Wisest defame el the Yelled Oaks will& No sassal et the Wows laws. TINS

U.S.C., bodies 703 end 794„ the treeminies er revelellee el which is say osayow to on sasoNtotioM wool Is lombibbod bY
TR-59-0000-00604

Pagel-39

I

pipi

14

0

0

0

04

14

0

44

gt0

4 •

CI3 .

494

C

o

e

w ad

4 8

0

U)

*44

0 ti)

0 id

0 44

c0

0

71/. dammeat centalas belennehon 019/119.tb ooll000l dolonoo of IM Wiled boars 91619 IM assealse of Ms bpl000po Loom, Title

MOM

Is. U.S.C.. WI/9 793 mod 79/. No 9...014.1n/ .1.1.9iw wow maw... ....1d1.31.1
A

•

.12

to

a

a a

a

,r4

may rv%

gfn

V 4) 0

4

M .

•

4

4.

a

fiv

*14 a

2

TR-0000-00604

Page 1.40

-CONFIDE-PAL

-CONFIDENTIAt

IMMO

ThIs doevelom tamt•Ins offedles the bk... II. WIWI Siolse wield* 9.• vaesalag of the Ilmaessam laws. MU U.S.C.. &M.. 793 Nod 794. do Inosolosion or reveleIlme al which Is goy aniumer eumelIteetabi gems Is prebIblood by low.
14

a • 8

a

•

•

40

eva

•••1

1n4

ty

0

O

•

•

SJ

tad

be

td

A

te

4.0

V

.44

•••

TR-59-0000-00604

Page 1-41

This decoesent ceoteins hifernietioe affecting the oeflasel defense of Ow WOW Steles within the weaning of the 41100000444 lows. TNN It Y.S.C. Sodlois 793 end 794. tho weounissien er reveletloo of which In ear woolior to ON iwoollswized memo is prebibilee by low.
TR-59-0000-00604

Page 1-42

-GWENT*

Table 4-6. Structural Weight of Titan-Boosted First Added Stage

(Using 12,000-lb Thrust F 2 - N2H4 Nomad Engine).

Pressure

Feed (lb)

Turbopump

Feed (lb)

. Propellant Tanks 208 133

Intertank Structure and Thrust Frames 214 178

Pressurization System 110 90

Start Rocket Casings (2) 20 20

Forward Interstage Structure 46 46

Motor (wet) with Flexure and Actuators 129 125

Turbopumps and Plumbing (wet) -- 39

Plumbing 19

Residual Propellants (1 per cent) 115* 115

*

TOTAL 861* 746*

Add 10 pounds for polar orbit.

This decennia enotelas inisnatothsa *Nesting the ..Need &low et the Yelled Seale* within the senselne 

ofe of the IlsoWeed. Lows, Till. 111,"U.S.C., logien 793 end 714, the trantatisslee er revelation of which In any swetwor te an weertberbed wan Is Onadbited be lew•
TR-59-0000-00604

Page 1-43

alternate designs of the first added stage. With slight modifications, this stage

Can be mounted on either the Atlas or the Titan.

The final stage has storable propellants, nitrogen tetroxide and hydrazine;

they are pressurized, and thus boiloff during long coast periods is avoided.

Since weight is not extremely critical in this design, three engines are used,

one 2500-pound thrust engine for perigee burning and two 1250-pound thrust

engines for apogee burning. The use of different engines at apogee and perigee

burning eliminates the reignition problems.

Table 4-7 is a weight summary of the Titan-boosted 24-hour satellite

vehicle, and Table 4-8 describes a possible allocation of the payload suitable

for either the interim or the initial systems when the Atlas or Titan is used as

boosters. • Although the weights allocated to the components are rather generous,

a considerable amount of weight remains unassigned. Guidance weight has been

increased by 100 pounds, to make the specifications easier to meet. However,

it is likely that the weights of various components will in general decrease in

time, and consequently the figures given in Table 4-8 are very conservative.

.....CANWID6r14AE :"

This docesteot eenteles lotervatehte effeetine the nothteel dame of the UMW Stew width' Nn smoseles of the ludowere lows. TItie

U.S.C. Soctloo 747 sod 794. Mn tronsadsslon er teveleden of whidt is art etemeer to es oneritherfeed woo Is prohibited he
TR-59-0000-00604

Page I- 44

Table 4-7. Weight Summary for Titan-Boosted Ve I

Equatorial Orbit

(Cape Canaveral)

(pounds)

Polar Orbit

(Vandenberg AFB)

(pounds)

First Added Stage

Initial Weight 20,000 20,000

Propellant Consumed 11,514 12,505

Burnout Weight 8,486 7,495

Structure and Residuals

Jettisoned 861 871

Final Weight 7,625 6,624

Final Stage

Batteries Jettisoned 71 12

Initial Weight at Perigee 7,554 6,612

Propellant Consumed 4,105 3,593

Burnout Weight at Perigee 3,449 3,019

Batteries Jettisoned 40 40

Initial. Weight at Apogee 3,409 2,979

Propellant Consumed 1,519 1,123

Burnout Weight at Apogee 1,890 1,856

Structure and Residuals 586 577

Gross Payload 1,304 1,279

Table 4-8. Payload. Weight Summary (Titan Booster,. Equatorial Orbit).

Interim

(pounds)

Initial

(pounds)

Payload Capability 1304 1304

Communications 176 220

Guidance 470 470

Attitude Stabilization 70 70

Vernier Jet System TO 70

Power 135 201

Cooling 18 51

Total 939 ' 939 1082 1082

Unassigned 365 222

•-r,

; 44s.

Tbk thmemeet ceeholas lehoweellee 4/tedium the selistel 4111161111111111 of tie Uoitott *des Wahl the wommtiog et Me Espleemp• Lew*, TIN,

U.S.C.. &Wiwi 793 wed 794. the beesaddieso et meddles of which ht elm reaseer to es etteedwind penes Is ptehlblted by low.
TR-59-0000-00604

Page 1-45

DISTRIBUTION

Copy No. 

1-40 R. A. Park

CONFIULii

One

TISs Macaws amislos IshoosIlss sfieelsos detsme of So Wiled Skates with% Ike Nossiss of lie Seisesss LOW% WOO IS, U.S.C., Ssill.. 713 sod 794, On itellealettell et revolsnes tof which Is say war I. on emmalmorlaed gismo b Prehlblasa bY