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NASA

Technical

Paper

3083

1991

National Aeronautics and

Space Administration

Office of Management

Scientific and Technical

Information Program

Launch Vehicle Integration

Options for a Large Earth

Sciences Geostationary

Platform Concept

James L. Garrison

and Lawrence F. Rowell

Langley Research Center

Hampton, Virginia

Abstract

Concepts are derived for the packaging of the

components of a large Earth sciences geostationary

platform to be launched in either the Space Shuttle

or the Titan IV Complementary Expendable Launch

Vehicle. Geometric data from a proposed conceptual design for the spacecraft, antenna sizing results

from thermal and structural finite element analyses, and independent mass and volume estimates are

used to determine sizes, shapes, and masses of the

major platform components and support equipment.

Solid modelling software is used to evaluate proposed

launch vehicle integration concepts in terms of meeting volume and mass constraints, checking for interference between components, verifying that centerof-gravity locations are within vehicle specifications,

and designing suitable interface structures that do

not violate any of these constraints. Construction at

Space Station Freedom is assumed, and a space-based

orbital transfer vehicle is determined to be necessary

for inserting the spacecraft, once assembled, into geostationary orbit.

Introduction

Global changes in our environment are a documented reality, and new efforts are being undertaken to monitor the important variables reflecting

changes in our oceans, atmosphere, land, and living

species. Some variables will need to be monitored on

a nearly continuous basis so that regional events can

be observed at a frequency sufficient for understanding, modelling and, eventually, forecasting. For this

reason, Earth science sensors will have to be flown

on geostationary platforms in order to meet these

temporal resolution requirements. The size and performance requirements for these platforms are unprecedented for geosynchronous orbit, and thus the

development of new enabling technologies for science

sensors, spacecraft systems, and data and information systems may be required.

Researchers at Langley Research Center have

conducted a systems study of a large "secondgeneration" Earth sciences geostationary platform

that would accommodate the large optical instruments and microwave antennas needed to meet the

scientific requirements beyond the year 2000. The

straw-man platform design that was used in the overall study is based on the conceptual design (fig. 1)

developed by Ford Aerospace Corp. (study entitled

"Geostationary Platform Bus Study--For Earth Observation Sciences, Volume II--Comprehensive Report," WDL-TR 11066, Dec. 1987). This spacecraft concept is recognized as being too large to be

launched fully assembled. The extremely stringent

surface tolerance requirement of the solid microwave

radiometer reflector dictates on-orbit construction.

Thus, this configuration requires on-orbit assembly

and an orbital transfer vehicle (OTV) to place the

erected spacecraft into geostationary orbit.

The objectives of the overall study are (1) to

quantify performance issues and feasibility of flying

very large platforms at geostationary orbit, (2) to

identify needed technology developments to enable

or enhance the mission, and (3) to identify infrastructure and support equipment requirements. This

report, one of several produced during the study, addresses the launch packaging and vehicle integration

issues. It draws upon other reports for the configuration details (ref. 1) and antenna structural design

(refs. 2 and 3).

With the baseline platform concept as outlined in

the Ford study and the assumption of an erectable

solid surface 7.5-m antenna and a deployable 15-In

membrane antenna for the passive microwave radiometers, this study was conducted to determine

possible methods of integrating the spacecraft components into the launch vehicle for transport to low

Earth orbit (LEO). Components consistent with the

specifications of the baseline design are assumed, except for specific changes arising from subsystem design of individual components of the platform.

Only two existing U.S. launch vehicles have the

weight and volume capacity to launch all the platform components into LEO. These are the Space

Shuttle and the Titan IV Complementary Expendable Launch Vehicle (CELV). This report limits consideration to these two launch vehicles with their

present operational capacities and performance.

Only the state-of-the-art Titan IV is considered,

and far-term Space Shuttle enhancements such as the

aft cargo carrier (ACC) are not employed. However,

it is assumed that a space-based orbital transfer vehicle (SBOTV) will be available. This decision results

because all existing or near-term upper stages have

insufficient weight capacity and structural requirements of the assembled spacecraft make a low-thrust

orbital transfer necessary. Additionally, several scenarios proposed assume that an orbital maneuvering

vehicle (OMV) will be available based at Space Station Freedom.

To verify launch feasibility, proposed packaging

concepts are checked for violation of launch vehicle payload envelopes, interference between packaged components, compatibility of payload with vehicle support hardware, and acceptability of payload

center-of-gravity (e.g.) location. Required interface

and support hardware is conceptually designed and
includedin the payload,alongwith the spacecraft

components,forevaluationof all theaforementioned

criteria.

Component Mass and Volume

Estimation

A complete listing of the spacecraft components

with their mass and volume estimates is given in table 1. The net dry platform mass is 6681 kg, and

the net packaged volume of the spacecraft is 89 m 3.

These components are modelled with the GEOMOD

solid modelling program (ref. 4). Unless further specified, mass estimates are obtained directly from the

Ford study and volume estimates are derived from

GEOMOD solid models created from the geometric

data in that study. A redesign of the two antennas in

references 2 and 3 produced revised estimates of the

masses and volumes for all components associated

with the passive microwave radiometers. The purpose of this section is to summarize additional work

that was performed to estimate masses and volumes

that could not be explicitly obtained from the Ford

study or the other geostationary platform design reports (refs. 1 to 3).

Instruments and Support Subsystems

According to the baseline platform design (fig. 1),

instruments with fine pointing requirements are

placed within an instrument module located on the

spacecraft truss structure near the 15-m antenna.

Underneath the instrument module is the housekeeping module, where most of the spacecraft support

systems are located. Instruments and support components not included within these two modules are

distributed along the platform truss structure. Estimates are made in the Ford study of the masses of the

instrument module, housekeeping module, distributed instruments, and spacecraft support systems.

The Ford study gives the instrument module a

mass estimate of 3008 kg and the housekeeping module an estimate of 500 kg. For the purposes of this

study, these two large mass contributions have to be

further divided so that many of the individual components or subsystems can be treated as discrete objects. A mass estimate of 20 kg for the high-gain antenna is taken from an instrument mass breakdown

for the baseline platform and is subtracted from the

total mass of the instrument module and charged to

the command and data handling mass budget. An

estimate of 45 kg is obtained for the lower central

cylinder by assuming that areal density is constant

on all panels of the structural model of the modules

shown in figure 2. This mass is subtracted from that

of the housekeeping module. Finally, the mass of the

two solar array booms, estimated to be 13 kg, is included with that of the housekeeping module.

Both the housekeeping and instrument modules

are combined into a single entity with a net mass of

3456 kg and a center of gravity located 4.09 m above

the interface between the housekeeping module and

the lower central cylinder. A division of platform

mass based on subsystem function is also included in

the Ford study.

The net mass of the housekeeping and instrument

modules is broken down according to subsystem category as follows:

Science instruments (payloads), kg . . 1043

Command and data handling, kg .... 87

Spacecraft controls, kg ........ 81

Attitude and orbit control, kg ...... 125

Propulsion, kg ............ 173

Thermal control, kg ......... 272

Electrical power, kg ......... 397

Structure (both modules), kg ..... 639

Integration, kg ............ 339

Housekeeping and instrument modules,

total, kg .............. 3456

Station-keeping propellant, estimated at 1435 kg,

will be loaded into the housekeeping module following assembly at Space Station Freedom. Therefore,

this mass is not included as part of the housekeeping module for launch packaging but is added to the

payload manifest as part of the OTV propellant.

Other component masses not obtained directly

from the Ford study include the 10 kg of the solar

sail, 4 kg for the sail itself and 6 kg for its boom.

The package for the solar sail is a cylinder (}.53 m

in diameter and 1.50 m in height. North and south

solar array panels are given mass estimates of 76 and

73 kg, respectively. These masses include the active

cavity radiometer (28 kg) on the south panel and the

X-ray imager (31 kg) on the north panel.

Passive Microwave Radiometers

Two large passive microwave radiometers, 7.5

and 15.0 m in diameter, are mounted on the platform truss. These provide measurements of atmospheric temperature and moisture profiles at frequencies of 60, 90, 118, 160, and 183 (3Hz by the 7.5-m

radiometer and at 6, 10, 18, 22, and 37 GHz by the

15-m radiometer according to specifications in the

Geostationary Platform Study performed by Lockheed Missiles & Space Co., Inc. (Final Report,

Contract NAS8-36103, Modification No. 13. DR-9,
Nov.1988).Bothmicrowaveradiometersareof the

offset-fedCassegraindesign,with the7.5-mantenna

havingan erectablestructurewith a solid surface

madeup of hexagonalelementsand the 15-mantennahavinga deployablestructurewith a membranesurface.

Forthepurposeof makingmassandvolumeestimates,theantennasarebrokendownintofiveparts:

(1) strongback truss, (2) reflecting surface, (3) feed

array, (4) subreflector, and (5) mast boom to support the subreflector. Development and evaluation

of these designs are covered in more detail in reference 3 for the 7.5-m antenna and in reference 4 for

the 15-m antenna.

The design of the 7.5-m passive microwave radiometer is based on the precision segmented reflector (PSR) concept of reference 5. The 15-m radiometer design is based on the geo-truss concept

(ref. 6).

Strongback truss structure. Estimations of

the shape, volume, and mass of support structures

are obtained from structural models of the antennas generated by the Large Advanced Space Systems

(LASS) antenna sizing software (ref. 7). These models are shown in figure 3. The packaged size of the

7.5-m-antenna strongback is estimated to be 0.82 m

in diameter by 2.38 m long. Its mass is determined

from LASS to be 131 kg. For the 15-m antenna, this

package has a diameter of 1.97 m, a length of 1.30 m,

and a mass of 250 kg.

Within LASS, the package sizing is based on

either of two schemes for folding truss elements. One

option involves folding them inward, while the other

folds them outward. For the 7.5-m erectable antenna,

either option is acceptable. An inward folding scheme

is chosen because it has the shortest length and is

the most compact. Because the larger antenna is

deployed with its thin membrane surface attached to

points on the truss backing, the packaging scheme

that folds the elements inward must be used so that

they do not puncture the surface upon deployment.

Reflecting surface. The surface of the 7.5-m

antenna is divided into 18 hexagonal segments (2.0 m

between edges) in accordance with specifications

from the antenna design work of reference 2. These

segments are small enough to be packaged side by

side in sets of two within the Space Shuttle or Titan IV payload envelopes. Mass of the solid surface is computed from an areal density of 10 kg/m 2.

This value is an antenna sizing approximation that

includes the additional mass of hardware fittings and

control system components (ref. 5) and produces a

surface mass estimate of 441 kg. The mass of the

surface membrane, assumed to be constructed of aluminized type-H polyimide film, is estimated to be

4 kg.

The volume of the membrane is found by scaling

the packaged surface of the 5-m, three-bay geo-truss

proof-of-concept model (ref. 8) by antenna surface

area. In this model, the surface mesh fits within a

cylinder 0.67 m in diameter and 0.70 m in height (i.e.,

0.25 m3). A package volume of 2 m 3 is obtained for

the antenna by scaling up to 15 m.

Because the membrane must remain attached to

the support truss when in the undeployed state, the

package shape is a cylinder with the same diameter as

that of the folded strongback (1.97 m) and attached

to tile front of the strongback. This requires a

membrane package height of 0.72 m.

Feed arrays. Feed arrays for microwave radiometers as large and complex as those considered

in this study contribute a significant amount of mass

and volume. While the need to include contributions

of the feed arrays in the mass and volume estimates is

recognized, little information on their design for the

geostationary platform exists. In order to roughly estimate their masses and volumes, reference feed array

designs are scaled by wavelength for each of the 10

frequencies for which microwave radiometry data are

sought. The reasoning behind this sizing technique

is that the waveguides and other microwave components must be shaped based on the wavelength of

the radiation that they are carrying, and these parts

contribute the most to the mass and volume of the

feed assemblies. Effects of electronic components in

the feed arrays that cannot be scaled directly with

wavelength are approximated by applying a rough

multiplicative factor to these mass estimates. A factor of 1.5 (ref. 3) is assumed for the feed array of the

15-m antenna and a factor of 2.0 (ref. 2) is assumed

for that of the 7.5-m antenna.

The reference feed array design for mass estimates

consists of 33 horns operating at 1.5 GHz with an

approximate total mass of 500 kg. This feed array

is described in reference 9. This mass is scaled

linearly from the reference wavelength of 0.20 m to

that corresponding to the operating frequency of each

antenna. The net masses of the two feed arrays

are found by summing these individual feed masses.

These calculations produce an estimate of 443 kg for

the feed array of the 15-m antenna and 72 kg for the

feed array of the 7.5-m antenna.

Volume and shape of the feed arrays are estimated

by a similar method. Reference 9 does not give

3
sufficientdata on the geometryof the feedarray,

sodimensionsarescaledfrom measurementstaken

on a singlehorn from the feedarray of a 15-m

hoop-columnantennatestarticle.Thishornhasan

overalllengthof 0.46m andoperatesat a frequency

of 11.6GHz. It is a circularhorn with a frontal

areaof 0.02m2. To maintaincompatibilitywith

the baselinefeedarraydesignand with the mass

estimateabove,33hornsareassumedto be used

for eachoperatingfrequency.The actualnumber

of hornsrequiredis basedonthe numberof beams

neededandwhetheror notelectronicsteeringof the

radiometerisimplemented.Thesefactorscannotbe

assessedat thepresentime,sothe33-hornfeedarray

isconsideredto bearepresentativedesignandisused

astheworkingbaseline.

With the assumptionthat the additionalelectroniccomponentsthat cannotbescaledsimplywith

wavelengtharelocatedbehindthe feedhorns,the

multiplicationfactorsof 1.5and2.0areincludedin

thelengthestimatesofthefeedarraysforthe15-and

7.5-mantennas,respectively.Theselengthestimates

comefromthelengthof thelongestfeedhornin each

arrayasscaledby wavelength(0.09and0.89m for

the7.5-and15-mantennas,respectively).Afactorof

1.103isincludedinthefrontalareaestimatebecause

circularhornscannotfit togetherperfectlywithout

someamountof wastedspacebetweenthem. The

geometryof this arrangementis shownin figure4.

Scalinglengthandcross-sectional areabywavelength

andapplyingthe aforementionedmultiplicativefactors resultsin the feedarrayof the 7.5-mantenna

beingsizedat 0.64by 0.64by 0.18m andthe feed

arrayofthe15-mantennabeingsizedat 2.69by 1.26

by 1.33m.

Subre_ector. The masses of the two subrefiector

assemblies are taken from estimates used in structural models of the radiometer assemblies in references 2 and 3. Mass estimates for the subreflectors

of the 7.5- and 15-m antennas are 13 and 52 kg, respectively. GEOMOD models of these components,

based on drawings and data in the Ford study, are

modified to represent stowed configurations.

Mast booms. A point of departure from the

baseline design of the Ford study is the use of deployable boom structures based on the minimast test

article (ref. 10) to extend the subreflectors out to the

proper distance for high-resolution scanning. Masses

and volumes of these booms are found by scaling

those of the minimast test article by boom length

(10- and 20-m booms are used for the 7.5- and 15-

m antennas, respectively). Stowed boom lengths of

0.35 and 0.70 m and corresponding masses of 48 and

95 kg are predicted for the 7.5- and 15-m antennas,

respectively. The deployment mechanism required

for these booms is a large structure, a significant factor in the mass and volume totals of the packaged

system. This mechanism and other support hardware are described subsequently.

Truss Structure Package

The main section of the geostationary platform

is constructed of seven 3-m cubic bays with beam

elements similar to those proposed for Space Station

Freedom. Support for the antennas is estimated

from drawings of the platform to be the equivalent

of approximately two more bays. This estimate is

based on the length by which the main truss structure

would have to be extended to attach directly to the

antennas. At the present time the exact structural

design of the interface between the main structure

and the antennas is not defined.

Each of the nine bays (seven on the spacecraft

bus and two assumed for attachment of the antennas) contains four longerons, four battens, and four

diagonals, with an additional four battens included

in the last bay. Longerons and battens are 3.0 m

in length and diagonals are determined from basic

geometry to be 4.24 m in length. In total, the platform requires seventy-six 3.0-m elements and thirtysix 4.24-m elements.

The 3.0-m elements are small enough to be

stacked in 3.0-m-wide rectangular containers that can

be placed transversely in either the Space Shuttle or

the Titan IV. The 4.24-m elements, almost a_ long as

the diameter of the Space Shuttle or Titan [V payload envelopes, are packaged in another container,

which is oriented longitudinally.

Estimation of the volume required for this packaging is based on an assumption that these elements,

which are 0.05 m in diameter, would be spaced at

least one diameter away from each other. A multiplicative factor of 2.0 is included to allow for the contribution of other components such as wiring harnesses,

joints, assembly hardware, and a packaging structure

to hold each of the 112 truss elements in place during

launch and assembly.

Packaging volumes required for the structure and

related hardware are accommodated by two containers 3.00 by 1.50 by 0.60 m for the 3.0-m elements and

one container 4.24 by 0.60 by 0.25 m for the 4.24-m

elements. The 3.0-m elements are packaged into

two containers to allow placement on both sides of

the housekeeping module base and to facilitate easy
accessfor assemblyaswellasto distributethecomponentmassaboutthepayloadcenterline.

Thebreakdownof structuralmass(377kg from

the Ford study) betweenthe threetruss element

packagesi basedonvolumesoasto giveaconstant

density to each package. A multiplying factor of 1.5

is used to roughly account for the launch integration structure and miscellaneous assembly hardware

packaged with the structural components. A smaller

multiplicative factor is chosen for the mass estimates

than that for the volume estimates because of the

wasted space between packaged elements implicit in

the assumption that these elements are placed one

diameter apart. This empty space contributes to the

volume of the package but not to its mass.

The net mass of each of the two 3.0-m packages

is estimated to be 253 ks, and mass of the 4.24-m

element package is estimated to be 60 ks. The

50-percent mass contribution for launch integration

structure is included as part of the launch support

hardware and not as part of the payload. This

contributes a total of 377 kg to payload mass and

189 kg to launch support. A volume of 4 m 3 is

contributed to payload, and 2 m 3 is contributed to

launch support.

Vehicle Considerations

Two launch vehicles are considered in this study,

the Space Shuttle and the Titan IV expendable

booster. These two are chosen because they are the

only near-term vehicles that possess sufficient mass

and volume capacity to launch the proposed geostationary platform. In some scenarios, an OMV

based at Space Station Freedom is assumed for

transfer of payloads to Freedom from lower orbits.

All options considered in this study depend on a

reusable SBOTV to transfer the assembled platform

to geosynchronous orbit.

Space Shuttle

Payload mass properties. Present Space Shuttle payload weight capacity is 24 297 kg (53 500 lb)

for launch into a low Earth orbit of 28.5 ° inclination and 204-km (110 n.mi.) altitude. For launch

to the 407-km (220 n.mi.) altitude of the proposed

Space Station Freedom the capacity is predicted to

be 17690 kg (39 000 lb). The advanced solid rocket

motor (ASRM), a future enhancement to the Space

Shuttle, will increase these respective payload capacities to 30 164 kg (66 500 lb) and 23 133 kg (51 000 lb)

by the middle 1990's (ref. 11).

The center of gravity (c.g.) of the entire payload,

including any shuttle integration hardware, must

remain within the acceptable envelopes specified in

the Space Shuttle manuals (ref. 12). Modifications

to the c.g. specifications made as of September 1989

(private communication from Steve Gaylor, Rockwell

Space Operations Company, Johnson Space Center,

Houston, Texas) were taken into account in this

study.

Payload volume constraints. Any payload

within tile Space Shuttle cargo bay, must be confined

within the payload envelope. This envelope, shown

in figure 5, is a cylinder 4.57 m (180 in.) in diameter

by 18.29 m (720 in.) in length with an enclosed

volume of 300 m a. An interference check within the

solid modelling software is used to verify that none of

the components of the packaged system (except the

attachment pins of the support cradles) penetrate

the orbiter payload envelope, and it also verifies

that none of the packaged components unexpectedly

touch each other or occupy the same space.

To fully verify that the payload dynamic envelope

is not violated during any stage of flight, deflection

of the payload structure under the flight loading conditions and vibration environment must be checked.

Such an analysis is beyond the scope of this study,

so the assumption is made that the Space Shuttle

integrated payload is stiff enough and that a large

enough margin is left between the payload and the

envelope boundaries that this kind of interference will

not occur.

Payload integration. Each payload pallet

mounted in the Space Shuttle cargo bay is supported

in a statically determinate manner by three pinned

longeron fittings, one stabilizing longeron fitting, and

one keel pin. Pin attachment points are customarily

measured in inches in the standard orbiter-referenced

coordinate system from reference 12. The origin and

axis orientation of this coordinate system are shown

in figure 5.

Location of these attachment points is limited to

those listed in reference 12. There are two kinds

of attachment points: active, for payloads that are

removable, and passive, for payloads that are not.

In all scenarios considered in this study, the pallets

have to be removable; therefore, they all require

attachment at active pin locations.

Cradles were designed to attach several individual

payload elements onto pallets and to provide for

attachment of the pallet longeron and keel pins. The

volume contribution of the cradles is determined from

solid models of them, and mass is computed from

an estimated average density. This density is found

5
from the assumptionthat thesesupportstructures

areapproximately20percentsolidandconstructed

mostlyof aluminum. Aluminumhasa densityof

2800kg/m3,so560kg/m3isusedasanapproximate

supporthardwareaveragedensity.

For moredetailedsizing,a structuralanalysisof

the responseof thesecradlesto typicallaunchaccelerationsis required.Suchan analysisis beyond

thescopeof this study. Onlygeometricalconsiderationsareusedin the arrangementof components

intothe threepalletsandthedesignof launchsupport cradles.Theseconsiderationsincludeproviding hardpointsfor supportof themajorspacecraft

components,attachingto allowablekeeland trunnionpin locations,preventingcomponentsfrom interferingwitheachother,andnotviolatingtheSpace

Shuttlecargo bay payload envelope. Therefore, the

design of the support cradles represents a very preliminary concept for these structures. It is developed

primarily to estimate the contribution to the mass

and volume of the launch support hardware.

Remote manipulator system (RMS). Use of

the remote manipulator system (RMS) to remove

the pallets from the Space Shuttle puts important

constraints on their placement. Reference 12 requires

a 0.61-m (24-in.) clearance between a payload item

being removed with the RMS and any other cargo

or support hardware in the shuttle bay and a 1.22-m

(48-in.) clearance between the forward-most pallet

and the forward bulkhead of the cargo bay. The

latter specification can bc waived, however, if the

pallets are removable, as they are in all concepts

considered in this study.

Another consideration in using the RMS is the

placement of grapple fixtures on each of the three

payload modules. Figure 6 shows envelopes containing areas that are within reach of the RMS arm as

specified in reference 13. Positions in these plots are

measured with respect to the base of the RMS arm,

located at coordinates 17.26 m, -2.74 m, 11.30 m

(679.5 in., -108.0 in., 444.8 in.) in the standard

orbiter-referenced coordinate system of reference 12.

Reference 12 also specifies that the total distance

from the grapple fixture base to the center of gravity

of a payload attached to its end effector must be

less than 3.96 m (13.0 ft) and that the perpendicular

distance from the grapple fixture centerline must pass

within 3.35 m (11.0 ft) of the pallet center of gravity.

The range of movements allowable at the three

joints of the RMS places further limitation on its use.

These limits are as follows:

Shoulder yaw:

Shoulder pitch:

Elbow pitch:

Wrist pitch:

Wrist yaw:

Wrist roll:

180 ° to 180 °

145 ° to 2°

2 ° to 160 °

120 ° to 120 °

120 ° to 120 °

-447 ° to 447 °

Another constraint on placement of RMS grapple

fixtures is that no surrounding structure can violate

an envelope that surrounds the fixture. The dimensions of this envelope are defined in reference 12. A

contribution of 13 kg to the total package mass must

be included for each grapple fixture.

Launch loads. In this study, no calculations

involving the inertial loads imposed on the payload

components because of the Space Shuttle ascent accelerations were made. Further refinement of the

concepts proposed in this study will require this information to evaluate reaction forces at the keel and

trunnion pin attachments and to size the support

structure cradles in each of the pallets. For this reason, quasi-static launch loads for three flight regimes

of the Space Shuttle ascent trajectory are reproduced

below from reference 12. The load factors are given

in g units in the same orbiter-referenced coordinate

system used for specifying locations within the cargo

bay.

Flight event

High-dynamic-pressure

boost

[ntegrated vehicle boost

(maximum X-direction load)

Drbiter boost (maximum

I X-direction load)

Load factor, g units, in

X- Y- Zdirection direction ,direction

--1.9 :t:OAO 0.25

- .50

2.9 +0.06 4).15

2.6 +.02 .20

3.17 0 -0.60

3.05 0 .80

Transient launch accelerations are specified in reference 12 to be -3.2, 1.4, and 2.5g in the X-, Y-, and

Z-directions for the Space Shuttle ascent trajectory.

Titan IV

Payload mass properties. The payload capacity for the Titan IV in the no-upper-stage (NUS)

configuration is 11793 kg (26000 lb) to 444 km

(240 n.mi.) LEO. Unlike the Space Shuttle, there

are no detailed published criteria for c.g. position for

the Titan IV in the NUS configuration. As long as
thepayloadc.g.isnotoffsetmuchfromthecenterline

of thevehicle,it will notbeconsideredasa factorin

thispackagingstudy.

Payload volume constraints. Platform components are packaged into the 23.16-m (76-ft) payload shroud of a Titan IV NUS configuration. This

offers a payload envelope with a volume of 296 m 3

(10453 ft3).

Interference between the platform components is

checked using solid modelling software in a method

similar to that used for the shuttle model. The

Titan IV shroud payload envelope from the Titan IV User's Handbook (Martin Marietta Denver Aerospace, MCR-88-2541, Contract F04701-85-

C-0019, June 1987) is checked to ensure that no part

of the payload, except for the interface ring on the

bottom of the support structure, extends outside of

it.

Payload integration. The Titan IV CELV is

intended to launch either assembled or deployable

satellites and, as such, does not provide for attachment of many smaller payloads in a way that the orbiter longeron and keel pins do. Integration of the

individual geostationary platform components into

this launch vehicle requires conceptually designing

a support structure that will, during launch and

OMV transfer, retain three payload modules similar to those proposed for use on the Space Shuttle.

A simplified design of this structure is conceived for

the purpose of estimating its mass contribution.

Launch loads. Load factors used to size the

launch support structure are taken from those representative of Titan III launch performance, because

published values are not readily available for the Titan IV. The Titan III static accelerations assumed

in this analysis are based on those in reference 14.

Three sets of loads are selected that represent the

worst cases. The load factors for these cases are as

follows:

1. 4.0g longitudinal (compressive), 1.8g lateral

2. 1.0g longitudinal (tensile), 1.8g lateral

3. 4.6g longitudinal (compressive), 1.1g lateral

From these load factors and a minimum allowable

factor of safety of 2.0, the column supports are sized

based on buckling load and Von Mises stresses, and

the thrust tube is sized for buckling from thin-walledtube theory.

Orbital Maneuvering Vehicle

Using the orbital maneuvering vehicle (OMV) to

return each cargo pallet to Space Station Freedom,

from either the Space Shuttle or the Titan IV, for

assembly is based on the OMV design reference ntission (DRM) number 10. In this DRM, the OMV

is capable of retrieving a 22 680-kg (50 000-1b) payload module from a Space Shuttle in LEO, boosting it to Space Station Freedom 204 km (110 n.mi.)

above, and returning with another 22 680-kg module.

This operation uses 4371 kg (9636 lb) of expendables,

which include 3855 kg (8489 lb) of bipropellant fuel,

444 kg (979 lb) of hydrazine, and 72 kg (159 lb) of

cold gas.

Estimates of the expendables that will be used for

retrieval of pallets in the scenarios presented in this

study are based on the assumption that the values

obtained for the DRM can be scaled according to the

total mass of the OMV payload, the OMV, and its

fuel. This assumption implies a constant expenditure

of propellant per unit mass for the same altitude

and plane change. Calculated from the DRM, these

ratios are as follows: bipropellant, 0.062 kg/kg net

mass; hydrazine, 0.007 kg/kg net mass; and cold gas,

0.0001 kg/kg net mass.

These calculations assume the same change of altitude as used in DRM number 10, namely, 204 km

(110 n.mi.) from a Space Shuttle orbit of 204 km to

a Space Station Freedom orbital altitude of 407 km

(220 n.mi.). As such, they represent a conservative scenario in that the pallets have to be transferred to the highest orbit proposed for Space Station

Freedom.

Note that another DRM, number 12, may also be

applicable. This mission involves moving a 34 019-kg

(75 000-1b) payload and OTV combination 1 km from

Space Station Freedom in the same orbit and returning only the OMV. It is derived for the movement of

an assembled spacecraft and OTV combination away

from Freedom Station for ignition of the OTV. Because the target is closer and does not require an

altitude or plane change, only cold gas is used as

a propellant. In this study the retrieval of pallets

from a co-orbiting Space Shuttle is not considered,

and DRM number l0 is chosen as the baseline because its operation is closest to the planned retrieval

scenario.

Orbital Transfer Vehicle Considerations

A proposed beginning-of-life (BOL) platform

mass of 8116 kg (6681-kg platform and 1435-kg

station-keeping propellant) is too large for any existing upper stage system. In addition, the Ford study

7
specifiesthat thegeostationaryplatformtrussstructure is designedbasedon assumedorbital transfer

accelerationsof 0.lg.

The decisionto usean SBOTVin this study

is the only significantdeviationfrom a basicconstraintonusingonlyexistingor near-termsystems.

An SBOTVisdeemednecessarybecauseof thelimitedweightcapacitiesofall existingupperstagesand

theneedto providea low-thrusttransferto geosynchronous orbit. Also, a reusable SBOTV eliminates

the need to include this upper stage in the payload

manifest.

From figure 7 (ref. 15), the OTV delivery propellant requirement is 25 850 kg. From reference 16, the

average density of the 02 and H2 (taking into account

the proper mixture ratios) is 260 kg/m3; therefore,

the required fuel volume is 99 m 3. Station-keeping

propellant required for the platform on orbit is estimated in the Ford study to be 1435 kg of nitrogen

tetroxide and monomethyl hydrazine. Reference 16

gives the density of this propellant, averaged by mixture ratio, as 1210 kg/m 3. This requires an extra

1 m 3 of fuel volume for station-keepin_ propellant.

Therefore, a total of 27285 kg (101 m J) of propellant must be transported to LEO in addition to the

6681 kg of platform hardware.

Space Shuttle Pallet Design

In all the scenarios using the Space Shuttle, tile

platform components are packaged into three pallets.

This arrangement of components is identical for every

Space Shuttle scenario, so a brief explanation of

the packaging within the pallets is given here before

presentation of the specific scenarios.

The first pallet, designated "A," is illustrated in

figure 8(a). It contains the following components:

instrument and housekeeping modules, solar panel

arrays, truss structure, infrared interferometer spectrometer, lower central cylinder and OTV interface,

geodynamic laser ranging experiment, high-gain antenna, multichannel microwave radiometer, and the

global positioning system (GPS) antenna. The net

mass of this pallet is 5486 kg.

The second pallet, designated "B," is illustrated

in figure 8(b). It contains the 15-m passive microwave radiometer feed array, strongback, surface

membrane, and mast boom and the 7.5-m passive

microwave radiometer boom. This pallet has a net

mass of 1851 kg.

The third pallet, designated "C," is illustrated

in figure 8(c). It contains the following payload

components: subrefleetors for both antennas, solar

8

sail, and 7.5-m passive microwave radiometer truss

structure, surface segments, and feed array. Its net

mass is 978 kg.

Separation of the platform components into three

pallets is done according to the three main structural areas of the platform: the housekeeping and

instrument modules and truss structure, the 7.5-m

passive microwave radiometer, and the 15-m passive microwave radiometer. Assembly considerations

are also taken into account in the arrmlgement of

platform components into the three pallets described

above.

The housekeeping and instrument modules are

assumed to be the first components used in the

assembly sequence. They are placed on pallet A.

The lower central cylinder and OTV interface are

packaged for launch perpendicular to their on-orbit

deployed positions to prevent them from violating the

Space Shuttle payload envelope and to allow for the

proper clearances for use of the RMS arm.

Once the lower central cylinder is properly connected, the housekeeping and instrument modules

are removed from the temporary attachment location on the structure of Space Station Freedom and

taken to a dummy OTV port located elsewhere on

the station. This dummy port provides electricity,

data communications, and fueling for the reaction

control jets. Communications through this port are

dsed for checkout of the instruments and onboard

systems inside the housekeeping module. This may

be done in addition to, or as a substitute for, moving

individual instruments to a separate integration and

checkout facility on the space station.

Construction of the truss structure is assumed

to begin at the base of the housekeeping module

and move outward. Structural elements are stowed

in three packages attached to the bottom of the

housekeeping module to permit easy access to them

during the assembly sequence.

Once the platform truss structure is assembled,

the packaged 15-m antenna strongbaek and mesh are

removed from pallet B. This antenna is deployed and

attached to hard points on the spacecraft truss.

The minimast boom deployment mechanism

forms the second half of pallet B. This mechanism is

attached to points along Space Station Freedom and

then operated automatically to open the minimast

booms stowed inside of it.

Placement of the grapple fixture for pallet B on

the base of the minimast deployment mechanism can

present future problems from the standpoint of visual

cues for the RMS operator because it cannot be seen
exceptby acameraontheRMSendeffector.Alternativesincludeplacementof the grapplefixtureon

thetopof thepalletor orientationof theentirepallet 90° in pitch(privatecommunicationfromLaura

S. Mann,JohnsonSpaceCenter,Houston,Texas).

Toplacethegrapplefixtureontopof thepalletwill

requireadditionalintegrationstructurebecausethe

packagedantennacomponentsareverydelicateand

cannotserveasload-bearingpointsas the deploymeritmechanismstructuredoes.For thesereasons,

placementof the grapplefixtureon the aft sideof

thispalletisrecommended.

The surfaceand structural elementsof tile

7.5-m antenna form the basis of pallet C. The strongback truss is erected from this set first. Solid surface

elements are attached to the assembled tetrahedral

truss, and then the entire antenna is connected to

the hard points on the spacecraft structure. Launch

support structure and the used deployment mechanism will have to be returned to the ground on a

future Space Shuttle flight.

To determine the mass and volume properties of

the complete packaged system in the Space Shuttle

orbiter cargo bay, the contribution of launch support

hardware must be estimated. This hardware is separated into four areas for sizing: the support cradles

for attachment of the pallets, the RMS grapple fixtures, the mechanism required to deploy minimast

booms, and the integration structure to hold disassembled truss components.

The three sets of Space Shuttle integration hardware for pallets A, B, and C are estimated to have

masses of 162, 277, and 246 kg and volumes of 0.29,

0.48, and 0.44 m a, according to the assumptions

made previously in this section. Three RMS grapple fixtures are required, one for each pallet, with a

mass of 13 kg each.

The minimast boom deployment mechanism, ineluded in pallet B, is considered as part of the support

hardware, rather than as part of the payload, because it does not remain on the assembled platform.

Folded booms loaded into it are included as payload.

The design for this mechanism is based on that of a

test article constructed as a boilerplate model. Volume is computed to be 1 m a from this solid boilerplate model of the mechanism. Mass, however, is

taken from estimates of the flight model version to

be 721 kg from "Preliminary System Requirements

Review: Mass Properties Status Report #8, DRD

SE-28, Harris Corp., Sept. 18, 1987.

From the section entitled "Truss Structure Package," integration hardware for the truss structure,

which is included in pallet A, is estimated to contribute 189 kg and 2 m 3 to the launch support inass

and volume, respectively.

Scenarios

Four different scenarios are considered for launch

of the geostationary platform components to Space

Station Freedom. One involves a single Space Shuttle

launch, one uses the Titan IV expendable booster,

and two involve dual Space Shuttle launches. All

four assume that the assembled platform will be

transferred to geosynchronous orbit by an SBOTV.

Scenario Number 1: Shuttle Launch and

OMV Retrieval to Space Station Freedom

Description of scenario. This scenario involves launch of the platform to LEO as three specialized pallets in the Space Shuttle cargo bay. One

at a time, these pallets are removed with the Space

Shuttle RMS and retrieved with an OMV for transfer

to Space Station Freedom.

At Freedom Station, the pallets are broken down

and those instruments that require checkout before

assembly are taken to the appropriate facility. The

platform is then assembled by the crew at Freedom

using a combination of extravehicular activity (EVA)

and telerobotics. Propellant for the OMV and OTV

will be allocated to another shuttle launch.

Placement of components. The proposed

arrangement of geostationary platform components

into three pallets inside the Space Shuttle cargo bay

is shown in figure 9. Figure 10 gives the specific locations of trunnion pins that support the pallets.

Pallet A is attached to the longeron support

pins at the 25.35-m (998.20-in.) station and to the

keel and longeron pins at the 29.85-m (1175.20-in.)

station. Pallet B is attached to the longeron supports

at the 19.96-m (785.80-in.) station and to the keel

and longeron supports at the 21.46-m (844.80-in.)

station. Pallet C is attached to the longeron supports

at the 16.56-m (652.07-in.) station and to the keel

and longeron supports at the 18.06-m (711.07-in.)

station.

The interference check in GEOMOD confirms

that none of the components of the packaged system

unexpectedly touch each other or violate the Space

Shuttle cargo bay dynamic envelope.

Mass properties. Net mass of the packaged system is 8315 kg and net volume is 93 m 3. Support

mass includes the three cradle assemblies, RMS grapple fixtures, minimast boom deployment mechanism,

and truss structure integration hardware. Together
thesecontribute1634kg of massand4 rn3 of volumeto launchsupporthardware.Centerof gravity

for the SpaceShuttleintegratedpayloadis located

at coordinates24.26m,-0.003m, 10.12m (955.1in.,

-0.1 in., 398.3in.) in theorbiter-referencedcoordinatesystemdefinedin figure5. As plottedin figure11,this e.g.positionis withintheacceptableenvelopes,althoughlittle clearancefromthemaximum

forwardpositionisfound.A moredetailedsummary

of masspropertiesof thesystemandeachofthethree

palletsisgivenin table2.

Remote manipulator system considerations.

Spacing between each of the three pallets and the

cargo bay bulkheads is shown in figure 10. Clearance between the payload and the forward bulkhead

in this scenario is 0.94 m (37.2 in.). These spacings

are within the specifications given earlier.

These clearance requirements do not provide

enough space in the payload bay for a space station

docking adapter to bc included. Therefore, the orbiter is incapable of attaching directly to Space Station Freedom. The use of the OMV to retrieve each

of the payload pallets one at a time and return them

to Space Station Freedom is proposed to rectify this

problem.

Numerical data used in evaluating the constraints

on use of the RMS arm are given in table 3. Drawings

of the RMS arm configuration used in removing each

of the pallets are shown in figure 12. The RMS

grapple fixture envelopes are shown for each of the

three pallets in figure 13.

The RMS arm is modelled to reach each of the

three grapple points and is included in the GEOMOD

interference checks used to search for violation of the

payload dynamic envelope. From this information,

the arm is verified as not touching any other parts of

the cargo as it retrieves each of the pallets.

In terms of payload clearance, reach envelopes,

attached payload mass properties, allowable joint

movement, and grapple fixture clearance envelopes,

this analysis shows the feasibility of using the RMS

in the present Space Shuttle packaged configuration.

OMV retrieval considerations. A summary

of the fuel expenditure estimates for each of the six

segments of the OMV retrieval operation is given in

table 4. From these computations, net expendables

consumed are 3380 kg of bipropellant, 389 kg of

hydrazine, and 63 kg of cold gas. These represent

83, 73, and 84 percent, respectively, of the OMV

capacities.

10

Assembly considerations. As illustrated in

figure 12, the order of pallet removal is first A,

then B, and then C. Operational difficulties could

arise in the future because of this order of removal.

Pallets are arranged from heaviest to lightest in

moving from the aft to the forward position in order

to force the launch payload center of gravity aftward

and to keep it within an allowable range.

Placement of the grapple fixture on the aft side

of pallet B requires that pallet A be removed first.

This order of deployment moves the payload center

of gravity forward as each pallet is removed. If

a contingency is allowed for immediate return of

the orbiter in an emergency, then having partially

removed the cargo could cause a violation of the

center-of-gravity envelope and result in aerodynamic

return problems. These concerns were pointed out

in the previously mentioned private communication

from Laura S. Mann. Full consideration of the

operational consequences of this proposed scenario

is beyond the scope of this study.

Scenario Number 2: Titan IV Launch and

OMV Retrieval to Space Station Freedom

Description of scenario. A scenario for placing the platform in orbit using the Titan IV begins

with a launch to LEO. The platform components are

retained inside the launch support structure, which

was conceptually designed previously. Once in orbit, this package is separated from the final stage of

the Titan IV, and an OMV based at Space Station

Freedom mates with a docking port on the top of this

structure. The OMV then transfers the complete system to Freedom Station.

The payload then docks with the dummy OTV

port on Space Station Freedom using the OTV

adapter on the lower central cylinder. Then the

OMV separates from the payload forward end. The

launch support structure is disassembled, and platform components are removed by telerobotic operators at Space Station Freedom. Because the Titan IV is unmanned and is not intended to reach

Space Station Freedom altitudes, an OMV is needed

to move the entire payload, once on orbit, to Space

Station Freedom for removal of the platform component modules.

Placement of components. The arrangement of payload components on the launch support structure inside the Titan IV 23.16-m (76-

ft) payload shroud is shown in figure 14. The

locations of payload groups within the Titan IV

shroud are shown in figure 15. The thre_ modules are arranged with the housekeeping module
at the bottom, as mentionedpreviously,to allow its OTV dockingmoduleto be attachedto

the dummyport on SpaceStation Freedom. The

15-m antenna assembly, with the minimast deployment mechanism, is placed on a platform located at

the 7.43-m (-292.5-in.) station in the Titan IV coordinate system defined in the Titan IV User's Handbook. Surface segments and the strongback truss

structure for the 7.5-m antenna are placed on a platform at the 11.24-in (442.5-in.) station, and the

OMV interface is located above them at the 15A8-m

(-609.5-in.) station. Tile individual payload modules

are illustrated in figure 16.

An interference check with GEOMOD determined that none of the packaged components, or the

launch support structure, unexpectedly touch each

other or violate tile Titan IV payload envelope.

Launch loads and structural design of support equipment. Inertial loads from the payload sets at the 11.24-m (442.5-in.) and 7.43-m

(-292.5-in.) stations are transferred through four column supports to a conical thrust tube, which also

transfers the loads from the 5311-kg module. A drawing of this simplified support structure is shown in

figure 17.

This payload support structure is designed from

the static launch loads given previously in order to estimate its contribution to the packaged system mass

and volume. Aluminum is chosen as the structural

material. Adequate strength is provided by 0.10-mdiameter columns with thicknesses of 0.008 m. The

thrust tube has a top radius of 2.29 in, a bottom

radius of 1.57 m, and a height of 1.29 m to attach

properly between the forward support structure and

the Titan IV payload interface ring. The thickness

of this thrust tube is 0.006 m to provide adequate

support for the launch loads.

With a density of 2800 kg/m 3 for aluminum, this

support structure has a mass of 801 kg. A 30-percent

contingency in this estimate is added to account for

attachment hardware and non-loadbearing structural

elements to bring its contribution to support mass to

1041 kg. Volume of this structure is 3 m 3.

Mass properties. The net mass of the packaged

system is 8645 kg, and net volume is 95 m 3. This

mass breakdown is summarized in table 5. Location

of the center of gravity is 5.37 m, 0.06 m, 0.04 m

(-211.5 in., 2.5 in., 1.6 in.) in the Titan IV coordinate system. As mentioned previously, there are

no detailed published criteria for the payload c.g. location for the Titan IV. The e.g. offset only 0.08 m

(3.2 in.) from the vehicle centerline is not expected

to present a problem.

The platform components included as payload remain unchanged from the Space Shuttle launch design. Support mass includes the 1041 kg determined

above, the OMV retrieval grapple fixture (assumed

to be the same as those on the shuttle, with a mass of

13 kg), the minimast boom deployment mechanism,

and the truss structure package. Total support structure contribution to mass is 1964 kg and contribution

to volmne is 6 m 3.

OMV retrieval considerations. With the

same assumptions for use of the OMV as is done in

scenario number 1, it is estimated that the OMV uses

1883 kg of expendables in retrieving the Titan IV

payload module from a 204-kin (ll0-n.mi.) orbit.

This expenditure is broken down as follows: transfer

of empty OMV from Space Station Freedom to LEO

(bipropellant, 587 kg; hydrazine, 66 kg; and cold

gas, 9 kg) and retrieval of Titan IV payload module

(bipropellant, 1082 kg; hydrazine, 122 kg; and cold

gas, 17 kg). Net expendables required are 1669 kg

of bipropellant, 188 kg of hydrazinc, and 26 kg of

cold gas. These amount to 41, 36, and 35 percent,

respectively, of the OMV capacities.

Assembly considerations. The platform components are arranged into three sets, as with the

Space Shuttle configurations. Special considerations

for use of the Titan IV include extension of the

OTV docking adapter on the lower central cylinder

to the base of the Titan IV interface ring. This allows the complete packaged system, once retrieved

by the OMV, to be docked to the dummy OTV port

on Space Station Freedom. Once docked securely to

Freedom Station, the launch support structure is disassemt)led and platform components are removed for

checkout and assembly.

Scenario Number 3: Shuttle Launch

Direct to Space Station Freedom

Description of scenario. This scenario involves direct launch of the platform to Space Station

Freedom in two Space Shuttle flights. Three pallets,

identical to those used in scenario number 1, are divided between the two flights. Pallet A is placed on

the first launch, and pallets B and C are included on

the second launch. Both flights can accommodate a

Space Station Freedom docking adapter.

Each orbiter docks at Space Station Freedom

using the standard interface. The pallets are removed

from each flight at Freedom. After the second flight,

the platform is assembled. Assembly operations are

11
assumedto be similar to thoseusedin scenario

number1.

Placement of components. The proposed arrangement of the pallets for the two Space Shuttle

launches is shown in figure 18. Figure 19 gives the

locations of trunnion pin attachment points for the

first launch, and figure 20 gives this information for

the second launch. These positions are measured in

inches in the standard orbiter-referenced coordinate

system.

Pallet A is launched on the first flight attached

to the longeron support pins at the 22.56-m (888.07-

in.) station and to the keel and longeron pins at the

27.05-m (1065.07-in.) station. Pallets B and C are

launched on the second Space Shuttle flight. Pallet C

attaches to the longeron supports at the 19.46-m

(766.13-in.) station and to the keel and longeron supports at the 20.96-m (825.13-in.) station. Pallet B

is attached to the longeron supports at the 23.80-

m (939.20-in.) station and to the keel and longeron

supports at the 25.35-m (998.20-in.) station. The

docking adapter is located in the standard position,

with its centerline at the 15.72-m (619-in.) station.

Support hardware included within each of the

three pallets is identical to that proposed for scenario

number 1.

An interference check, similar to that used on

the single launch configuration, is performed within

GEOMOD to verify that none of the packaged components unexpectedly touch each other or occupy

the same space. The GEOMOD interference check

also verifies that the Space Shuttle payload envelope

and the grapple fixture clearance envelopes are not

violated.

Mass properties. Net mass of the packaged

system on the first launch is 6189 kg and net volume

is 82 m 3. On the second launch, net mass is 3532 kg

and net volume is 28 m 3. Launch support contributes

a total of 1067 kg and 11 m 3 on the first launch

and 1973 kg and 11 m 3 on the second launch. This

includes the docking adapter, which adds 703 kg

of mass and 9 m 3 of volume to each flight. The

e.g. for this flight is located at coordinates 23.18 m,

-0.01 m, 10.07 m (912.3 in., -0.42 in., 396.3 in.)

in the standard orbiter coordinate system. For the

second flight, e.g. position is located at 20.21 m,

0.003 m, 10.06 m (795.5 in., 0.1 in., 395.7 in.).

Centers of gravity on both flights are within the

acceptable envelopes, as shown in figure 21.

As in scenario number 1, this e.g. is very close

to the maximum forward position allowed. However,

12

as discussed in the next scenario, this arrangement

provides the option of dividing the OTV propellant

between these two flights instead of requiring a dedicated launch, as would be the case in scenario number 1. If this propellant is included, it moves the e.g.

for both flights aft a significant distance. To provide

some flexibility and a factor of safety in integrating

OTV propellant containers on these two launches, if

that is desired, a movement of the dry payload e.g.

as fax forward as possible may be advantageous. A

more detailed summary of mass properties for each

launch is given in table 6.

Space Station Freedom RMS considerations.

When the Space Shuttle is docked at Space Station Freedom, the space-station-based RMS arm is

assumed to be used to remove the pallets from the

cargo bay. If a space-station-based RMS arm is unavailable, the arm inside the Space Shuttle can also

be used. Requirements for payload clearances, mass

properties, and grapple fixture envelopes are assumed

to be the same as, or more liberal than, those specified for the orbiter-based RMS.

Spacing between each of the pallets and cargo bay

bulkheads for both launches is shown in figure 19 for

the first launch and in figure 20 for the second launch.

All these clearances are greater than the 0.61-m (24-

in.) minimum specified for the orbiter-based RMS.

Since the three pallets are identical to those used

in the previous single launch configuration, the c.g.

offsets from the grapple fixtures are equal to those

for that configuration. Refer to table 3 for the

pallet mass properties in relation to use of the RMS.

As shown previously, these offsets are within the

allowable limits, and no interference exists with the

grapple fixture envelopes.

Scenario Number 4: Two Shuttle

Launches, Including OTV Propellant

One important factor to consider in packaging

the geostationary platform components is the large

amount of propellant required for the OTV transfer.

This is computed to contribute 27285 kg of mass

and 101 m 3 of volume. In any concept proposed, this

supplemental mass will require at least one additional

launch of either the Space Shuttle or an expendable

vehicle. Scenario number 3, in which the payload is

divided between two launches, provides a sufficient

amount of unused volume in which to include this

propellant.

In this scenario, the OTV propellant is packaged

into the excess volume on each of the two launches.

The three pallets will be located in the same positions

within the cargo bay. Each Space Shuttle launch will
thuscontainsomeplatformcomponentsaswellms

someOTVpropellant.

In figure22, availableareaswherethe propellant canbelocatedareindicated.Onthefirstflight

therewill be46m3 of volumebetweenthe 29.65-m

(1167.44-in.)and33.07-m(1302-in.)stations(allowingthe0.61-m(24.00-in.)marginfortheRMSarm

to removethepropellantcontainers).Onthesecond

flight,117m3of volumewill beavailablebetweenthe

25.35-m(998.2-in.)and33.07-m(1302-in.)stations.

Toevaluatetheoptionof dividingthisadditional

payloadbetweenthetwoSpaceShuttleflightsin this

scenario,the contributionto massand volumeof

thetankageandsupportstructureis assumedto be

30percentof the propellantalone. No furtherassumptionsaremadeasto the exactnatureof this

support. Net massand volumecontributionsincurredby includingOTV andstation-keepingpropellantonthetwoSpaceShuttleflightsare35470kg

and 131m3. This estimateof propellantandsupport equipmentmasssignificantlyexceedsthe projectedcapacityof theSpaceShuttlewith theASRM

to LEO(204km) of30164kg. Therefore,theoption

ofusingasecondSpaceShuttlelaunchforthepropellantalonewill not befeasibleunderpresentshuttle

capabilities.

If 10612kg (39 m3) are includedin the first

flight,thec.g.is movedaft to the28.34-m(1115.86-

in.) stationandthe payloadmassis increasedto

16801kg. If 24859kg (92m3) areincludedin the

secondflight, the c.g.is movedaft to the 28.09-m

(1105.98-in.)stationandthe net massis increased

to 28392kg. Thesenewc.g.locationsareindicated

in figure23.

By includingthis propellantwith the payload

manifest,dividingplatformcomponentsbetweentwo

launchesbecomesmoredesirable. Note that althoughthe first flight is belowthe projectedSpace

Shuttlecapacity(17690kg)to attainthelowestorbit proposedfor SpaceStationFreedom (407 km),

the second flight will slightly exceed the 204-km altitude capability, so this option is dependent on future

Space Shuttle capability upgrades such as the ASRM.

With such a large payload, the second launch

will probably require OMV retrieval to Space Station

Freedom. Therefore, the docking adapter will not

be necessary and can be removed, slightly lowering

the predicted payload weight but not enough to

change the conclusions about the feasibility of this

scenario. Another possibility will be to remove the

OTV propellant using OMV sorties first, and then

change orbits and dock with Space Station Freedom

using the docking module to then unload the pallets.

Space Shuttle operations such as that have not been

investigated any further.

Summary of Payload Integration Concepts

(Scenarios)

A summary of the three geostationary platform

payload integration concepts considered in this study

is given in table 7. The breakdowns in net mass

and volume among platform components, launch integration hardware, and OTV propellant are shown

graphically in figure 24 for scenario number 1, in figure 25 for scenario number 2, in figure 26 for scenario

number 3, and in figure 27 for scenario number 4.

Scenario number 1, a single Space Shuttle launch

with OMV retrieval, has the advantage of packaging all the platform into one flight using only present

Space Shuttle capabilities. This advantage would be

offset by the need for a second mission to launch

OTV propellant, although separating the spacecraft

and the propellant could be advantageous. Its primary disadvantage is the reliance on many separate

aspects of future space architecture, including the

OMV, OTV, and Space Station Freedom, as well as

the need for two separate crews, one at Freedom and

one on the orbiting Space Shuttle. This complexity

is compounded by the fact that, with all three pallets

in one flight, direct docking with Space Station Freedom is impossible. Also, using a dedicated OTV and

station-keeping propellant launch after the platform

is assembled may not be practical because the payload mass of such a launch exceeds projected Space

Shuttle with ASRM capabilities for LEO.

Use of the Titan IV, as in scenario number 2,

has the advantage of not being dependent on Space

Shuttle launch schedules. It also does not require

a second crew. The disadvantage of this scenario

is the need to provide nonstandard launch support

hardware. This concern could give rise to problems

unforeseen at the present time if this design option is

pursued further. Other than the nonstandard launch

support structure, the only special hardware required

is an OMV based at Space Station Freedom.

Separating the geostationary platform package

into two Space Shuttle flights, as in scenario number 3, allows docking at Space Station Freedom.

However, this scenario presents the disadvantages of

having to store the components launched on the first

flight in either a disassembled or a partially assembled form. Also, scheduling an additional launch into

the Space Shuttle manifest could be difficult. A final

disadvantage is that there will still be the need for

another mission to get the OTV propellant up.

13
Requiringtwoflightsforthepayloadmaystill bc

attractiveasa meansof launchingthe platformif

the OTV propellantis includedin the flights. By

separatingthepayloadbetweentwoflightsit is conceivablethat the OTV propellantcouldbeaccommodatedwithintheexcessvolumeandmassonboth

flights,if the advancedsolidrocketmotor(ASRM)

is available.However,havingthe platformassembledandcheckedout beforethe OTV propellantis

launchedis anadvantageof usinga dedicatedpayloadflight andadedicatedpropellantflight. If this

propellantis included,thenthe secondlaunchwill

requireretrievalbyanOMVbecauseits payloadwill

exceedSpaceShuttleperformancelimits to Space

StationFreedom altitudes.

Recommendations and Options

Packaging and integration concepts for launch of

the components of a large Earth sciences geostationary platform in either the Space Shuttle or the

Titan IV Complementary Expendable Launch Vehicle have been derived. On a conceptual level, all

these concepts meet launch vehicle requirements and

do not exceed projected performance specifications.

The comparison of the four can be based primarily on mission complexity, amount of nonstandard

hardware and operations required, and level of space

technology infrastructure assumed.

All four scenarios considered in this study require

Space Station Freedom and an SBOTV. Dependence

on either of these pieces of future space infrastructure

can put the schedule of the geostationary platform

in question. However, the only scenarios that do

not require Space Station Freedom or an SBOTV

entail either significant modifications to the basic

geostationary platform design or utilization of farterm technology.

Future options might include the following: assembly of the platform on an orbiting Space Shuttle,

launch of the 7.5-m microwave radiometer assembled

in the ACC of the Space Shuttle, use of proposed

upgrades to the Titan IV, or application of existing upper stages. Consideration can also be given to

launching individual components packaged on Spacelab payload accommodation pallets. Studies of this

option show that sufficient payload volume is difficult

to obtain because of the space taken up by the pallets

themselves. Also, removal of the individual platform

components is difficult without a heavy reliance on

EVA.

If the dependence on existing launch vehicle

technology is relaxed, the heavy lift launch vehicle

(HLLV), the Advanced Launch System (ALS), or

14

Shuttle-C could be considered as future options. Further evaluation of these options and their impact on

utilization of launch vehicle capabilities is beyond the

scope of this study.

Recommendations to improve the basic platform

concept for the next design iteration include reducing its mass below 4535 kg (10000 lb) and strengthening the truss structure to allow use of a Centaur

upper stage for transfer to geosynchronous orbit. Reducing the amount of assembly required in LEO is

also desirable because this could eliminate the dependence on Space Station Freedom. A deployable

design below 4535 kg is optimum because this could

be launched directly to geosynchronous orbit on a

Titan IV/Centaur. Options for a partially deployable platform include launch of the main bus and

the Centaur upper stage into LEO by a Titan IV

and then assembly of the remainder of the spacecraft

by a Space Shuttle crew.

Concluding Remarks

This paper has demonstrated, on a very preliminary basis, several options for placing the components of a large Earth sciences geostationary platform at Space Station Freedom. The four scenarios

considered, three using the Space Shuttle and one using the Titan IV, were all shown to be viable mission

choices, dependent on near-term hardware only.

NASA Langley Research Center

Hampton, VA 23665-5225

May 9, 1991

References

1. Pidgeon, David Joseph: A Subsystem Design Study of

an Earth Sciences Geostationary Platform. M.S. Thesis,

George Washington Univ., July 1989.

2. Farmer, Jeffery T.; Wahls, Deborah M.; and Wright,

Robert L.: Thermal-D_stortion Analysis of an ATttenna

Strongback for Geostationary High-Frequency Micwwave

Applications. NASA TP-3016, 1990.

3. Wahls, Deborah M.; Farmer, Jeffery T.; and Sleight,

David W.: On-Orbit Structural Dynamic Performance of

a 15-Meter Microwave Radiometer Antenna. NASk TP3041, 1990.

4. I-DEAS TM Geomod@. Solid Modeling and Design,

User's Guide, Level _. Structural Dynamics Research

Corp., c.1988.

5. Collins, Timothy J.; and Fichter, W. B.: Support Trusses

for Large Precision Segmented Reflectors: Preliminary

Design and Analysis. NASA TM-101560, 1989.

6. Freeland, R. E.: Survey of Deployable Antenna Concepts. Large Space Antenna Systems Technology- 1982,

E. Burton Lightner, compiler, NASA CP-2269, Part 1,

1983, pp. 381--420.
7. Leondis, Alex: Large Advanced Space Systems ComputerAided Design and Analysis Program Programmer Document. NASA CR-159191-3, 1980.

8. lager, J. A.: Status of Deployable Geo-Truss Developinent. Large Space Antenna Systems Technology 1982,

E. Burton Lightner, compiler, NASA CP-2269, Part 1,

1982, pp. 513 525.

9. Foldes, Peter: A Design Study for the Use of a Multiple

Aperture Deployable Antenna for Soil Moisture Remote

Sensing Satellite Applications. NASA CR-178154, 1986.

10. Adams, Louis R.: Design, Development and Fabrication

of a Deployable/Retractable Truss Beam Model for Large

Space Structures Application. NASA CR-178287, 1987.

11. National Space Transportation System Program Schedule.

NSTS 08178-17, NASA Lyndon B. Johnson Space Center,

Sept. 1989.

12. Shuttle Orbiter/Cargo Standard Interfaces, Revision I.

ICD 2-19001, NSTS-07700, Volume XIV, Attachment 1

13.

14.

15.

16.

(ICD 2-19001), NASA Lyndon B. Johnson Space Center,

Sept. 16, 1986.

System Description and Design Data Payload Deployment and Retrieval System. NSTS (}770(}. VohlIne XIV,

Appendix 8, Revision J, NASA Lyndon B. Johnson Space

Center, May 2. 1988.

Dynamics and Loads Section, Martin Marietta Corp.,

Denver Div.: Program Titan IIIB Design Loads Report

(Final). SSD-CR-66-105 (Rev. 1) (Contract AF04(695)-

694), June 1967. (Available from DTIC a_s AD 816 408.)

Kofal, Allen E.: Orbital Transfer Vehicle Concept D@

inition and System Analysis Study. t_blume IX Study

Extension Results. NASA CR-183549, 1986.

Sutton, George P.: Rocket Propulsion Elements An Introduction to the Engineering of Rockets, Fifth ed. John

Wiley & Sons, Inc., e.1986.

15
Table 1. Platform Mass and Volume Estimates

Platform component Mass, kg Volume, m 3

15-m passive microwave radiometer:

Strongback truss

Membrane surface

Feed array

Mast boom

Subreflector

7.5-m passi_'e microwave radiometer:

Strongback truss

Surface panels

Feed array

Mast boom

Subrefleetor

Instrument and housekeeping modules

Infrared interferometer spectrometer

250

4

443

95

52

131

441

72

48

13

3456

860

Lower central cylinder and OTV interface

Main structure

Geodynamic laser ranging experiment

Multichannel microwave radiometer

High-gain antenna

Solar arrays (includes active cavity radiometer and X-ray imager)

GPS antenna

Solar sail

Contribution from smaller components a

45

377

100

100

20

149

15

10

54

4

2

4

1

1

0

5

0

0

1

Total platform f 6681 89

aContribution of components with volumes listed as 0.

16
Table2. MassPropertiesforScenarioNumber1

Component

PalletA

Pallet B

Pallet C

Total packaged system

Mass, kg

5486

1851

978

8315

26.92

20.14

17.15

24.26

Center of gravity, m,

in orbiter-referenced

coordinate system

y Z

-0.01 10.11

0 10.08

0.01 10.24

0 10.12

Table 3. Remote Manipulator System (RMS) Reach Analysis

for Scenario Number 1

Module c.g. in orbiter-referenced

coordinate system, in.:

X ...........

y ...........

Z ...........

Grapple fixture location in orbiterreferenced coordinate system, in.:

X ...........

y ...........

Z ...........

Grapple fixture location unit vector:

X ...........

y ...........

Z ...........

Total distance from grapple

to module c.g., in. . .

Pallet A

1055.8

--.2

398.0

1113.7

0

460.1

0

0

1.0

Pallet B Pallet C

793.0

0

396.9

851.9

0

364.2

1.0

0

0

675.3

.5

403.0

681.5

60.7

438.7

0

0.5

0.9

84.9 67.3 70.3

Perpendicular distance from grapple

to module c.g., in. 57.9 58.9 34.8

RMS arm joint orientation, deg:

Shoulder pitch .....

Shoulder yaw .....

Elbow pitch ......

Wrist pitch ......

Wrist yaw .......

43.3

14.0

61.1

72.2

0

-46.2

23.8

116.6

109.0

-23.8

-88.6

89.6

133.0

-75.5

--.3

17
._

=

x

_q

o o

0

o

_p

o

cp

0_

o

0

rae'_

18
Table 5. Mass Properties for Scenario Number 2

Center of gravity, m, in

Titan IV coordinate system

Component

Payload set at 4.14-m (163.0-in.) station

Payload set at -7.43-m (-292.5-in.) station

Payload set at 11.24-m (-442.5-in.) station

Launch support structure and grapple fixture

Total packaged system

Mass, kg

5311

1708

572

1054

8645

X

-2.95

-8.90

12.13

-8.18

5.37

-0.01

.35

.01

0

0.06

0.07

.41

.01

0

0.04

Table 6. Mass Properties for Scenario Number 3

Center of gravity, m, in

orbiter-referenced coordinate system

Component Mass, kg x y z

First flight: pallet A 5486 24.14 -0.01 10.11

First flight: Space Station Freedom

docking adapter 703 15.72 0 9.74

Total packaged system, first flight 6189 23.18 -0.01 10.07

Second flight: pallet B 978 24.50 0.01 10.24

Second flight: pallet C 1851 19.64 0 10.08

Second flight: Space Station Freedom

docking adapter 703 15.72 0 9.74

Total packaged system, second flight 3532 20.20 0 10.06

19
Table 7. Summary of Mass and Volume for Four Scenarios

Scenario

Number 1

Number 2

Number 3,

first launch

Number 3,

second launch

Number 4,

first launch

Number 4,

second launch

Launch

vehicle

weight Payload

capacity, mass,

kg kg

24267 6681

11 793 6681

17690 5122

17690 1559

17690 5122

30 164 b 1559

Support

hardware

mass_

kg

1634

1964

1067

1973

3330

7881

Total

OTV packaged

tirol system

mass_ mass_

kg kg

7542

19692

8 315

8 645

6189

3 532

15994

29132

a With 23.16-m (76-ff) shroud.

b Projected mid-1990's capacity to 204 km(ll0 n.mi.) with ASRM.

Launch

vehicle

envelope

volume_

m 3

30O

296 a

30O

30O

30O

300

Payload

volume,

m 3

89

89

71

18

71

18

Support

hardware

volume,

m 3

4

6

11

10

19

32

OTV

fuel

volume,

m 3

39.1

91.7

Total

package

system

volume,

m 3

93

95

82

28

130

142

20
go

(D

e_©

e_

.e©m

21
Instrument

module

i

Housekeeping

t module

Lower central

cylinder

Figure 2. Internal structure of instrument and housekeeping modules. All dimensions are in meters.

22
J

7.5-m antenna 15-m antenna

Figure 3. Finite element models of passive microwave radiometer structures.

23
/

/

t

Within the (_

triangle ABC t

\

/

I

/

/ \

/

/ //

I II

Area A = _ R 2

Total area of feed array

Y_,Area of individual feeds 3 =R2

6

= 1.103

Figure 4. Wasted area in fitting circular feed horns into array.

/

/

24
4.57

(180)

X

10.16

(400)

Figure 5. Space Shuttle orbiter-referenced coordinate system. Dimensions are in meters (inches).

25
y = 360 in.

y= 160 in.

y = 260 in.--_

y = 80"

y=Oin.-_

I

-700 -._ O0 50O

y = 460 in.

y=Oin.

-300

x= 120

x=O"

x = 60

Z, in.

\

x=Oin.

x = 60 in.

x = 180 in.

x = 240 in.

I

-700 -500

x = 300

x = 320

x = 420 in.

iY, in.

500 700

x = 400 in.

x=Oin.

_--x = 60 in.

-400

Figure 6. Reach envelope of shuttle RMS (measured in RMS-centered coordinate system).

26
%

0

rr"

O_

_E

23'
i Geodynamic laser ranging experiment

f Solar panel

/ Instrument module _-_

L Infrared

interferometer

spectrometer

/- Multichannel microwave

antenna / radiometer

_ / ,/- 3-m truss elements

_ package

I t _'- 4.24-m truss

_y_::::__ elements

_'--_'- package

. '- Housekeeping module

Z/ eel pin

(a) Pallet A.

_- 15-m passive microwave

/f__ radiometer membrane surface

15-m passive // _-"- _ _.

microwave _ /T_--_

radiometer feed array _C_'_m__l_ I /_\\ 1. \_ _ 15-m passive microwave radiometer

i !t i_ t I_ strongback truss

Minimastboomdeployment_ J- _---'_:_,__f_'_lt /2

mechanism _//'_ \_ "_i _///1 11/.._'- 15-m passive

\ \\_ _l l_// _'_Y_//_ microwave

/_/_ _- 7._:mpassivemiorowave

__ _dJ_/- radiometer boom

_- Kee_ pin

(b) Pallet B.

Grapple "-_ __ 7.5-m passive microwave fixture radiometer truss

\ _ _._ structure

Longeron _ _ _ 7.5-m passive microwave

pins radiometer surface

segments

Longeron

pins

15-m passive microwave

radiometer subreflector

Launch support structure

_|_ _ Keel pin

L Solar sail

(c) Pallet C.

Figure 8. Space Shuttle orbiter cargo pallets.

28
Figure 9. Space Shuttle arrangcment of pallets for scenario number 1.

29
Grapple fixtures//_

0.635 /

(2_=

0.940

_,1(37 ) 30l

0.610

(24)

L I

I

Station 165 180 199 214 253 298

18.061

(711.07) 21.458

(844.80)

25.354

16.563 (998.20)

(652.07) 19.959 29.850

(785.80) (1175.20)

(a) Top view.

(b) Side view.

Figure 10. Top and side views of Space Shuttle orbiter packaged configuration for scenario number 1. All

dimensions are in meters (inches).

3O
30rx 103

60 -<103 28J

261

241

50 - _ 221

,_ 201

40- _ 18t61

_ 14] E 30 - --

o 101

_'20 -

n 81

61

10 ," 41

oL

Capability to

204 km ("

OV 103, OV 104,

and OV

Payload c.g.

102

2 4 6 8 10 12 14 16 18

Distance from forward end of payload bay, m

100 200 300 400 500 600 700

Distance from forward end of payload bay, in.

Capability to

407km (220 nmi.)

(a) X-direction.

x 103

60-

25

50-

20

__40-,_

E

"o 15

3o-_

.9o a_ 10

>, 20-

(3-

10

x 103

Capability to V

204 km

(110 nmi.)--L-_.

i.= =

5

0

-.50 -.40 -.30 -.20 -.10

rOY 103, OV 104,

....... _ Capability to

..... _ _ 407km

",.= ""r'''_ "" (220 nmi')

Payload c.g. location

J _3 i 0 .10 .2 .30

Y distance from centerline of payload bay, m

% :s 6 _ t'o

Y distance from centedine of payload bay, in.

I I

.40 .50

_5

(b) Y-direction.

120 - 3_

Capability to 407 km

(220 n.ml.), OV 102

r to 204 km

Payload c.g. OV 102

Capability to 204 km

(110 n.mL), OV 103,104,

and 105

ICapability to 407 km

(220 n.mf.), OV 103, 104, and 105

' _, ' _ ' 1'2 ' t'6 ' _0 ' 2'. ' 2'8×IO3

Payload mass, kg

i'o _o _o _o ' _o 0'0×1o3

Payload mass, Ib

(c) Z-direction.

Figure 11. Space Shuttle payload c.g. locations for scenario number 1.

31
Elbow - Wrist --_

Shoulder-_ _ [- End effecter__

(a) Pallet A first.

Elb°w -_ _ Wris___t 1

%

_nd effector

(b) Pallet B second.

Elbow-_

Shoulder-_ _

/--Wrist

/- End effector

,%

(c) Pallet C third.

Figure 12. Removal of pallets with RMS for scenario number 1.

32
(D

"_ ID.

O5 0

o_

33
23.16-m (76-ft)

payload shroud

OMV grapple

fixture

Launch support

st ru ct u re

Figure 14. Titan IV launch packaging for scenario number 2.

34
MV grapple fixture _O - a-load envelope 'q

\ _3"itan W NOS P Y

pport structure

orioino! _63. 1

1 " 43-1 cOJ°_il gradn_t

.24 system

._5.48 (,.442.5) (-292.5) \

(,-609.5)

,f.direction

Figure 15. T_t_n IV N_JS packaged c°nfigur_t_°n for scenario number 2. A_ d_menS_OnS are in meters

(inches).
7.5-m passive microwave radiometer

strongback truss

(a) 572-kg module.

7.5-m passive microwave

radiometer surface

segments

15-m passive microwave radiometer

strongback truss

/-- 15-m passive microwave

__ _ _ \ \\ /radiometer feed array

15-mpassivemicrowave __1'

/

-_ _--=\ i_ _ 15-m passive microwave 7.5-m passive microwave _--=_ radiometer subreflector

radiometer boom

(b) 1708-kg module.

4,24-m truss

elements--

/-

Housekeeping module -_ j

Instrument module

ii panel Launch integration structure

Geody I a Infrared interferometer

ranging experiment spectrometer

(c) 5311-kg module.

Figure 16. Titan IV NUS payload modules for scenario number 2.

3-m truss elements

Lower central

cylinder and OTV

interface

36
T--

CO

¢5

_V

t'M

O

04

/ I

/

.1F

V

4.58

572 kg

1708 kg

-- 5311 kg

Thrust tube

Figure 17. Simplified Titan IV launch support structure model. All dimensions are in meters unless

otherwise indicated.

37
Figure 18. Space Shuttle packaging for scenario number 3.

38
29.653

(1167.44)

\

Grapple fixture 33.071

(1302.00)

I

Station 225 270

15.723 22.560

(619) (888.07)

27.050

(1065.07)

(b) Side view.

Figure 19. Top and side views of Space Shuttle packaged configurations (first flight) for scenario number 3.

All dimensions are in meters (inches).

39
Station

Grapple

25.354 33.071

(998.2) (1302)

i

194209 238 253

i

15.722 23.800

(619) (939.20)

2O.958

(825.13)

19.460

(766.13)

25.354

(998.20)

(a) Top view.

_ i_Grapplefixtures _ - I_ 1

(b) Side view.

Figure 20. Top and side views of Space Shuttle packaged configurations (second flight) for scenario

number 3. All dimensions are in meters (inches).

4O
30x 103

60 i 1°3 281-

261-

241.

50 22t2ol-

-_40 _ 181-

E30 _ 14t-

._ o 12t.

o #.loI. _-2o

10

O. 0 _ J

0

103

50

_ 30'

o

_ 20'

I:1.

10

Capability to _ |

204 km {110 n.mL)--_

_= ' • ' • _ Capability to

ov 103,ov 104, I1 ...... ,_IL-_. 407 km (220 n.mi,)

OV 102 _ h

_;:yloadc.g location(first aunc)

Payload c.g. location (second launch)

| i I | i 1 i l 1 i i i i i 1 i ]

2 4 6 8 10 12 14 16 18

Distance from forward end of payload bay, m

I_O 2;0 3;o 4_o s;o 6_o 7;0

Distance from forward end of payload bay, in.

(a) X-direction.

lO3

251-Capability to j._

/ 204kin

20 t (110 n.mi.)--_

/'/

a. 5

0 t i | i

-.50 -.40 -.30 -.20 -.10

-_5

.r OV 103, OV 104,

_IL / and OV 105

• " ' " • " " _ Capability 1o

(first launch)

/- Payload c.g. location

(second launch)

0 .10 . 0 . 0 .40 .50

Y distance from centerline of payload bay, m

-_0 -35 (_ 5 I'0 15

Y distance 1tom centerline of payload bay, in.

(b) Y-direction.

120

- e 31'

* i

_c 80- _ 2I

i40 g -o I

e_ e_

E-41- E-1

o o

 8o'

"_ i ,o

_-120L _-3 ,

o

Payload c.g. location

(first launcn)

"_i \ f- Capability to 407 km

\ / (220 n.ml.), OV 102

f-- Capability to 204 km

_0 n.ml\), OV 102

i i

 a ,oa0c.o., ,,on i i !

(second'launch) -* _,-,f,_""'_'_ 204 km (110 nmi.),

_OV 103,104, and 105

J L_ Capability to 407 km 220 n mi )

O_P103, 104, and 10_ " "'

' ' ' ' 1_2 ' 1'6 ' 2 ' ' ' J 3 4 8 0 24 28 _<tO

Payload mass,| kg I I

t'o 20 30 4o _o & ×1o3

Payload mass. Ib

(c) Z-direction.

Figure 21. Space Shuttle payload c.g. locations for scenario number 3.

41
25.35

(998.2)

0.61

(24)

29.21

(1150.1)

33.07

(1302)

1

0.61

I

-,,_ Available volume Propellant

cargo c.g.

location

29.65

(1167.4)

/ 0.61

(24) ,,

31.36

(1234.7)

33.07

(1302)

0.61

,%

Available _-- Propellant

cargo c.g.

location

Figure 22. Volume available on Space Shuttle launches for OTV propellant for scenario number 4. All

dimensions are in meters (inches).

42
7O

6O

5O

40

E

_" 3o

-13,.

2O

10

x 103

x 103 Projected mid-1990's capability with ASRM to 204 km (110 n.mi.)

30-

28-

26

Projected mid-1990's capabilit,i/with ASRM to 296 km (160 n.mi. t

I_-- Total payload c.g. location

(second launch)

24

22

20

Projected mid-1990's capability with ASRM to 407 km (220 n.mi.)

I to

204 km (110 n.mi.}

_18

u}

_ 16

E

N 14

g_ 12

10

OV 103, OV 104,

and OV 105

IIIIIIIIIII

@

ill ii

c.g. location

(first launch)

OV 102

to

407 km (220 n.mi.)

0

0 2 4 6 8 10 12 14 16

Distancefrom forwa_ end ofpayload bay, m

I I I I I I I

0 100 200 300 400 500 600

Distancefrom forwa_ end ofpayload bay, in.

18

I

7OO

Figure 23. Space Shuttle payload c.g. locations for scenario number 4.

43
Support,

1634 kg

Payload,

6681 kg

(80%)

Support,

4 m 3

(4%)

Payload,

89 m 3

(96%)

(a) Mass breakdown. (b) Volume breakdown.

Figure 24. Mass and volume breakdowns for scenario number 1.

Support,

1964 kg

(2

Payload,

6681 kg

(77%)

Support,

6 m3

(6%)

load,

89 m 3

(94%)

(a) Mass breakdown. (b) Volume breakdown.

Figure 25. Mass and volume breakdowns for scenario number 2.

44
Support,

1067 kg

Payload,

5122 kg

(83%)

Support,

11 m3

(13%)

load,

71 m 3

(87%)

(a) Mass breakdown, first launch. (b) Volume breakdown, first launch.

Support,

1973 kg

(56%_

yload,

1559 kg

(44%)

Support,

10 m3

(36%)

Payload,

18 m3

(64%)

(c) Mass breakdown, second launch. (d) Volume breakdown, second launch.

Figure 26. Mass and volume breakdowns for scenario number 3.

45
Payload,

5122 kg OTV fuel,

) 28om3

(24

Payload,

OTV fuel, 71 m3

7545 kg

(47%) (60%)°/

Support,_ _

Support, (16%)

3330 kg

(21%)

(a) Mass breakdown, first launch. (b) Volume breakdown, first launch.

OTV fuel,

19 692 kg

(68%)

Payload, Payload,

1559 kg 18 m 3

(5%) (15%)

Support,

7881 kg

(27%)

OTV fuel, 72 m 3

(59%)

Support,

32 m 3

(26%)

(c) Mass breakdown, second launch. (d) Volume breakdown, second launch.

Figure 27. Mass and volume breakdowns for scenario number 4.

46


Report DocmnentationPage Nal,ona AeronauhcS aqd

Space Admit, s[rat,c,.

l. Report No. 2. Governnmnt Accession No. 3. Reeipient's Calalog No.

NASA TP-3083

4. Title and Subtitle

Launch Vehicle Integration Options for a Large Earth Sciences

Geostationary Platform Concept

7. Author(s)

James L. Garrison and Lawrence F. Rowell

9. Performing Organization Name and Address

NASA Langley Research Center

Hampton, VA 23665-5225

12. Sp{msoring Agency Name and Address

National Aeronautics and Space Administration

Washington, DC 20546-0001

5. Rel)ort Date

July 1991

6. t)erfl)rining Organization Code

8. Perfl_rming Organization Report No.

L-16819

10. Work Unit No.

506-49-21-02

l 1. ('[mtra('l or Grant No,

13. Type of Report and Period Covered

Technical Paper

14. Sponsoring Agency Code

15. Supplementary Notes

i6. Abstract

Concepts are derived for the packaging of the components of a large Earth sciences geostationary

platform to be launched in either the Space Shuttle or the Titan IV Complementary Expendable

Launch Vehicle. Geometric data from a proposed conceptual design for the spacecraft, antenna

sizing results from thermal and structural finite element analyses, and independent mass and

volume estimates are used to deternfine sizes, shapes, and masses of the major platform

components and support equipment. Solid modcqling software is used to evaluate proposed

launch vehicle integration concepts in terms of meeting volume and mass constraints, checking

for interference between components, verifying that center-of-gravity locations fall within vehicle

specifications, and designing suitable interface structures that do not violate any of these

constraints. Construction at Space Station Freedom is assumed, and a space-based orbital

transfer vehicle is determined to be necessary for inserting the spacecraft, once assembled, into

geostationary orbit.

17. Key Words (Suggested by Author(s))

Geostationary platform

Launch packaging

Launch vehicles

Launch compatibility

19. Security Classif. (of this report)

Unclassified

18. Distribution Stat, ement

Unclassified Unlimited

i

20. Unclassified Security Classif. (of this page) 21. 47No. of Pages 22. A03 Price

Subject Category 16

NASA FORM 1626 OCT 86 NASA-Langley. 1991

For sale by the National Technical Information Service, Springfield, Virginia 22161-2171