<<

The Space Congress® Proceedings 1982 (19th) Making Space Work For Mankind

Apr 1st, 8:00 AM

The Expendable and Development

Harry J. Masoni Project Manager, Systems Engineering Laboratories, Space and Group

Follow this and additional works at: https://commons.erau.edu/space-congress-proceedings

Scholarly Commons Citation Masoni, Harry J., "The Expendable Vehicle and Satellite Development" (1982). The Space Congress® Proceedings. 4. https://commons.erau.edu/space-congress-proceedings/proceedings-1982-19th/session-4/4

This Event is brought to you for free and open access by the Conferences at Scholarly Commons. It has been accepted for inclusion in The Space Congress® Proceedings by an authorized administrator of Scholarly Commons. For more information, please contact [email protected]. THE EXPENDABLE VEHICLE AND SATELLITE DEVELOPMENT

Harry J. Masoni Project Manager, Systems Engineering Laboratories Space and Communications Group Hughes Aircraft Company El Segundo, California

of ABSTRACT missions are those which avail themselves the satellite's stationary or sweeping view The role of the expendable (ELV) of the to economically get a global look in the development of satellite systems is at portions of, or the entire, earth's surface. an outgrowth of the bal­ Definite advantages are thus afforded for explored. Essentially of listic missile which was first used for mili­ vital earth tasks such as: communications tary applications 4 decades ago, the ELV is wideband information via high frequency line- used for a wide spectrum of peacetime appli­ of-sight links, large area photographic cations including the launching of earth observations applicable to , high which provide inter- and resolution photography for assessment of orbiting aid intra-national communications, large area visible earth resources, and photography for weather analysis and predic­ for aircraft and ships. tion, and high resolution imagery capability for earth resources evaluation. The launch It is possible to define any of the above mis­ constraints and considerations which the sions in terms of a generic con­ satellite designer must address are reviewed. figuration, as depicted in Figure 1, which consists of a basic spacecraft bus and mission A survey of three developed by Hughes . The spacecraft bus functions to: Aircraft Company for launch by expendable 1) establish and maintain a desired and boosters is included. The HS 376 series of attitude following insertion into an approxi­ communications satellites are capable of mate orbit by the expendable launch vehicle, transmitting digital data, digital and analog 2) command and control the spacecraft and voice, and television in a typical geosynchro­ retrieve bus status and payload data via radio nous orbit application. The Geostationary link, and 3) provide power, thermal control, Meteorological Satellite (GMS) records and and mechanical support for the payload. transmits visual and IR images of the entire earth. The multispectral scanner (MSS) is an Communications payloads typically consist of imagery system used aboard the low altitude any up- and downlink radio system wherein LANDSAT spacecraft which creates images in wideband information transmitted via radio four spectral bands for evaluation of earth waves from an earth terminal is received, resources. amplified, and routed to a distant earth ter­ minal. There it enters into the local ground INTRODUCTION network. * Information transfer occurs in near real time. Payloads for non­ The development of advanced electronics and communications type spacecraft receive imaging materials since an expendable or field/particles information and transmit launch vehicle successfully placed a satellite this data to mission control. This informa­ into earth orbit over 2 decades ago, has given tion may be stored on board the spacecraft rise to a vast array of spacecraft missions bus and transmitted at an appropriate time. designed to benefit mankind. The objectives missions fall into two gen­ The expendable launch vehicle (ELV) is a key of these peaceful space­ eral categories: science exploration and element in the design evolution of every earth applications. Scientific missions deal craft system. It dictates the shape and vol­ with the uncovering of new knowledge of the ume within which the spacecraft must fit and earth and its environment, the , the maximum allowable weight which can be and the stars that cannot be obtained via lofted into the desired . The earth-bound instruments. Earth applications acoustic noise and mechanical vibration envi-

4-16 ronment produced by the launch vehicle largely harness not only ties together these PAF sub­ establishes the design criteria for the space­ systems, but provides the flow of power and craft structure. signal transfer between the ground and the spacecraft during prelaunch checkout. This paper is confined to the earth orbiting satellite and its interplay with the expend­ Certain satellite missions require spin sta­ launch vehicle. Due to its broad use and bilization during operation subsequent to application, the geosynchronous orbiting separation of the spacecraft from the launch satellite mission will be used as an example vehicle. This spin rate can be supplied in developing the role of the ELV and its im­ either by a spin table on the launch vehicle's pact on the . The discussion upper stage or by a system on and methodology, however, can be easily board the spacecraft bus. Of course, the extended to nongeosynchronous missions. A latter technique adds weight, cost and com­ survey of three satellites is provided in the plexity which the satellite designer must, hope that the reader will develop an apprecia­ once again, be cognizant of early in the sys­ tion for the wide range of applications tem development. afforded by satellite technology. TRAJECTORY PLANNING LAUNCH VEHICLE INTERFACES Different satellite missions often require The mechanical interface between the space­ different orbit plane orientations and craft and the launch vehicle represents the periods. Some examples are shown in the brief single most important constraint on the space­ survey presented in the later sections of this craft design. The interface is depicted for paper. For purposes of identifying the a three stage launch vehicle in Fig­ interrelationship of the expendable vehicle ure 2. For every launch vehicle and fairing, and the spacecraft design for this important a payload dynamic envelope is defined by the mission phase, the earth synchronous, or geo­ launch vehicle supplier. The dynamic envel­ stationary, orbit will be used as an example. ope is the volume in the fairing within which the dynamic motions of the spacecraft must be The is one which affords contained while the fairing is attached during the satellite a stationary position in space the boost phase. The launch vehicle supplier as viewed from the earth. The orbit must must calculate the maximum excursions of the therefore lie in the equatorial plane. It fairing in order to define this volume. Snub- must be circular to eliminate any retrograde bers against the fairing may be required to motion. Finally, it must have a 24 hour hold these excursions to acceptable levels. period. The synchronous alti­ Various windows, some removable, are required tude is approximately 35,800 km. in the fairing for access to and inspection of the spacecraft. The spacecraft designer The launch vehicle is a multistage , must use the launch vehicle dynamics and it places the spacecraft into a coast environment to calculate the dynamic motions transfer orbit that has an apogee at synchro- of the spacecraft. nouse altitude. The transfer orbit is inclined relative to the equatorial plane. The importance of understanding this inter­ Perigee velocity is approximately 1524 m/sec face can not be emphasized strongly enough. and a final boost velocity of 1829 m/sec is It relates, for example, directly to the required to circularize the orbit and cor­ allowable solar panel area of the spacecraft rect inclination due to off- launch. and hence, its power-producing capability. Available power is directly related to system The most weight efficient way to achieve syn­ performance for many mission payloads. The chronous orbit is to have the spacecraft per­ designer must be aware at the inception of the form the apogee orbit injection with a high design process whether or not deployments of thrust solid , often referred to antennas, solar panels, or scientific instru­ as an (AKM). This allows ments are required to mitigate the size and the smallest launch vehicle for a given in- shape constraints imposed by the launch orbit spacecraft weight, but the spacecraft vehicle payload envelope. must cope with large offset thrust pitching moments produced by the solid rocket. These A payload attach fitting (PAF) attaches the perturbing forces can be readily attenuated by spacecraft to the propulsive third stage, spinning the spacecraft along an axis paral­ depicted as a solid rocket motor for the lel to the thrust. In some instances, a Delta application. The PAF also provides a liquid propulsion system may be used in place platform for mounting the electrical, mechan­ of the solid rocket. The attendant lower ical, and ordnance equipment needed for third thrust results in greatly reduced perturbing stage operation. These include a forces. subsystem, a sequencing subsystem, and the spacecraft separation system. An electrical 4-17 A less efficient method of apogee injection, chronous transfer and the synchronous equa­ but one that requires no spacecraft control torial missions. The velocity, VPARK, ° during the transfer orbit, consists of using 185 n.mi. parking orbit is about 7800 m/sec. a launch vehicle upper stage with multiple The velocity increment required to inject from burn capability. The first burn is done at this orbit into the synchronous transfer perigee to achieve transfer orbit injection orbit is about 2440 m/sec, so VQH fo r the syn­ and a second burn is performed at apogee for chronous transfer orbit is about 10,240 m/sec. injection. The upper stage For the synchronous equatorial orbit, an addi­ can then be ejected, freeing the satellite tional velocity increment of 1830 m/sec is for normal mission operations. required to inject from the inclined transfer orbit. Thus the total impulse velocity needed Both techniques are illustrated in Figure 3. beyond VP/\RK is approximately 4270 m/sec for Since the expendable launch vehicle is the a V CH of 12,070 m/sec. most expensive element of the system and its price increases monotonically with spacecraft Note that for the synchronous equatorial mis­ weight, most synchronous orbit spacecraft sion, €3 is 24. For interplanetary missions, carry their own apogee kick motor. In the 63 depends heavily on the transit time; the entire stable of current expendable launch shorter the transmit time the greater €3. , only the Titan IIIC possesses the Direct flight to and , for example, capability of placing a satellite directly requires a 63 of 9 and 12, respectively, for into geostationary orbit. Note from the fig­ transmit times less than 1 year. ure that the spacecraft must be able to reorient its attitude via onboard propulsion PAYLOAD WEIGHT OPTIMIZATION independent of the technique used. Maximization of in-orbit payload is a critical PERFORMANCE CONSIDERATIONS performance tradeoff analysis which the satel­ lite system designer must conduct. For a For every satellite mission, there is a veloc­ geosynchronous mission, this tradeoff involves ity that is characteristic of that mission. matching the spacecraft's apogee impulse capa­ This "characteristic velocity" is not single bility with the launch vehicle capability. valued, however, and many different missions can have the same characteristic velocity. Assume the launch vehicle is required to The importance of this parameter is that is inject a payload (spacecraft + AKM) into syn­ allows the launch vehicle performance to be chronous transfer orbit, relying on the AKM characterized by a single curve. Such curves to complete the injection into geosynchronous are very useful for the spacecraft designer orbit. The technique for maximizing the pay- in the preliminary design phases. The defini­ load is illustrated in Figure 5. Synchronous tion of characteristic velocity, VQ|_|, is transfer orbit payload is plotted against given as: launch vehicle perigee plane change (or trans­ fer orbit inclination). This performance is easily calculated and is also provided by the "CH V £AV launch vehicle supplier. One can also com­ PARK pute the AKM capability translated back to the synchronous transfer orbit payload. As where VP^RK is the circular velocity in a the perigee plane change executed by the given parking orbit and ZAV is the sum of all launch vehicle increases, the required apogee the velocity increments required to establish plane change decreases and, hence, the allow­ final orbit. For high energy missions (e.g., able payload weight increases. If the AKM earth escape for interplanetary propellant loading is just right, the AKM and flight) an energy parameter, C^ 9 is used. It launch vehicle capability curves meet at near is related to characteristic velocity by the zero perigee plane change ^and the payload expression: weight is maximized. If the AKM is oversized, the velocity increment will inject the space­ craft into the wrong orbit. In such cases, = V the launch vehicle is required to run an inef­ CH ficient boost trajectory to effectively expend PARK the excess energy. If the AKM is undersized, the proper orbit can only be obtained by reducing the payload weight and the launch where K is the earth's gravitational constant vehicle must be programmed to supply the •and Rp/\R« is the parking orbit radius. matching perigee plane change. As one might expect, AKMs are rarely perfectly matched Typical expendable launch vehicle perfor­ because of the cost of developing new motors mance is shown in Figure 4 where spacecraft and changing launch vehicle capabilities. weight is plotted against a multiple scale which includes 63 as well as characteristic velocity. As an example, consider the syn- • 4-18 Since every extra pound that can be put in electronic links between the many islands of orbit can reap substantial rewards in system and ASEAN countries by Perumtel; performance, the payload weight optimiza­ TELSTAR 3, American Telephone and Telegraph's tion process is critical. The satellite sys­ new long distance high speed digital and tem designer must thoroughly understand the video system; and launch vehicle performance capability in order , dedicated to the distribution of cable to carry out the optimization process. television for Hughes Communications, Incor- portated. Satellite system lifetimes range SATELLITE SYSTEM SURVEY from 7 to 10 years. The three satellites discussed are represen­ The basic HS 376 communications payload con­ tative of the various payloads developed by sists of two polarization selective reflectors Hughes Aircraft Company for launch by expend­ which share a common aperture and a channel­ able launch vehicles. In each case the satel­ ized, single conversion repeater. Incoming lite designer has had to confront and solve (uplink) signals are collected and amplified the problems discussed earlier. He has had by the antenna where they are routed to wide­ to constrain critical satellite dimensions to band receivers that supply additional gain fit within the launch vehicle payload envel­ and downconvert to the desired transmit fre­ ope, sometimes incurring increased cost and quency. The signals are then chanelized by complexity as a result of the need for deploy­ an input multiplexer. Each channel provides ments and improved efficiency. He phase and amplitude equalization, gain con­ has had to optimize the boost trajectory to trol, and high power signal amplification. eke out every last ounce of available pay- The signals are then recombined in an output load weight and still, in some cases, suffer multiplexer prior to transmission via the reduced lifetimes or have to develop new, down!ink antenna. light materials. He has had to construct com­ plex spacecraft structural math models to Shaped transmit and receive antenna coverage ensure that the spacecraft will survive the is achieved through multiple feed horn tech­ severe loading produced by the boost environ­ niques. The front reflecting surface is ment. All these efforts are recurring themes horizontally polarized and is RF transparent in the development of a satellite system. to vertically polarized beams which are bounced off the rear reflector. Frequency To discuss the design process as it relates to reuse is afforded via polarization and/or the impact of launch vehicle constraints on spatial diversity. any one of these satellites would be beyond the scope of this paper. Instead, the designs Although the main part of the HS 376 space­ are presented simply to expose the reader to craft body spins in space to maintain gyro­ the diverse applications of satellite systems. scopic stability, the antenna and communica­ tions equipment shelf are despun to permit HS 376 Communications Satellite the antenna to point consistently toward its target on earth. Hughes 1 newest spacecraft for satellite com­ munications, designated the HS 376, is shown Every HS 376 has its own solid propel!ant in Figure 6. Its concept developed out of a apogee motor. Originally matched for launch need for more powerful satellites for which aboard the Delta 3914, subsequent growth in a choice of launch vehicles could be made. capability of the Delta family has resulted While most HS 376s currently developed are in an undersized apogee motor for most cur­ destined for launch aboard the Delta 391O/ rent missions 3920, the spacecraft can also be launched on the . An outer solar panel that The HS 376 spacecraft is a classic example of in space to nearly double the solar design innovation that has sought to mitigate power generating capacity and a folding, space the constraints imposed by the launch vehicle. saving antenna system afford cost-effective The use of deployments, light and strong com­ launch flexibility. posite structures, and increased solar cell efficiency has afforded the HS 376 a communi­ A wide variety of payloads have been ordered cations capability exceeding that of earlier for both C band and K band communications tasks spacecraft launched from larger, expen­ including; C and D» covering the most sive expendable vehicles, populated areas and all of for Tel sat Canada; SBS, offering K band communications Geostationa^ to business customers in the continental with weighted beams for the Designed to pinpoint and photograph dynamic most densely populated areas; Westar IV, V 9 weather patterns over the Pacific » the and VI, with which will serve Geostationary Meteorological Satellite (QMS), customers in the continental United States, shown in Figure 7, operates at 140 E longitude Hawaii 5 and Puerto Rico; B, providing over the equator, directly south of Tokyo, 4-10 . The satellite has a key role in the Also carried on GMS-1 is a space environmental World Weather Watch program, described by monitor (SEM), developed by Nippon Electric meteorologists as the most detailed study of the earth's atmosphere ever attempted. The Company, which is used to investigate ener­ effort involves the United States, Japan, the getic particle activity emanated by the , the , the and assist in studying the effect of extreme , the World Meteorological solar activity on earth's Organization, the Internal Council of Scienti­ systems. The SEM detects solar protons, fic Unions, industries, and about alpha particles, and solar electrons. 145 nations that contribute daily surface and atmospheric measurements. The first GMS was launched from Cape Canaveral in July 1977 by NASA; the new generation (GMS-2) Many conditions of weather that daily affect was launched in August 1981 by Japan's National the lives of people in various parts of the Space Development Agency (NASDA), from their world exist for such relatively short periods Tanegashima Space Center, using a Japanese that polar-orbit weather satellites fail to N-II rocket. To match the payload capacity of observe them. However, the synchronous orbit the N-II launch vehicle, a spacecraft weight GMS, from its stationary vantage point reduction of 36 kg was required. This was approximately 35,800 km above the equator, has accomplished by using lighter materials and the operational capability to keep a third of fewer but more efficient solar cells. the earth's surface under constant observation and to acquire almost instantaneous informa­ LANDSAT and the Multispectral Scanner tion on rapidly changing weather patterns. from space for earth resource An instrument developed by Hughes' Santa management constitutes one of the most practi­ Barbara Research Center for making pictures cal applications of . With of the earth's cloud cover in daylight and the Earth Resources Survey Program inaugurated darkness is carried aboard the spin stabilized in 1972 by the highly successful LANDSAT GMS-1. This instrument, called a visible and (originally named the Earth Resources Tech­ infrared spin scan radiometer (VISSR), senses nology Satellite, or ERTS), the user community radiation from the earth and its atmosphere. is provided data that are unique and useful The optical consists of a scan ­ in many resource management fields. ror and and secondary mirrors. It is mounted on the spinning portion of the space­ LANDSAT circles earth from a sun synchronous craft and scans the earth using a combination 919 km circular orbit every 103 minutes or of the natural spin motion of the spacecraft 14 times a day. The is from north to (100 rpm) for east-west scan, and a mirror south at an angle of 99° retrograde to the which steps 0.2 mr every revolution for north- equator. On each north to south pass the south scan. The telescope uses folded optics satellite crosses the equator at 9:30 a.m. and images the scene onto eight visible and two local time. Each pass covers a region 185 km IR detectors via fiber optics. The visible wide with some overlap between passes. After detectors are photomultip!ier tubes and provide 18 days, or about 252 passes, the satellite 1.25 km resolution while the IR detectors are returns to the same overhead position. The -cadmium telluride and are sized to sun synchronous orbit ensures that lighting provide 5 km resolution. The camera has the angles are little changed for contiguous areas capability to send back black and white, imaged and for subsequent images of the same television-like images of one-third of the area. earth every 30 minutes, day or night, enabl ing meteorologists to identify, monitor, and track Hughes designed and built the Multispectral severe windstorms, heavy rainfall, and Scanner (MSS) which performs the actual imag­ typhoons. ing on LANDSAT. MSS is designed to scan a 185 km swath on earth. An oscillating mirror Picture data gathered by the VISSR is formatted provides the cross-track scan while the satel­ in a multiplexer/modulator and transmitted to lite 's orbit motion provides the in-track scan. the ground through S band electronics and This mirror, together with a two element all antenna. The same S band system is used as a reflective optical system (telescope) and a repeater to relay ground processed imaging data bank of 24 detectors, produce colocated images for facsimile reproduction at distribution in four spectral bands; one image is in the points in the Western Pacific area. In addi­ green spectral region, one is in the red, and tion, meteorological observation data from the other two are in the near infrared. A surface collection points (ships, buoys, and single image is approximately square (about weather stations) can be relayed to the cen­ 185 km on a side) and is made up of 7-1/2 tral processing center at the NASDA Tsukuba million pixels (picture elements), each 80 Space Center in Japan. meters square. Each pixel is quantized into

4-20 one of 64 intensity levels, and the values are used to map dynamic changes multiplexed in foraging areas into a single 15 Mbps data stream and resulted in great cattle savings in on board the spacecraft and transmitted to California and New Mexico. Icelanders have receiving sites. The received data are been able to direct their sheep to clear pas­ recovered and demultiplexed into the origi­ tures and to avoid ice damaged ranges. MSS nal channels and formatted for binary ocean ice information also permits the Ice­ recording on multitrack tapes. landers to ship their sheep at the premium spring age. Thus by looking down from its near , the MSS aboard LANDSAT can transmit nearly The list goes on and on, with long term 9000 highly detailed pictures a week for benefits in geology, , , processing, analyzing, recording, and estab­ oceanography, agriculture, and urban develop­ lishing standards tables for use by 300 prin­ ment. Equally significant is that the data cipal investigators in 50 nations. is not only fast, with excellent ground reso­ lution; it is obtained at substantial savings From comes reports that satellite data over the more conventional methods. show marked differences in tributary positions from other recent maps. Some of The MSS has also felt the impact of expendable these discrepancies are by as much as 19 km launch vehicle performance constraints. The and 90° in direction. latest version of the MSS will fly on the new Multimission Modular Spacecraft (LANDSAT D), Across the South Atlantic Ocean, Ghana reports shown in Figure 8. Due to a heavier payload MSS information is being used experimentally complement, the circular orbit altitude is to control locusts by spotting the breeding necessarily reduced from 919 to 705 km. In grounds. Reports from the African nation of order to maintain similar area coverages and Mali indicate that MSS data are being used to data rates, significant modifications to the make maps of remote areas, for the routing of MSS had to be incorporated. new roads, for water exploration, and for many other vital guidelines. The spectacular achievements of MSS in carry­ ing out the Earth Resources Remote Sensing A half a world away, uncharted reefs have been Mission have prompted NASA and the user com­ detected off . By swinging over the munity to set even broader goals, requiring Bahama Islands, MSS information has produced increased instrument capability. To satisfy maps more accurately than any current naviga­ these new requirements, Hughes is developing tion charts. Area acreages for Texas and and building a sensor called Thematic Mapper Oklahoma have been categorized as to range which has major performance improvements over and pasture, forest, and water by.using satel­ the MSS. First, its resolution is 30 meters- lite information. In Nebraska, Illinois and 2.5 times better than the 80 meters currently New York, shallow substrate, water-bearing obtainable. Second, it has seven spectral rocks were detected. Polluted water drift was bands versus four for the MSS. In addition, charted off the coast of New Jersey. the spectral separation of these bands is especially tuned to established user needs. MSS data has enabled geologists to identify Its radiometric precision is 256 levels — alternative sources for oil, gas, and other four times better than the 64 levels of the minerals in Alaska. In land use mapping,- MSS. The Thematic Mapper will fly with MSS-D maps were obsolete before they were com­ on the LANDSAT-D spacecraft. Together, they pleted, often taking 1 to 2 years. An equi­ promise to provide still new and ever chang­ valent or better MSS picture can be rendered ing information on the quality of life on this in 30 to 40 hours. . Three weeks of poor grazing can cost 113 kg of beef per animal, therefore MSS pictures were

4-21 MISSION PAYLOAD

SPACECRAFT BUS ORBIT

IMAGING/FIELDS AND PARTICLES ACTIVE \ RADIO EARTH RELAY RADIO LINK FOR SPACECRAFT CONTROL AND DATA RETRIEVAL EARTH AND OTHER CELESTIAL BODIES

FIGURE 1. GENERIC SPACECRAFT MISSION

4-22 PLANE

ENVELOPE

PAYLOAD

SEPARATION

VEHICLE

SATELLITE

LAUNCH

(PAF)

DELTA

FITTING

ATTACH

STAGE)

THREE-STAGE

ENVELOPE

2.

(THIRD

'PAYLOAD

FIGURE

^FAIRING

MOTOR

ROCKET

SOLID REORIENT SATELLITE TO

POSITION a OR S/C

US

COAST TRANSFER ORBIT ORIENT TO APOGEE INJECTION . APOGEE ORBIT ATTITUDE / INJECTION / AV = 1830 m/sec

\ SHROUDSHR OR UPPER SPACECRAFT STAGE (S/C) US (US)

LAUNCH VEHICLE AND SPACECRAFT

FIGURE 3. SPACECRAFT PLACEMENT IN EARTH SYNCHRONOUS ORBIT 1600 00 hO O§ 1400

Sg 1200

DC O 1000 SYNCHRONOUS TRANSFER ORBIT e soo

Q < 600 O

400

SYNCHRONOUS 200 EQUATORIAL ORBIT

I I I 0 9.5 10.0 10.5 11.0 11.5 12.0 12.5 13.0 CHARACTERISTIC VELOCITY, KM/SEC I I I I I 0 10 20 30 40

LAUNCH ENERGY, C3-KM2/SEC2 FIGURE 4. EXPENDABLE LAUNCH VEHICLE PERFORMANCE

4-25 1600

1500 MATCHED APOGEE MOTOR

1400 O UNDERSIZED APOGEE MOTOR Q 1300 O > 1200 QL 5 1100 cc O o: 1000 LU ULto 900 LAUNCH oc VEHICLE K to CAPABILITY 800 O Z O CC 700 Q Z 600 PLANE CHANGE FOR to MAXIMUM IN-ORBIT PAYLOAD

I I I I I t 46 8 10 12 14 16 18 20 22 PERIGEE PLANE CHANGE BY LAUNCH VEHICLE, DEGREES J______I I 29 25 23 21 19 17 15 13 11

TRANSFER ORBIT INCLINATION, DEGREES

FIGURE 5, LAUNCH VEHICLE/ARM OPTIMIZATION

4-26 TELEMETRY g AND COMMAND § BICONE £ ANTENNA

DUAL SHARED APERTURE GRID REFLECTORS

FIXED FORWARD ANTENNA SOLAR FEEDS PANEL DESPUN REPEATER THERMAL SHELF RADIATOR TWTA

REACTION BATTERY CONTROL PACK (8) (4)

PROPELLANT TANK (4)

EXTENDIBLE APOGEE KICK AFT SOLAR MOTOR PANEL

FIGURE 6. HS 376 COMMUNICATIONS SATELLITE

4-27 DESPUN ANTENNA ASSEMBLY

VISSR

VHF DBA SUPPORT ANTENNA ARRAY DESPIN BEARING ASSEMBLY (DBA)/ FORWARD ROTARY JOINT THERMAL BARRIER

VISSR COVER ELECTRONIC AND SUNSHADE EQUIPMENT SHELF SOLAR PANEL THRUST TUBE AND STRUT ASSEMBLY REACTION CONTROL SUBSYSTEM

APOGEE KICK AFT THERMAL MOTOR BARRIER

FIGURE 7. GEOSTATIONARY METEOROLOGICAL SATELLITE

4-28 HIGH GAIN ANTENNA GLOBAL POSITIONING SYSTEM ANTENNA \

ANTENNA MAST

WIDE BAND MODULE SUBSYSTEM MODULE

PROPULSION SUN SENSORS MODULE

POWER MODULE MULTISPECTRAL SCANNER SIGNAL CONDITIONING AND CONTROL UNIT SOLAR ARRAY PANEL THEMATIC MAPPER SB AND X BAND ANTENNA ANTENNA FIGURES. LANDSAT-D SPACECRAFT

4-29