march 1913 i follow-yn astronomy analyses sortie missions
TRANSCRIPT
Contract NAS8-28144DPD No. 282DR No. MA-04
FinalReportAddendum/ March 1913
Follow-ynAnalyses Astronomy
SortieMissionsDefinitionStudy
(NAS A-CR-31 892) ASTRONOMY SORTIE MISSION N73-22770DEFINITION STUDY. ADDENDUM: FOLLOW-ONANALYSES Final Report (Bendix Corp.)280 p HC $16.00 CSCL 03A Unclai
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Itek
I
Contract NAS8-28144DPD No. 282DR No. MA-04
FinalReportAddendum
AM 1002-1005
March 1973
ASTRONOMYSORTIE MISSIONDEFINITION STUDY
FOLLOW-ONANALYSES
Prepared for:
National Aeronautics and Space AdministrationGeorge C. Marshall Space Flight CenterHuntsville, Alabama
BENDIX CORPORATIONNAVIGATION AND CONTROL DIVISIONDenver, Colorado 80201
ITEK CORPORATIONOPTICAL SYSTEMS DIVISIONLexington, Massachusetts 02173
MARTIN MARIETTA CORPORATIONDENVER DIVISIONDenver, Colorado 80201
FOREWORD
This document is submitted in accordance with the data require-ments of the Follow-On Effort to the Astronomy Sortie MissionsDefinition Study, George C. Marshall Space Flight Center ContractNAS8-28144, Modification Number 2 and the Data Procurement Docu-ment Number 282, Data Requirement MA-04.
This volume is an addendum to the Astronomy Sortie Missions StudyFinal Report issued in September of 1972. This document containsthe results of the follow-on effort and does not cover the re-sults of the original Astronomy Sortie Missions Definition Studywhich are contained in the four volume final report issued inSeptember of 1972.
Comments or requests for additional information should be directedto:
Dale J. Wasserman, PD-MP-AAstronomy Sortie Missions Definition StudyContracting Officer's RepresentativeGeorge C. Marshall Space Flight CenterMarshall Space Flight Center, Alabama 35812
or
William P. Pratt, MS 8102Astronomy Sortie Missions Definition StudyStudy ManagerMartin Marietta Denver DivisionDenver, Colorado 80201
ii
CONTENTS
I. INTRODUCTION . . . . . . . . . . . . . . . . . ... I-1
andI-2
II. STUDY OBJECTIVES .................... II-1
III. GUIDELINES AND PRINCIPAL ASSUMPTIONS . . . . . . . ... III-1
IV. BASIC DATA GENERATED AND SIGNIFICANT RESULTS . . . . . . IV-1A. UV Instruments ..................... IV-1B. Nondeployed Solar Payload . . . . . . . . . . . . ... IV-34C. On-Orbit Access ................... T. IV-76D. Planning Data ................... .. IV-139
V. COST ESTIMATES ................... .. V-lA. Costing Approach, Methodology, and Rationale ...... V-1B. Cost Estimates by WBS Element ............. V-2C. Cost Estimates by Common Items, Telescope, Array Groups,
and Instruments ................... V-39D. Technical Characteristics Data ............. V-56E. Work Breakdown Structure Dictionary and Diagram . . .. V-113
thruV-121
VI. SUGGESTED ADDITIONAL EFFORT . . . . . . . . . . . . . . VI-1and
VI-2
VII. REFERENCES ....................... VII-1
Fiur__________________________________
IV-1 UV Celestial Sphere Viewing . . . . . . . . . . . . IV-7IV-2 UV Viewing Constraints . .............. IV-8IV-3 Instrument Survey Operations . . . . . . . . . . . . . . IV-9IV-4 Launch Window Variations with Inclination and Time of
Year ...... . .. .. . .. . .. .. .. ... . . IV-12IV-5 UV Instrument On-Orbit Operation Cycle . . . . . . ... IV-13IV-6 UV Instrument Refurbishment . . . . . . . . . . . . . . IV-14IV-7 Modified ST-100 in Sortie Lab Airlock . . . . . . . . . IV-16IV-8 ST-100 Instrument Accommodation . . . . . . . . .... IV-16IV-9 ST-100 Platform Installation with and without Pallet .. IV-19
iii
IV-10IV-liIV-12IV-13
UV Outer Gimbal Platform with and without Pallet .TV Outer Gimbal Platform Sizing Study . . . . . . .ASM Solar Payload Deployed X-POP . . . . . . . . .ASM Solar Payload Modified Baseline Deployed X-POP
IV-14 ASM Solar Payload Nondeployed X-POP . .IV-15 ASM Solar Payload Nondeployed X-POP or ZIV-16IV-17IV-18IV-19IV-20IV-21IV-22IV-23IV-24IV-25IV-26IV-27IV-28IV-29
RCS Weight . . . . . . . . . . . . . . .Access Concepts . . . . . . . . . . . . .Photoheliograph . . . . . . . . . . . . .Stratoscope III . . . . . . . . . .Infrared Telescope . . . . . . . . . . .Photoheliograph, Airlock/Hangar ConceptStratoscope III, Airlock/Hangar ConceptIR Telescope, Airlock/Hangar ConceptStratoscope III on Gas Bearing MountPhotoheliograph on Mechanical GimbalPhotoheliograph on Gas Bearing MountStratoscope III on Mechanical GimbalIR Telescope on Gas Bearing Mount . . . .Effect of Wavefront Error on MTF . . .
. . . . . . .
-POP ....
. . . . . .
. . . . . .
. . . . ..
. . . . ..
. . . . ..
. . . . ..
. . . . ..
. . . . ..
. . . . ..
. . . . ..
. . . . ..
IV-30 Normalized Angular Resolution vs Wavefront ErrorIV-31 Gas Bearing Gimbal at System Center of GravityIV-32 Basic Two-Mirror Telescope Relations . . . . . . .IV-33 Obscuration by Fold Mirror . . . . . . . . . . . .IV-34 Two Fold Options for Photoheliograph . . . . . . .IV-35 Obscuration Analysis for Folded Photoheliograph .IV-36 Obscuration Analysis for Folded IR Telescope . . .IV-37 Change in Resolution of IR Telescope vs Obscura-
tion . . . . . . . . . . . . . . . . . . . . . .IV-38 Effect of Fold Distance on Baseline IR TelescopeIV-39 Obscuration Analysis of Folded Stratoscope IIIIV-40 Diffuse Reflection Off Baffle in Folded SystemIV-41 Problem of Warm Structures in IR Telescope Sensor
Field of View . . . . . . . . . . . . . . . . . .IV-42 Airlock/Hangar On-Orbit Access ConceptIV-43 Concept CGs vs Shuttle Orbiter CG Constraints . . .IV-44 Turnaround Schedule . . . . . . . . . . . . . . . .IV-45 Schedule for Telescopes, Arrays, and InstrumentsIV-46 Schedule for Subsystems Used with Intermediate
Telescopes and Arrays . . . . . . . . . . . . . .IV-47 Schedule for Subsystems Used with Small UV Instru-
ments . . . . . . . . . . . . . . . . . . . . . .V-l WBS Diagram .
iv
IV-21IV-29IV-37IV-41IV-45IV-49IV-72IV-79IV-81IV-83IV-85IV-89IV-91IV-93IV-95IV-97IV-101IV-103IV-105IV-110IV-110IV-111IV-112IV-113IV-115IV-115IV-116
IV-116IV-117IV-118IV-119
IV-120IV-123IV-127IV-140IV-142
IV-143
IV-144V-121
I7 _
Table…__ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _
IV-1 Small UV Instruments .IV-2 UV Instrument Mission Mode Implications . . . . . . .IV-3 ST-100-SI Platform Characteristics . . .IV-4 ST-100 Platform Modifications/Improvements . . . . . .IV-5 UV Platform Capabilities . . . . . . . . . . . . . . .IV-6 UV Outer Gimbal Platform Installation Weights . . . .IV-7 UV Instrument Accommodation Requirements . . . . . . .IV-8 UV Instrument Support Hardware . . . . . . . . . . . .IV-9 Baseline Configuration Weights . . . . . . . . . . . .IV-10 Modified Baseline Configuration Weights . . . . . . .IV-11 Nondeployed X-POP Configuration Weights . . . . . . .IV-12 Nondeployed X-IOP or Z-POP Configuration Weights . . .IV-13 Stabilization Momentum Requirements, X-IOP . . . . . .IV-14 CMG Momentum Requirements, X-IOP . . . . . . . . . . .IV-15 RCS Momentum Requirements, X-IOP . . . . . . . . . . .IV-16 RCS Impulse Requirements, X-IOP . . . . . . . . . . .IV-17 Summary of Angular Momentum Requirements for CMG Con-
trol . . . . . . . . . . . . . . . . . . . . . . . .IV-18 Summary of Angular Momentum Requirements for RCS Con-
trol . . . . . . . . . . . . . . . . . . . . . . . .IV-19
IV-20IV-21IV-22IV-23
IV-24IV-25IV-26IV-27IV-28IV-29IV-30IV-31V-1V-2V-3V-4V-5V-6
Weight Summaries for CMG and RCS Monopropellant Sys-tems . . . . . . . . . . . . . . . . . . . . . . . . .
Power Summaries for CMG Systems and RCS . . . . . . . .Comparison of Solar Payload Concepts . . . . . . . . . .Summary of Telescope Weights for Hangar Concept . . . .Summary of Telescope Weights for Gas Bearing and Mechan-ical Gimbal Concepts ................
Hangar Concept Weight Summary . . . . . . . . . . . . .Gas Bearing Concept Weight Summary . . . . . . . . . . .Mechanical Gimbal Concept Weight Summary . . . . . . . .Baseline Weight Summary .. . . . . . . . .Access Operations ...................Operations Effectiveness . . . . . . . . . . . . . . . .Cost Estimates .....................Concept Comparison ...................Subsystem Equipment and Operations . . . . . . . . . . .UV Instruments Subsystems Equipment . . . . . . . . . .IR Telescope, Stratoscope III, and Photoheliograph .Array Groups, A, B, C, D & E . . . . . . . . . . . .Small UV Instruments and UV Instrument Groups . . .WBS Dictionary ...................
IV-4IV-23IV-27IV-28IV-30IV-31IV-32IV-33IV-39IV-40IV-47IV-51IV-57IV-58IV-61IV-61
IV-70
IV-70
IV-73IV-74IV-75IV-107
IV-108IV-128IV-129IV-130IV-131IV-132IV-133IV-134IV-138V-40V-45V-48V-51V-54V-114
v
.
I. INTRODUCTION
This document includes the results of the design analyses, tradestudies, and planning data that were developed during the 6-monthfollow-on effort to the Astronomy Sortie Mission Definition Study,Contract NAS8-28144. This study was performed for the George C.Marshall Space Flight Center by an industry team led by the DenverDivision of Martin Marietta and supported by the Bendix and ItekCorporations. The Bendix Corporation was responsible for thepointing, control, and stabilization analyses for the astronomyinstruments and for the Shuttle stabilization system. The ItekCorporation was responsible for the definition of the astronomyinstruments and the analyses of their performance and operations.
To understand the information contained in this document it isnecessary to have the background information that was derivedduring the original 9-month Astronomy Sortie Mission DefinitionStudy. This information is contained in the four-volume finalreport (Ref 1).
The primary objective of the 6-month follow-on effort was to pro-vide more in-depth analyses of several key areas identified dur-ing the original study. A summary of each of the key areas ispresented in the following paragraphs.
UV Instruments - The purpose of this task was to analyze and con-ceptually define a group of small ultraviolet (UV) instruments,and to define the support hardware and operations necessary toaccommodate the instruments. As a result of this task, eightsmall WUV instruments that might be flown on the sortie missionswere identified. It was recommended that these instruments bedesignated as "flight of opportunity" instruments that could beflown on any inertial attitude sortie mission that had excesscargo bay volume and payload weight capability. The support hard-ware defined was based on the "flight of opportunity" mode ofoperation, and the major piece of support hardware identified wasan instrument mount that would house instruments up to 1.17 meters(46 in.) in diameter.
I-1
NondepZoyed SoZar PayZoad - The purpose of this task was to eval-uate solar payload configurations that do not require deployment
of the entire payload out of the cargo bay. Based on the analy-ses performed, we recommend a solar payload configuration that
does not require deployment, and hence does not require a deploy-ment mechanism. In addition, we recommend stabilizing the Shuttle
orbiter in a Z-POP inertial attitude (Shuttle's longitudinal axislies in the orbit plane with the Z-axis perpendicular to the or-
bit plane) throughout the 7-day mission.
On-Orbit Access - The objective of this task was to investigateseveral alternative methods of providing on-orbit shirtsleeve ac-cess to the focal planes of the infrared (IR), stratoscope, andphotoheliograph telescopes. The recommended approach for pro-viding the on-orbit shirtsleeve access is a configuration wherethe telescope optics are in the vacuum environment external tothe Sortie Lab, and the focal plane instruments are located in
the Sortie Lab. Light is brought from the telescope opticsthrough a passage in the Sortie Lab bulkhead and made availableto the instruments. A mechanical gimbal provides 2 degrees ofwide angle pointing, and fine pointing and stabilization is ob-tained by image motion compensation within the telescopes. Dur-ing the study, the IR telescope was dropped from consideration
for on-orbit access because of the potential problems associatedwith the cryogenic temperatures.
Planning Data - The planning data developed include the cost and
schedules associated with the astronomy instruments and/or theirsupport hardware.
Cost Estimates - The costs are defined in terms of DDT&E, produc-
tion, and operations, and are presented in a parametric fashionthat will allow any payload or payload combination to be costed.
I-2
II. STUDY OBJECTIVES
The specific objectives of this study were:
1) Analyze and conceptually design a group of small survey-typeUV instruments and define the support hardware and operationsnecessary to accommodate these instruments during the AstronomySortie Missions. The UV instruments baselined for the studywere the Tifft, Morton, and Carruthers instruments that weredesigned for the ST-100 platform. In addition, derivativesof these baseline instruments were to be investigated for ap-plicability to the sortie missions. The support hardware tobe defined was the ancillary hardware that was required inaddition to the Sortie Lab and pallet. The operations to bedefined included all phases of the mission including ground,prelaunch, launch, on-orbit, and postflight operations;
2) Evaluate the desirability of deploying payloads out of theShuttle cargo bay. The objective was to examine the require-ment for rotating the entire solar payload through a 90-degreeangle in order to satisfy the viewing requirements. Thistask was initiated because of a change in the ground rulesrequiring that the deployment mechanism weight (approx 2000lb) and volume (approx 5 linear ft of cargo bay) be chargedto the payload. Previously, the weight and volume were chargedto the Shuttle orbiter;
3) Evaluate the advantages and disadvantages of providing on-orbit shirtsleeve access to the focal planes of the IR,stratoscope, and photoheliograph telescopes. The primaryreason for this task was the strong desire by the scientificcommunity to have on-orbit access to the focal planes of thetelescopes. Consequently, several alternatives were to beinvestigated that would provide this capability;
4) Provide parametric planning data that will enable alternativeflight schedules and instrument complements to be evaluated.Because this study does consider representative type astron-omy instruments and "straw-man" flight schedules, it was de-sirable to develop the planning data in a parametric fashionthat would allow rapid evaluation of many alternatives.
II-1
III. GUIDELINES AND PRINCIPAL ASSUMPTIONS
The guidelines and principal assumptions for the study were pro-vided by the NASA/MSFC Contracting Officer's Representative as apart of the contract statement of work. They were:
1) The Astronomy Sortie Missions Definition Study Final Report(Ref 1) was the baseline for the follow-on effort;
2) The Sortie Can Conceptual Design document (Ref 2) was thebaseline for the definition of the Sortie Lab and pallet;
3) The Space Shuttle Baseline Accommodations for Payloads docu-ment (Ref 3) was the baseline for the definition of the SpaceShuttle interfaces and capabilities;
4) The UV instruments that were to be considered during the fol-low-on effort were those instruments designed for operationwith the ST-100 platform and included--the UV camera proposedfor use in the UV photographic survey experiment by Dr. Tifftat the University of Arizona, the all-reflective spectrographproposed for the UV stellar spectrometry experiment by Dr.Morton of Princeton University, and the Schmidt image con-verter spectrograph proposed for use in the far UV spectrometryexperiment by Dr. Carruthers at the Naval Research Labs;
5) Theroperational mode for the sortie missions was 7 days dura-tion with two scientific crewmen available for 24-hour-per-day operation of the experimenter.
III-1
BASIC DATA GENERATED AND SIGNIFICANT RESULTS
The general results of the follow-on effort are: designate thesmall UV instruments for the astronomy sortie missions as flight-of-opportunity instruments, and provide accommodations for theseinstruments using state-of-the-art ancillary hardware; do not re-quire deployment of the astronomy payloads from the Shuttle car-go bay; and provide on-orbit access to the focal planes of thephotoheliograph and stratoscope payloads at a modest increase incost and weight.
The specific tasks that are covered in this chapter include: UVInstruments; Nondeployed Solar Payload; On-Orbit Access; andPlanning Data.
A. UV INSTRUMENTS
The emphasis during the follow-on study was on the Sortie Mission
concept and its adaptability for providing the means to conductUV experiments. NASA-MSFC identified a group of small UV surveytype telescopes as representative telescopes to be flown. Thefirst task in the study was to examine the experiment definitionscontained in the available documentation on these telescopes withemphasis on the aspects that would affect the Astronomy SortieMission Definition Study. Itek was responsible for examining theexperiments proposed and comparing them with other small or med-ium sized astronomy instruments formerly associated with rocketflight, balloon, or small unmanned satellite programs. Bendixperformed the comparison of pointing and stabilization require-ments of the instruments with the available performance data onthe NASA-MSFC ST-100 Stellar Instrument (SI) platform. An alter-native configuration for an instrument platform and mount was alsoconsidered. Mission profiles for the WUV survey instruments wereestablished to cover all mission phases, identifying those opera-tions that influence the instrument, Sortie Lab, Pallet, or Shuttledesigns or interfaces.
1. Ground Rules and Assumptions
The principal ground rules and assumptions that were used duringthe performance of this task were provided by the NASA/MSFC COR,or were carried over from the earlier portion of the study. Theywere:
IV-1
IV.
1) The baselined UV survey instruments consisted of the ST-100stellar payload experiments. These experiments were consid-ered as representative types of UV telescopes that might beflown. The original group consisted of--the two 15-centimeter(6-in.) survey imaging cameras developed by Dr. Tifft at theUniversity of Arizona (Ref 4), the objective grating telescopedeveloped by Dr. Morton of Princeton University, and theSchmidt telescope for spectroscopy developed by Dr. Carruthersat the Naval Research Laboratory;
2) Other instruments, particularly those designed for rocketflight and balloon-borne interfaces were considered;
3) Detailed designs of the proposed UV instruments were thesubject of other studies or programs;
4) Photographic film was the preferred recording media to beused with each experiment;
5) Film cassettes would be loaded with sufficient film to makeinflight replacement unnecessary during the 7-day mission.
2. Instruments Considered
In the period since the ST-100 experiments were first conceived,other small telescopes have been proposed or flown that couldfulfill the UV experiment mission objectives. Dr. Carruthershas a proposal for a modified Schmidt image-converting spectro-graph, which is a larger aperture version of his original instru-ment and provides increased capability. He has also proposed two40-centimeter (15.8-in.) complementary telescopes that providesimultaneous spectroscopy and imagery. The physical configura-tion is similar to the original survey imaging (Tifft) cameras.A coordination meeting with Dr. Tifft at the University of Arizonaindicated that an aperture increase for the survey imaging cameraswould logically provide a times two magnification to allow use of70 mm film. Alternatively, the film would remain 35 mm and theaperture increase would be used to decrease obscuration. The all-reflective spectrograph developed by Dr. Morton has been supersededby a larger payload for the Aerobee 170 rocket. This payload has alarger aperture, narrower field-of-view UV Echelle spectrograph.In addition, a balloon-borne instrument, a narrow-band scanningspectrometer described in NASA's Blue Book, Volume III, Section1 (Ref 5), was considered as a candidate instrument for the Astron-omy Sortie Mission. These instruments could fly on a Shuttle
IV-2
Sortie mission to perform celestial surveys and provide a base-line concept for a sky survey capability. Performance characteris-tics and accommodation requirements are shown in Table IV-1 for
the baselined experiments and for the similar small and mediumsize UV telescopes.
a. Tifft 15-Centimeter (6-in.) Survey Camera - The UV photo-graphic survey provides high quality photographic images of star0fields and emission regions taken in UV light from 1700 to 3000 A(170 to 300 nanometers). The camera system consists of twin UVcameras, which are identical except for the internal filteringused. One camera operates in a near-UV band (2300 to 3000 A);the other operates in the mid-UV band (1700 to 2100 A). Primaryemission features to be studied are the Magnesium II doublet at
0 0
2800 A, and the Calcium III intersystem line at 1909 A. The
stellar continuum observations will enable the UV energy distri-bution of a large number of stars of Spectral Type F or earlierto be examined, and will also provide a general classificationscheme based on image photometry. Three exposures will be takenin each band (camera). A wide-angle (20 to 30 deg) star-fieldcamera will be used with the UV survey cameras to take one ex-posure for each exposure cycle of the UV experiment in orderto assist in identifying the field of view. A sun sensor willdetect the presence of the sun or bright Earth, with 30 degreesof the look angle of the experiment.
b. Morton AlZ-RefZective Spectrograph - This far-UV stellarspectrometry experiment uses an all-reflective instrument with
an objective grating and Lithium Fluoride coated optics. Thespectral range observed will be from 900 to 1800 A (90 to 180nanometers) with dispersion spectra of 0, B, and A stars downto the sixth magnitude. Recording will be on photographic filmwith a resolution of 0.5 A. The spectrograph has a 6-degreefield of view. When the instrument is tracking the desired source,shutters will be opened and UV-sensitive film will be exposed fora period of 10 to 20 minutes. An ion chamber will be included tomonitor the local flux of UV at 1216 A in the direction of thesource. If local flux is high, the experiment exposure duration
will be shortened to prevent film fogging.
c. Carruthers Image-Converting Spectrograph - This far-UV stel-lar spectrometry experiment uses a Schmidt-type electronic imageconverter with objective gratings. The objective of this experi-ment is to obtain spectral and photometric data in the far-UVrange by means of wide-angle (15 deg) sky surveys using electrono-graphic techniques. The sensitive region of the spectrum will befrom 1200 to 1800 A (120 to 180 nanometers) with data recorded ona nuclear emulsion film. Exposure durations of 2.5, 6.25, 15, 40,100, 250, and 625 seconds will be used. Analysis of the data willallow determination of the far-UV brightness and spectral distri-butions of a large number of early-type stars. The experiment
IV-3
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data will help in determining effective temperatures, chemicalcompositions, and sources of opacity in these stars. Quantita-tive measurements of the absorbing properties, composition, den-sity, distribution, and temperatures of gases composing the inter-stellar medium (including galactic nebulae) will also be made.
d. Carruthers (Modified) Image-Converting Spectrograph - Thisexperiment will be a modification of the far-UV stellar spectrom-etry experiment described above. The aperture has been increasedfrom 10 to 15 centimeters, spectral coverage has been expanded toinclude 300 to 2100 A (30 to 210 nanometers), and angular resolu-tion is improved to 7.3 x 10-5 radians (15 sec). The instrumentwill be physically larger and will accommodate improved features,such as a removable slit collimator for diffuse source spectro-scopy, a rotating mount to introduce three different gratingsinto the energy path, and a sliding holder for Schmidt correctors.
e. Carruthers 40-Centimeter Telescopes - This experiment incor-porates two 40-centimeter aperture telescopes, one to performspectroscopy, the other to perform imagery at the same time on acommon target. The internal-grating electronsgraphic spectrographwill cover the spectral range of 900 to 1800 A (90 to 180 nanom-eters). It will be mounted at the focus of a Cassegrain-typetelescope, which will be equipped with an internal star trackersystem that will pivot the Cassegrain secondary mirror to keepthe observed image on the spectrograph slit. Field of view forthe spectrographic telescope will be 1.7 x 10-2 radians (1 deg).The imagery telescope will use a semitransparent-photocathodeelectronographic camera to obtain an image of the source. Thecamera will be placed at the focus of a Ritchey-Chretian orCassegrain telescope. Interchangeable photocathodes will be usedwith a rotating holder to provide spectral coverage from 900 to2100 A (90 to 210 nanometers). An externally generated errorsignal can be used to correct for pointing jitter by modulatingthe voltage on the deflection coils around the image tube. Fieldof view for the imagery telescope will be 2.8 x 10-2 radians(1.6 deg). The telescopes will provide complementary data, andwill be mounted on a common platform. Angular resolution ofeither telescope will be 5 x 10-6 radians (1 sec).
IV-5
f. Morton UV EcheZZe Spectrograph - The 15-centimeter apertureEchelle Spectrograph is an updated rocket-borne version of theMorton spectrograph described above. Spectral coverage for thisinstrument has been shifted and expanded to cover 1200 to 2300 X(120 to 230 nanometers); field of view has been reduced to 1.7 x10-2 radians (1 deg).
g. Kondo Scanning Spectrometer - The telescope spectrometer is
a versatile, high-resolution system for analyzing stellar UVspectra over a range of 2700 to 3000 X (270 to 300 nanometers).The instrument (an adaptation of a balloon-borne instrument) willbe used to investigate unique emission features, such as the
Magnesium II doublet of 2795 X and 2802 X (279.5 and 280.2 nanom-eters), or the Carbon IV and Silicon IV emission lines in the farUV. A fine scanning mode will be used for observing narrow ab-sorption lines, such as those caused by interstellar molecular
hydrogen. The Ebert-Fastie grating-optics configuration of thespectrograph will provide a spectral resolution of 0.1 X. The
telescope will use a tilted aplanatic Cassegrain optical config-uration, and will provide an angular resolution of 9.7 x 10-6radians (2 sec).
3. Mission Analysis
The intent of the small astronomical UV instruments is primarilyto study the complete celestial sphere at wavelength ranges notobservable by ground-based instruments. Results of the studywill be used to establish a dependable base of reference data andto make it possible to identify unusual celestial sources. Gen-erally, the UV instruments do not impose restrictions on missionorbit altitude or inclination with respect to the equator.
Missions are required at various times of the year to permit sur-vey of the entire celestial sphere and to view along the entireGalactic plane. Maximizing dark time is very important to theUV instruments (except the Kondo Scanning Spectrometer).
a. Operational Constraints - Viewing of the celestial sphere bythe UV instruments is restricted about the sun, Earth, and moon,and to the dark-time period of the orbit. In addition, targetsof interest for several of the instruments lie about 15 degrees
from the Galactic plane. Figure IV-1 shows the viewing restric-tions and the position of the galactic plane for an orbit on
July 15, 1977, during a new-moon condition.
IV-6
0.5 radian (28.5 deg) orbit inclinationto equator. Zero beta angle orbitinclination. Altitude 463 km (250 n mi).Cannot view within 0.524 radian (30 deg)of sun, Earth, and moon. July 15, 1977.
300 240 180
O Moon
Sun
Right Ascension
120
Figure IV-1 UV CeZestiaZ Sphere Viewing
The orbit inclination to the Earth's equator is 0.5 radian (28.5deg) with a zero Beta angle orbit inclination. The altitude is463 kilometers (250 n mi). The shaded areas that cannot be viewed
about the sun, Earth, and moon are the limits of 0.524 radian(30 deg). For the Kondo Scanning Spectrometer, which has view
limits of 0.79 radian (45 deg), each of these restricted zoneswill be 0.26 radian (15 deg) larger in radius than shown. How-
ever, the Kondo instrument (unlike the other UV instruments) mayview in daylight, so the impact of the larger viewing limits is
offset somewhat. For the UV instruments with the 0.524 radianviewing limit, data may be acquired only while in the dark side
of the orbit. The viewable portion of the celestial sphere dueto this restriction is indicated. It may be noted that at other
times of the year the position of these restrictions will be dif-ferent. Thus, by flying at other times, the entire celestial
sphere may be observed and specific targets of interest near theGalactic plane may be viewed.
IV-7
360 60 090
60
30
3O030
-60
9090
The nighttime only viewing period and the 0.524 radian (30-deg)look-angle constraint about the Earth are shown in Figure IV-2for a 463 kilometer (250 n mi) orbit.
0.367 radian (21 deg) angle betweenlocal horizontal and Earth's limb
Sun and MaximumMoon in Viewing ViewingNew Moon Period Angle ConePosition 2.83 radians
(162 deg)
Zero beta angle orbit inclination;463 km (250 n mi) altitude; cannotview within 0.524 radian (30 deg)of sun, Earth, and moon.
Figure IV-2 UV Viewing Constraints
At the altitude shown, the angle between the local horizontaland the Earth's limb is 0.367 radian (21 deg). Since the UV in-struments must view no closer than 0.524 radian (30 deg) of theEarth, the maximum available is a full-angle cone of 2.83 radians(162 deg), which is derived by accounting for the 0.367 radian(21 deg) angle between the local horizontal and the limb of theEarth, as well as the 0.524 radian (30 deg) limit to the limbimposed on the UV instruments. For the Kondo Scanning Spectrom-eter, the viewing limits are 0.79 radian (45 deg), and the re-sulting viewing cone is 2.31 radians (132 deg). With the sunand moon in new-moon condition as shown, the 2.83 radians (162deg) viewing cone is available whenever the spacecraft is in thenight view period. With a zero beta angle orbit inclination,the maximum duration of this night viewing period is about 35.3
IV-8
minutes for an overall orbit period of 93.7 minutes. The con-straint of operating these instruments only while in the darkperiod of the orbit has the greatest impact on the amount of sci-entific data that may be acquired. This constraint applies toall of the UV instruments analyzed except the Kondo ScanningSpectrometer. The limitation of viewing no closer than 0.524radian (30 deg) to the sun, Earth, and moon, can be observed with-out reducing useful data.
b. Survey Operations - While the survey instruments have what isconsidered wide fields of view from the astronomical standpoint(4 to 15 deg), each exposure sequence on one source region coversonly 0.006 to 0.03 steradian. Several of the UV instruments havean objective to survey the entire celestial sphere. Figure IV-3shows the number of fields in the celestial sphere for variousfull-cone fields of view.
Schmidt 15-degImage-Converting FlightSpectrograph(Carruthers)
All-ReflectiveSpectrograph 3 Flights(Morton)
Two ObservationPeriods Per Orbit;200 ObservationPeriods Per 7-DayMission
4 Flights
500 1000 1500 2000 2500
Total Fields of View in Celestial Sphere
Figure IV-3 Instrument Survey Operatzons
15-
10-
Full-ConeField ofView,degrees
5-
0
0 3000
IV-9
I I I I i ·
Four of the instruments that have been proposed are indicated onthe figure, together with the number of flights required to com-plete the survey for that telescope. In making these calcula-tions, the individual fields were overlapped in a hexagonal pat-tern to provide full coverage and coordination between areas.Two observation periods per orbit were assumed, with a total of200 observation periods in each 7-day flight. Note that full-sky is practical in sortie missions for all fields of view greaterthan about 5 degrees. For narrow fields of view, the number offlights required is greater than 15. Note that in a galacticsurvey program, no high-rate maneuvers of the vehicle are re-quired. The exposure program can be carefully prepared so as torequire only slight spacecraft maneuvers between exposure se-quences, or even no maneuvers at all.
Another special objective for the small UV instruments cannot bepreplanned. Of critical astronomical interest is whatever datacan be obtained during the early phases of novae ("exploding"stars) in the photon energy ranges not available to ground-basedinstruments. Should a survey mission be in progress when a brightnova is detected, the survey program would be abandoned in favorof a detailed study of this unusual object, with all availableinstrumentation. However, this may also require sizable vehiclemaneuvers. To minimize the time lost to reorient the instru-ments, it appears to be preferable to provide as much of this re-orientation as possible with the platform gimbals, and limit thespacecraft reorientation to a roll angle change. This leads tothe same platform mounting arrangment considered for the X-POP-Imode, where the wide-freedom gimbal axis is mounted perpendicularto the Shuttle longitudinal axis.
c. Orientation Requirements - The full-hemisphere requirementfor instrument orientation results in some specific acceptableplatform/spacecraft arrangement, depending on the orbital atti-tude of the Shuttle. The case where the Shuttle is maintainedin a fixed inertial orientation, with the longitudinal axis per-pendicular to the orbital plane (X-POP-I) was considered first.The only spacecraft orientation freedom available in the rollaxis, i.e., the payload bay can be oriented so that it opens outto any direction in the orbital plane. Since the ST-100 platformhas only one gimbal axis with wide freedom, the instrument linesof sight must be perpendicular to this platform axis, and theplatform gimbal axis itself must be perpendicular to the space-craft roll axis. More specifically, it means that the gimbalaxis with wide freedom must be oriented parallel to the directionof the Shuttle wings. This provides the possibility to orientthe line of sight in any direction in the celestial sphere, by acombination of platform gimbal angle and Shuttle roll angle.
IV-10
If the Shuttle is maintained in a fixed inertial orientation, butwith the longitudinal axis in the orbital plane (X-IOP-I), thepayload bay can be oriented so that it opens out toward any direc-tion in the celestial sphere by a combination of spacecraft rolland a rotation of the Shuttle longitudinal axis within the orbitalplane. This means that instruments looking out of the payload baycan be oriented to any direction without the need for wide gimbalfreedom on the stabilizing platform. Therefore, if no additionalconstraints on Shuttle orientation arise from other instrumentsin the payload bay (or from some currently undefined system re-quirements), the limited-freedom platform orientation capabilityrelative to the Shuttle spacecraft axes is not a constraint. Ifexternal constraints exist that limit the spacecraft freedomwithin the X-POP-I or X-IOP-I attitudes, it does not appear thata platform like the ST-100 with a single-wide-freedom axis can beefficiently used to carry the UV instruments. Noninertial atti-tudes, such as the X-POP-ZLV or X-IOP-XLV, are not easily handledby a platform with limited gimbal freedoms. In this case, wherethe spacecraft is rotating at approximately 1.16 mrad sec - 1 (4deg/min), the +15-degrees-freedom axis could limit exposure se-quences to less than 10 minutes.
d. Launch Constraints - Figure IV-4 shows the variations inlaunch window for 463 kilometer (250 n mi) altitude orbits atvarious inclinations that will result in averages of 15 and 30minutes of dark time per orbit.
The dark-time periods are averaged over the entire 7-day mission;therefore, some individual orbits in a mission may have the maxi-mum dark-time duration attainable of about 35 minutes, whileothers have no dark time at all. Note that for 15 minutes aver-age dark time, there is no restriction on launch time at an or-bit inclination of 28.5 degrees. At 60 degrees, there is norestriction for about 40 days at both of the equinoxes, and, atthe solstices, about 21.5 hours in each day are available forlaunch. At 90 degrees inclination, the launch window is nearlyconstant throughout the year, varying from about 17.9 hours toabout 18.4 hours per day. To achieve 30 minutes average darktime, a restriction in launch time for an orbit inclination of28.5 degrees must be observed for about 40 days at the summerand winter solstices. The launch window in this instance is about22.7 hours. At 60 degrees, the minimum launch window is about17.7 hours, increasing to about 24 hours at the equinoxes. For90-degree inclined orbits, the launch window is from about 8.7hours to about 10 hours for the 30 minute dark-time case. It may
IV-ll
be observed that for flights in which the primary experiments donot require other launch times, the launch windows for all orbitinclinations to provide 30 minutes average dark time are not un-duly restrictive. The worst limit leaves about 8.7 hours avail-able for launch each day, and in most cases, the launch windowis 18 or more hours per day.
15 minutesDark TimePer OrbitAverage in7 Days at250 n miAltitude
- i = 28.5°
D-X-\
-- i = 60°
i = 90O
i OrbitInclinationto Equator
24
16
8
J F M A M J J A S O N D
Calendar Year
30 minutesDark TimePer OrbitAverage in7 Daysat 250 n miAltitude
~! % _of~i %o-
1 i = 60°
_
= 90°
J F M A M J J A S O N D
Calendar Year
Figure IV-4 Launch Window Variations with IncZination and Time of Year
e. On-Orbit Operations - Typical on-orbit operations sequencesfor the three ST-100 UV instruments (Tifft, Carruthers, andMorton) are shown in Figure IV-5.
The overall cycle starts for a selected target as the spacecraftenters the dark period of the orbit. Note that this cycle isarranged to provide relatively short exposures using the SchmidtImage Converter Spectrograph, and that the other instruments arecoordinated with the Schmidt Spectrograph during the medium- andlong-duration exposures. It is necessary that film advance forthe three telescopes be performed simultaneously, as shown, when
IV-12
24
LaunchWindowEach Day,Hours
16
8
imaging by two or all of the instruments is concurrent. Thisnecessitates the idle times when exposures are not of equaldurations. On completion of one cycle, a second target is se-lected and a portion of the cycle is performed until the space-craft leaves the orbital dock period. For orbits with less than28 minutes dark time, the cycle may be shortened to the availabletime.
EnterOrbitalDarkTime
uverall
Sequence 1
(high intensity) Schmidt Image ConverterSpectrograph (Carruthers)
2.5 6.25 15 40 1 min 4 min 10 minSec sec Sec sec 40 sec 10 sec 25 sec
Sequence 2
Camera Idle Time
Film Advance (about 10 sec)
Film Exposure Time
(low intensity)
All-Reflective Spectrograph(Morton)
10 to 20 min
Figure IV-5 UV Instrument On-Orbit Operations CycZe
f. Instrument Refurbishment - The schedule of events requiredfor refurbishing the UV instruments has a duration of 10 weeks(land-to-land), as shown in Figure IV-6. In addition, a post-flight period of 16 weeks has been provided for scientific dataprocessing and analyses, and for planning subsequent missionsusing these instruments.
IV-13
Legend
O
FM -I
--2
i :
Weeks
I 1 2 13 1 4 1 5 1 6 I 7 18 19 110 I1 14 18 1221 26
Land, Safe, Remove Payload
Service, Pack and Ship Payload to Payload Integration Center
Inspect Payload
Remove Experiment Equipment
Prep Sortie Lab and Pallet
Install Telescopes
| Flight Readiness Tests
Ship to Launch Site, I Prep for LoadingLoad in Orbiter
and CountdownPrelaunch
Launch, IOn Orbit
Land, Postflight
59 Man-Months
___I---- __ __l_7- _
21 Man-Months
Figure IV-6 UV Instrument Refurbishment
The minimum manpower estimates shown are for a three-man crewdedicated to one UV instrument for the entire turnaround period,including postflight evaluation. During the land-to-land timeperiod, another specialist is required to interface with theSortie Lab and pallet. During instrument removal, installation,and test at the Payload Integration Center, an integration spe-cialist is also necessary. The maximum manpower estimates arebased on a UV flight configuration including three instruments.In this case, three crewmen are dedicated to each instrument overthe entire 26-week period, with one specialist to provide thecoordination for all instruments with the Sortie Lab and pallet.Three additional men (one for each instrument) are required dur-ing operations at the Payload Integration Center.
IV-14
12
GroundCrew
8
4
4. Mission Mode Alternatives
The UV instruments may be flown on compatible Shuttle sortie mis-sions that can accommodate the instruments and the mount. As pres-ently conceived, these instruments require an inertially-stabil-ized orbiter and minimal crewman participation to perform nonrepe-titive and periodic monitoring functions. Orbit constraints andlook-angle restrictions are not harsh, but launch window limitsmay be desirable, depending on orbit inclination and time of yearof the mission. Two modes of operation were considered during thestudy.
The first mode considered was a dedicated flight where the exper-iments and platform would be used on flights that were programmedfor their use. The scientific crewmen would be trained, and giventime to interface with the survey telescopes using an airlockintegral with the Sortie Lab. The second mode considered was touse the platform and instruments on a flight of opportunity, suchas an Earth resources mission or with a biology experiment. Aflight of opportunity would imply a minimum of crew training onplatform operation, almost no control and display requirements,and no EVA for data retrieval. The crew would be expected todeploy and store the platform and start the experiment, but op-eration would then be entirely automatic.
a. Configuration AZternatives - The three original ST-100 ex-periments and the five additional experiments were consideredfor performance and physical compatibility. Each experimentcould be flown individually. Two groupings of experiments wereconsidered: one group consisted of the original ST-100 package;the second contained the two Carruther's 40-centimeter telescopes.The experiments and their groupings were reviewed to determinetheir adaptability for use in the Sortie Lab airlock, with theST-100-SI platform, or on an outer gimbal platform.
A study was made to determine the feasibility of locating theST-100 platform in the Sortie Lab airlock defined in "Sortie CanConceptual Design," MSFC ASR-PD-D-72-2, March 1, 1972 (Ref 2).Because of the small size of the airlock, the unmodified platformwith its group of experiments will not fit. However, individualST-100 experiments can be accommodated by modifying the platformstructure, relocating equipment, proper counterbalancing of theplatform, and by minor modifications to the instruments. Themodified configurations are shown in Figure IV-7. None of theother UV experiments defined in this study will fit into the air-lock.
IV-15
___J 40 in.-j dia
Plat-form
60 in.
Deploy-mentMechanism
6-in. Survey Cameras (2)(Tifft)
All-Reflective Spectrograph(Morton)
Image-Converting Spectrograph(Carruthers)
Figure IV-7 Modified ST-100 in Sortie Lab AirZock
Each of the UV experiment groups defined in this study can physi-cally be accommodated on the modified ST-100 platform if not con-strained within the airlock. Three configurations are shown inFigure IV-8.
@~nm@E _ V1'
6-in. Survey CamerasAll-Reflective SpectrographImage-Converting Spectrograph
Internal GratingSpectrographImaging Camera
Scanning Spectrometer
Figure IV-8 ST-100 Instrument Accommodation
IV-16
In each case, except for the original ST-100 experiment comple-ment, the platform structure must be modified and equipment re-
located. Some of these modifications are unique to the experi-ments due to their shapes and sizes. The counterbalancing re-quired in each case is unique. The ST-100 platform is shown inFigure IV-9 installed in the Shuttle payload bay on the pallet
and without the pallet. It was assumed that sortie flights wouldgenerally have a pallet associated with the Sortie Lab. Experi-ments, such as biology, would not necessarily require a pallet,and the UV instrument platform would have to provide its own sup-porting structure to the Shuttle. In either case, the installa-tions minimize physical interfaces with the Shuttle to facilitatefast turnaround. This is particularly important when flying theexperiments on flights of opportunity. An outer gimbal platform(Fig. IV-10) could be installed in the Shuttle payload bay on thepallet and without the pallet, similar to the ST-100 platform in-stallation.
b. Dedicated vs Flight of Opportunity - The platform and itsinstruments can benefit from the presence and dedicated partici-pation of an astronomer only if: the results of the observationsare available in near real time; the astronomer can dedicate asignificant amount of his time to examining the results; and ameaningful reaction to those results can be taken. While allthree of these conditions could be met, it seems very unlikelythat any of them will be. To provide near-real-time results itwould be necessary to remove the film frequently, have equipmentfor processing the film, and provide facilities for film inspec-tion. This equipment is bulky, would consume valuable space in
the Sortie Lab, and its use would preclude the possibility ofgeneral flights of opportunity, because even astronomical flights
would not normally provide such facilities.
The astronomer will normally be heavily involved with the primaryexperiment or the larger telescopes that justify the flight andhe cannot devote much time to the smaller UV survey instruments.Also, after the first few flights all the problems of exposure orother uncertainties and surprises will be worked out, so auto-mated observations should be very reliable. It is possible thatinspection of a new survey photograph would show that a programshould be modified. It appears that there will be little require-ment for an astronomer to be on board as far as the small UV tele-scopes are concerned. This could change if electronic sensorsare used, or if a totally new instrument is being tried out thatrequires a great degree of attention.
IV-17 and IV-18
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The advantage of a flight of opportunity is that small instru-ments could be flown more frequently, giving more scientists op-portunities to try out ideas and instruments, and providingshorter lead time to useful data. However, many nonastronomicalflights will be Earth oriented rather than inertially stabilized,so that the instrument platform must have a continuous slew of4 degrees/minute. Even worse, the Shuttle bay will often beopen toward the Earth, so that a boom must be used to get theplatform out to the side where the stars can be seen. The Shuttleis stabilized to ±0.5 degrees with fairly high accelerations, andsuch a boom could be very difficult to use. For these reasonsit may be that unrestricted flight-of-opportunity capability willnot prove useful. The case for flights of opportunity on astro-nomical Sortie missions is much stronger. The Shuttle will beinertially stabilized, so that continuous slewing of the platformwill not be required. The survey instruments may have filmchanges and some performance monitoring, however, because theycould be designed for astronomical flights of opportunity, theoperation would be mostly automatic.
Table IV-2 is a summary of the advantages and disadvantages forthe dedicated mode and for the flight-of-opportunity mode.
TabZe IV-2 UV Instrument Mission Mode Implications
Mission ModeProgramConsiderations Flight of Opportunity Dedicated Mission
Astronaut Minimum Interface Modify Concepts andParticipation Using Existing Exper- Accommodation to Use
iment Concepts Crew Skills
Support Hardware Minimum Requirements AirlockRequirements Controls & Displays Resupply Film
Provide Adequate Film On-Board Filmfor 7 days Processing
Mission Inertially Stabilized Select Orbit Condi-Parameters Spacecraft tions for Desired
Inclination and Target AreasLaunch Window Con- Optimize Dark Timestraints
Operations Automatic after Interrupt Data TakingCheckout and De- to Change Filmployment Process and Interpret
Film
Scientific Satisfactory SatisfactoryObjectives
IV-23
c. Recommended Mission Mode - The flight-of-opportunity modewas selected, because the scientific objectives of the UV in-struments can be satisfied with minimum crew and hardware inter-faces, and the multiple missions necessary for complete surveyof the celestial sphere can be flown.
5. Platform Stabilization and Control Analysis
Possible types and characteristics of stabilized platforms re-quired for pointing and stabilization of the UV survey instru-
ments were analyzed during the study. The pointing and stabili-zation requirements of the instruments were compared with the
available performance data on the NASA/MSFC ST-100-SI Platform.Shortcomings of the ST-100 platform were discussed and necessary
improvements were identified. An outer gimbal configuration forthe platform was also considered.
For this analysis of stabilization and control, the items of par-
ticular interest concerning each instrument were:
1) Guiding (stabilization) accuracy requirement;
2) Pointing accuracy requirement;
3) Instrument field of view;
4) Instrument orientation requirements (location of sources ofinterest).
Other data values help to describe the instruments and to iden-tify physical compatibility with the stabilizing platform. Exam-ination of the data showed that some of the instruments are easierto accommodate than others. The instruments designed for rocket-flight interfaces generally do not require extreme stabilizationaccuracies. In contrast, instruments that were originally de-signed as balloon-borne payloads, are considerably more sophisti-cated and demanding.
a. ST-100 Platform - The ST-100-Stellar Instrument Platform wasoriginally designed as a stabilized mount for three small UV sur-vey instruments (boresighted on the platform) to be carried onan early Skylab flight. The mechanical design of the platformconsists of a conventional internal gimbal system. Each gimbal
axis includes two double-race ball-bearings, a brushless synchroresolver, and one or two brushless torquers (two torquers areused in the axis with high moment of inertia). The "inertial
IV-24
gimbal," or instrument mounting frame, must be specially tailoredto optimize the physical mounting arrangement of the instrumentsassigned to a specific flight. It is possible to accurately bal-ance the "inertial gimbal" with counterweights, so that the cen-ter of gravity of the assembly is located precisely at the pointof intersection of the three gimbal axes. This provides for gooddecoupling of the instrument mounting frame from motions of thespacecraft; residual coupling can be attributed primarily to bear-ing stiction and rolling friction, and to cable bending friction.A conventional mechanical-contact caging mechanism is included toprovide a rigid structural support during launch/reentry periods,and perhaps when high acceleration vehicle maneuvers arerequired.
To change the platform orientation, the synchro transmitters atthe control panel are rotated to selected gimbal angles, whichintroduces a bias in the control loops. The platform slews un-til the commanded orientation is reached when the resolvers atthe gimbals match the synchro transmitter settings. To hold thecommanded position, a triad of mutually orthogonal floated rateintegrating gyros is used, mounted (strapped down) on the "iner-tial gimbal." The output of each gyro is connected directlythrough an amplifier, shaping network, and a signal conversionmodule, to the corresponding brushless torquer. The nominal"crossover" point of the gimbal stabilization systems is esti-mated to be between 10 and 20 rad/sec- 1 (closed loop response ofabout 1.6 to 3 Hz).
One aspect of the ST-100-SI design, of particular interest tothe pointing control and stabilization area, involves the rategyros. For the original design, the gyro triad used Kearfott"King" rate gyros, with a drift/bias rating of 9.4 deg/hr(X2 prad/sec- ) if uncompensated. By careful bias compensation,the drift rate can be reduced to 0.03 to 0.04 deg/hr (X0.15 to0.20 prad/sec-1). These Kearfott units are small, dependable,high-output-level components, but they appear to be the majorsource of platform drift, limiting the guiding accuracy (stabil-ity) of the platform. The instrument payload that can be carriedwill depend on the moments of inertia of the "inertial gimbal"assembly, and on the torque delivered by the brushless torquers.The engineering model unit designed to carry the ST-100 instru-ments had the following characteristics:
1) I = 529 x 106 g cm2 = 39 slug ft2 ; T = 7.3 N m = 5.4 lb ft;x x
2) I = 390 x 106 g cm2 = 29 slug ft2 ; T = 4.9 N m = 3.6 lb ft;2.4 N m = 1.8 lb ft.
3) I = 84.5 x 106 g cm2 = 6.2 slug ft2; T = 2.4 N m = 1.8 lb ft.z z
IV-25
Freedom in the roll axis (outer gimbal, x) is ±1.74 radians =±100 degrees, and in the pitch (middle gimbal, y) and yaw (innergimbal, z) it is ±0.26 radian = ±15 degrees. The nominal designperformance characteristics of the ST-100-SI platform, as derivedfrom the available documentation, and from personal discussionswith NASA/MSFC personnel, are shown in Table IV-3.
To provide the platform characteristics required by the instru-ments (Table IV-1), the ST-100-SI platform requires some easilydefined improvements. The most obvious step in upgrading theplatform guiding error (stabilization accuracy) is to substitutethe original type of rate gyros for some of the currently avail-able units with improved drift rate performance specifications.Qualified units with nominal drift rate of 0.003 and 0.001 deg/hrcan be substituted for the original units in the same mountinglocation, with only a slight increase in unit cost.
For stabilization in the range of 2 to 10 arc seconds (10 to 50prad) in the presence of the anticipated torques generated bythe spacecraft motions and by the survey instruments during filmchange, it is generally necessary to resolve the gyro outputsthrough appropriate coordinate transformations. This would elim-inate the undesirable cross-coupling effects between the gimbalaxes that occur if the gimbals are not orthogonal. These coordi-nate transformation computations are easy to perform with simple,special-purpose mini-computer equipment. The frequency contentof the anticipated input disturbance torques extends to at least2 Hertz (12.5 rad/sec-1), depending partly on the type of stabil-ization used in the spacecraft. To overcome these input distur-bances to the required level of instrument stabilization, aclosed-loop response of at least 5 Hertz (31.4 rad/sec- 1) willbe necessary.
Torquer output capabilities may also require upgrading, particu-larly if more massive instruments, with awkward installation con-straints, are considered. To achieve the capability to launchand retrieve the platform in any selected orientation relativeto the spacecraft axes, it is necessary to use a later design ofthe platform that includes large bearings in the roll axis (outergimbal, axis of wide freedom).
IV-26
Table IV-3 ST-100-SI Platform Characteristics
SI English
Weight:
Platform without Instruments 115 kg 250 lb
Control and Display Panel 34 kg 75 lb
Platform with Instruments andInterface Equipment 290 kg 635 lb
Instrument Payload Capacity 300 kg 440 lb
Dimensions, without Instruments:
Height (x) 1.07 m 42 in.
Length (y) 1.27 m 50 in.
Width (z) 0.94 m 37 in.
Dimensions, with Instruments:
Height (x) 1.27 m 50 in.
Length (y) 1.83 m 72 in.
Width (z) 1.0 m 39 in.
Angular freedoms:
fx (roll) ±1.74 rad ±100 deg
6 (pitch) ±0.26 rad ± 15 deg
%z (yaw) ±0.26 rad ± 15 deg
Pointing Accuracy (3 axes) 9 mrad 0.5 deg
Stability (guiding error):
Short Term, rms 60 P rad 12 sec
Long Term (over 20 min) 250 p rad 50 sec
Stabilization System Closed- 10-20 rad sec- 1 1.6-3 Hzloop Bandwidth
Power Requirements:
Peak (stalled at full power) 790 W
Average (platform operating) 165 W
Average over one Full Orbit 120
Mounting Interface:
Flat flange with 8 holesuniformly spaced in acircular pattern
Flange Width 35 mm 1.375 in.
Flange Outer Diameter 406 mm 16 in.
Diameter of Hole Pattern Circle 381 mm 15 in.
Diameter of Mounting Screw Holes 9.5 mm 3/8 in.
IV-27
There are two reasons why it was considered desirable to haveShuttle spacecraft orientation data from the IMU available atthe platform operations control computer:
1) The crewman should be able to select exposure fields basedon right ascension and declination (celestial) coordinates;
2) If spacecraft angular rate signals are available from theIMU, it is possible to use these signals in the platformstabilization loops to further reduce the effects of Shuttlemotions during instrument data acquisition periods.
If these modifications and improvements are included, it appearsthat the ST-100-SI Platform can meet the pointing control andstabilization requirements for the small UV survey instruments.
Modifications and improvements that would significantly upgradethe ST-100 platform concept are listed in Table IV-4.
TabZe IV-4 ST-100 PZatform Modifications/Improvements
Assembly/Characteristic
Rate Gyros-Drift Bias
Closed-LoopResponse
Torquers
PointingAccuracy
Bearing
Performance
ST-100
0.4°/Hour(2x10- 6 rad/sec)
1.6 to 3.0 Hertz(10 to 20 rad/sec)
Platform GimbalAngles to±90 x 10- 4
radians(±30 m'n)
Modified/Improved
0.001°/Hour(0.5x10- 8 rad/sec)
5 Hertz(31.4 rad/sec)
Initial Orientationby Crewman.Pointing ChangesRelative to Ouptutof Improved RateGyros.
Comments
Major source of platform drift.Qualified units with lowerdrift rate in same mountinglocation.
Frequency content ofAnticipated input disturbancetorques extends to - 2 Hz(function of S/C stabilization).
Upgrade to support otherlarge instruments withawkward installationconstraints.
Input to platform operationscontrol computer to selectexposure fields based oncelestial coordinates.
Incorporate larger rollaxis bearings.
IV-28
As shown, stabilization accuracy can be upgraded by substitutingthe original rate gyros with currently available units with im-proved drift rate performance specifications. Qualified unitsare available that can be substituted in the same mounting loca-tions. The pointing control and stabilization loop should beextended to at least 5 Hertz to overcome the anticipated inputtorques. Initial pointing orientation will be performed by thecrewman, and changes to other targets will be derived from theoutput data of the improved rate gyros. The output capabilitiesof the brushless torquers will require upgrading to accommodatethe larger instruments. Larger roll axis bearings provide thecapability to deploy and retrieve the platform in any selectedorientation relative to the spacecraft axis.
b. Outer Gimbal Platform - The UV experiment groups identifiedearlier were examined to determine the smallest circular mount-ing platform capable of mounting any of the groups (Fig. IV-11).
46in.
nX 46 in.
Internal Grating 6-in. SurveySpectrometer Cameras (2)
Imaging Camera All-ReflectiveSpectrograph
Image-ConvertingSpectrograph
46 in.
ge-Converting Spectrograph Image-Converting Spectrograph
ce-mounted) (thru-mounted)
Figure IV-11 UV Outer Gimbal Platform Sizing Study
IV-29
Imag(fac
The size of the circle shown in the figure represents the insidediameter of the inner roll ring of a conventional three-axisouter gimbal system. All of the experiment groups fitted readilyon a 117-centimeter (46-in.) diameter platform, except for theST-100 group and the Image-Converting Spectrograph. By mountingone of the instruments on the far side of the platform and pro-viding a viewing hole in the platform, the ST-100 group could bemounted on the platform. The Image-Converting Spectrograph couldnot be face-mounted on the platform, but can be through-mountedas shown. The outer gimbal system is similar to the 213-centimeter(84-in.) system recommended in the earlier part of the study toaccommodate the intermediate class instrument. This system usesan azimuth table that provides complete 360-degree rotation, plusan elevation mechanism that gives a gimbal range of ±90 degrees.This azimuth-elevation system can achieve hemispherical coveragewithout requiring that the instruments and platform be deployedout of the cargo bay. The platform capabilities derived to sat-isfy the wide field-of-view UV survey instruments are listed inTable IV-5. Weights for the outer gimbal platform installation,with and without pallet, are given in Table IV-6.
Table IV-5 UV PZatform CapabiZities
Pointing Accuracy 5.8 x 10- 4 radian (2 m"n)
Stability 5 x 10-6 radian (1 siec)
Stabilization SystemClosed-Loop Bandwidth 31.4 rad/sec (5 Hz)
Gimbal Travel Hemispherical
Weight of Experiments 200 kg (440 lb)
c. Recommended PZatform - The outer gimbal mount configurationis recommended as the platform to accommodate the UV instruments.One of the factors considered was that several of the instrumentsanalyzed in this study may undergo significant changes duringdevelopment efforts. Since the ST-100 platform was configuredfor specific small experiments, its capabilities and versatilityare limited. Modifications of the ST-100 platform are necessaryto meet the requirements of the UV instrument groups that werestudied. In addition, the internal gimbal arrangement is diffi-cult to counterbalance when the experiment complement is unsym-metrical. Another important factor is that hemispherical view-ing capabilities of the outer gimbal concept enhance the operationof the telescopes on flights of opportunity. While higher weightand cost may be unavoidable, the advantages of the outer gimbalmount configuration outweigh the disadvantages.
IV-30
Table IV-6 UV Outer Gimbal PZatform Installation Weights
English,SI, kg lb
Gimbal System
Gimbal Rings 123.2 272Roll Drive 5.0 11AZ Stabilization Actuator 18.2 40EL and Pointing Stabilization Actuator 31.8 70AZ Yoke and Table 94.3 208AZ Drive 9.1 20Launch Locks 29.5 65Equipment Platform 4.5 10Support Equipment 31.8 70Eject Mechanism 23.1 51Miscellaneous and Cabling 10.4 23
381.0 840
With Pallet
Gimbal System 381 840Truss 33.2 71
414.2 911
Without Pallet
Gimbal System 381 840Truss 67.2 148
448.2 988
6. Support Requirements
a. Interfaces - The range of accommodation requirements for thesmall UV astronomy payloads is shown in Table IV-7 for the pro-posed flight of opportunity mission mode. These requirements in-clude the platform outer gimbal mount and truss supports, point-ing and control components, and the additional instrument supporthardware, except for the remote controls and displays (remotecontrols and displays add 44 Watts, 16.3 kilograms (36 lb), and0.2 kbps to the respective requirements).
IV-31
Table IV-7 UV Instrument Accommodation Requirements
Parameter Minimum Maximum
Size, Cargo Bay Length -Meters (in.) -2.54 (100) -2.54 (100)
Weight - kg (lb) 468 (1031) 633 (1396)
Data - kbps 1.0 21.0
Power - Watts 330 465
Pointing Accuracy - Radians
(min) 29 x 10 (10) 5.8.x 10- 4 (2)
Pointing Stability - Radians
(sec) 58 x 10- 6 (12) 5 x 10- 6 (1)
The maximum weight requirement reflects a UV instrument platformmounted with a drop truss structure in lieu of having a Shuttlepallet available. Maximum payload weight was developed usingthe ST-100 grouping payload weight of 176 kilograms (389 lb), thedrop truss structure weight of 450 kilograms (988 lb), plus ex-periment support equipment weight of 8.6 kilograms (19 lb). Maxi-mum power was derived by adding the ST-100 grouping power require-ment of 151 Watts to the experiment support equipment power re-quirement of 314 Watts. The maximum data rate of 20 kbps wasgenerated specifically from the proposed Scanning Spectrometer(Kondo) experiment. Pointing accuracy and stability must be pro-vided by the proposed outer gimbal mount and platform.
b. Hardware - The hardware recommended to support the small UVinstruments for flights of opportunity (Table IV-8) consists ofthe platform with gimbals, the mount, the gyro package, electronicsand experiment support equipment, and the payload controls anddisplays.
The platform consists of an outer gimbal system similar to thesystem recommended for the intermediate class instruments. Themount is capable of ±180-degree azimuth and 90-degree elevationtravel to provide hemispherical coverage independent of space-craft attitude. Rate gyros with drift rates of 0.5 x 10-8 radians/second (0.001 deg/hr) are incorporated to hold the initial com-manded position and to provide reference outputs for automaticprogramming to preselected targets. The control system providesa closed-loop response of at least 5 Hertz (31.4 rad/sec) toovercome disturbances from anticipated torques.
IV-32
Table IV-8 UV Instrument Support Hardware
Electronics and experiment support equipment mounted on the plat-form provides power, control, and status monitoring of the supporthardware and the UV instruments. The mini-computer is used toperform the coordinate transformation computations for platformpointing and to allow target programming based on celestial coor-dinates, not on platform gimbal angles as is the case in theST-100 platform designs. A closed-circuit TV system is providedfor crewman confirmation that the platform is on target beforebeginning an experiment exposure sequence. The reference cameratakes one exposure for each exposure cycle of the UV instrumentsto assist in identifying the field of view. A sun-Earth sensordetects the presence of the sun or bright Earth within 30 degreesof the instrument look angle. Local flux monitors are used toprogram exposure duration and to indicate amount and time of pos-sible data degradation. Functions and displays provided to thecrewmen include the platform and mount controls, experiment andsupport equipment controls, TV monitoring system, and status in-dicators.
IV-33
Platform (gimbals have ±2.5-deg stabilization travel)
Pointing Accuracy: 5.8 x 10- 4 radian (2 min)
Pointing Stability: 5.0 x 10-6 radian (1 sec)
Mount - Azimuth and Elevation
Travel: Hemispherical Coverage
Gyro Package
Drift: 0.5 x 10-8 rad/sec (0.001 deg/hr)
Electronics
Distribution Boxes (Power and Telemetry)
Programmer (Experiment Control and Sequencing)
Mini-Computer
Experiment Support
Vidicon Camera (Star Field Observation)
Reference Camera
Sun-Earth Sensor
Radiation Detector (SAA and Lyman-Alpha)
Controls and Displays
Platform and Gimbal Functions
Experiment Support
B. NONDEPLOYED SOLAR PAYLOAD
The objective of this task was to evaluate alternative methodsof satisfying the solar payload viewing requirements. The con-
cept recommended during the original 9-month Astronomy SortieMission Definition Study required that the entire solar payloadbe rotated through 90 degrees, so that the Sortie Lab's longitu-dinal axis was perpendicular to the Shuttle's longitudinal axis.The ground rule for the original study was that the deploymentmechanism was charged to the Shuttle orbiter. At the final per-
formance review, personnel from the Sortie Lab project at MSFCrequested that the solar payload concept be re-examined, sincethe latest Shuttle ground rule had the deployment mechanismcharged to the payloads. As a result of this task, the config-uration now recommended for the solar payload is a nondeployedconcept that does not require the deployment mechanism.
1. Ground Rules and Assumptions
The ground rules and assumptions that were used during the per-formance of this task are:
1) The deployment mechanism shall be charged to the payload -revised study ground rule;
2) The deployment mechanism shall be as defined by the GDC doc-
ument Shuttle/RAM Deployment Mechanism Conceptual Design(Ref 6). The deployment mechanism weight is 900 kilograms(2000 lb) and length is 1.52 meters (5 ft);
3) Deployed payloads shall not exceed 13.72 meters (45 ft) inlength - extracted from Sortie Lab Design Requirements (Ref
7);
4) A deployable pallet 9.75 meters (32 ft) in length weighs 544kilograms (1200 lb) more than a nondeployable pallet - ex-tracted from Sortie Lab Briefing to Solar Physics WorkingGroup (Ref 8).
IV-34
2. Alternative Configurations
Four solar payload configurations were evaluated. They consistedof the baseline deployed configuration defined during the origi-nal 9-month study, a modified baseline, and two nondeployed con-figurations. Three different inertial attitudes also consideredin the analyses were: X-POP (Shuttle's longitudinal axis perpen-dicular to the orbit plane); X-IOP (Shuttle's longitudinal axisin the orbit plane); and Z-POP (Shuttle's longitudinal axis inthe orbit plane and the Z axis perpendicular to the orbit plane).
The four configurations were evaluated in terms of total weightand length, compatibility with stellar payloads, momentum require-ments for the various inertial attitudes, compatibility with therevised ground rules, and viewing capabilities of the solar in-struments.
a. Baseline Configuration - The baseline configuration recom-mended at the completion of the original 9-month study is shownin Figure IV-12. The significant characteristics of this con-figuration are: the entire payload must be deployed from theShuttle cargo bay to enable the solar instruments to view thesun; the Shuttle maintains an X-POP inertial attitude throughoutthe 7-day mission; the deployment mechanism is charged to the
Shuttle orbiter; the orbit inclination is greater than 67 degreesto enable continuous sun viewing during the 7-day mission; andthe ancillary hardware is common with the stellar payloads. Adetailed description of the baseline configuration is presentedin Volume III, Book 1 of the original 9-month study Final Report(Ref 1).
The weight statement for this configuration is summarized in TableIV-9. The total payload weight of 13,067.5 kilograms (28,809 lb)is based on the use of three ATM-type control moment gyros (CMG)to stabilize the entire Shuttle in an X-POP inertial attitude.The weight also includes' the use of a 4.73-meter (15.5-ft) SortieLab that weighs 5755.1 kilograms (12,688 lb) and a standard palletthat weighs 1388 kilograms (3060 lb). The total payload lengthis 17.83 meters (58.5 ft), exclusive of the deployment mechanism.
IV-35 and IV-36
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Table IV-9 Baseline Configuration Weights [kg (Zb)]
Sortie Lab - 4.73 meters(15.5 ft)
Pallet - 13.2 meters (43 ft)
Stabilization SystemCMG (3)CMG Inverter (3)IMUControl and Input BoxSupports and Cabling
Mount and Gimbal SystemsForward Common MountAft Common MountOrdnanceForward Gimbal SystemAft Gimbal System
Electrical and Data System
Thermal Insulation
Solar InstrumentsPhotoheliographSolar Group
Total Payload
571.577.66.89.1
48.5
481.7481.79.1
585.6585.6
(1260)( 171)( 15)( 20)( 107)
(1062)(1062)( 20)(1291)(1291)
997.9 (2200)1967.2 (4337)
5,755.1 (12,688)
1,388.0 ( 3,060)
713.5 ( 1,573)
2,143.7 ( 4,726)
43.1 (
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The primary advantages of this configuration, and the reasons it
was selected as the baseline for the original 9-month study are:
1) The Shuttle could be maintained in an X-POP inertial atti-
tude, which is the best inertial attitude in terms of momen-tum requirements;
2) The ancillary hardware (i.e., mounts, stabilization system,
etc) is common with the stellar payloads;
3) The deployment of the entire payload minimizes any viewingconstraints on the instruments due to reflecting surfaces in
the field of view of the instruments:
4) The configuration was compatible with the study ground rulethat charged the deployment mechanism weight and volume to
the Shuttle orbiter.
IV-39
b. Modified BaseZine Configuration - One alternative that wasevaluated for possible use was a modified baseline configuration.This configuration (Fig. IV-13) was developed in an effort to becompatible with the modified study ground rules. To provide vol-ume in the Shuttle cargo bay for the hinged deployment mechanismit was necessary to shorten the overall length of the payload by1.52 meters (5 ft). This was accomplished by relocating the for-ward mount closer to the aft end of the Sortie Lab. This modifi-cation results in a limited capability azimuth table that is ac-ceptable for the solar payloads, but not for the stellar. Inaddition, it was also necessary to reverse the positions of thephotoheliograph and solar group telescopes in order to reducethe overall length.
The totel payload weight for this configuration is 13,702.9kilograms (30,209 lb) (Table IV-10).
TabZe IV-10 Modified BaselZine Configuration Weights [kg (Zb)]
Sortie Lab - 4.73 meters(15.5 ft) 5,755.1 (12,688)
Pallet - 12.05 meters (39.5 ft) 1,115.8 ( 2,460)
Deployment Mechanism 907.2 ( 2,000)
Stabilization System 713.5 ( 1,573)CMG (3) 571.5 (1260)CMG Inverter (3) 77.6 ( 171)IMU, 6.8 ( 15)Control and Input Box 9.1 ( 20)Supports and Cabling 48.5 ( 107)
Mount and Gimbal Systems 2,143.7 ( 4,726)Forward Common Mount 481.7 (1062)Aft Common Mount 481.7 (1062)Ordnance 9.1 ( 20)Forward Gimbal System 585.6 (1291)Aft Gimbal System 585.6 (1291)
Electrical and Data System 43.1 ( 95)
Thermal Insulation 59.0 ( 130)
Solar Instruments 2,965.1 ( 6,537)Photoheliograph 997.9 (2200)Solar Group 1967.2 (4337)
Total Payload 13,702.9 (30,209)
IV-40
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Lab weight of 5755 kilograms (12,688 lb), a pallet weight of1115.8 kilograms (2460 lb) and a deployment mechanism weight of
907.2 kilograms (2000 lb). The total payload length is 16.77meters (55 ft), exclusive of the deployment mechanism, which re-
quires an additional 1.52 meters (5 ft).
Although it was possible to modify the baseline configuration toallow an additional 1.52 meters (5 ft) for the deployment mechan-
ism, it was not possible to shorten the payload to the 13.72-meter (45-ft) length that is required for deployable payloads.
Therefore, this configuration was dropped from any future consid-eration since it did violate the new study ground rules.
c. NondepZoyed X-POP Configuration - As an alternative to deploy-
ing the entire payload out of the cargo bay, a canted configura-tion was investigated. This canted configuration (Fig. IV-14)is significant in that: payload deployment is not required;the Shuttle can maintain an X-POP inertial attitude; the solar
instruments are deployed from the cargo bay at offset angles thatallow both sets of instruments to view the sun; and the ancillaryhardware is compatible with the stellar payloads.
The characteristics of this configuration are compatible with themodified study ground rules, since the entire payload does not
require deployment. However, the major disadvantage of this con-figuration is the need for the solar instruments to view across
the top of the Shuttle cabin. Although the instrument fields ofview clear the Shuttle, reflections from the Shuttle surfaces
could still be detrimental. This is especially true of theCoronagraphs, which will view as far out as 30 solar radii. At
this distance the source signal will be relatively weak and re-flections off the Shuttle could be a significant problem. Special
coating or baffles on the Shuttle or solar instruments may ade-quately suppress the reflections, but could be impractical or
costly to implement.
The detailed weight statement for this configuration is shownin Table IV-11. The total payload weight is 12,227 kilograms(26,956 lb) and includes: the 4.73-meter (15.5-ft) Sortie Labthat weighs 5755.1 kilograms (12,688 lb); a lightweight pallet
that is nondeployable and is 13.6 meters (44.5 ft) in length andweighs 502.1 kilograms (1107 lb); and three ATM type CMGs that
will stabilize the entire Shuttle in an X-POP inertial attitudethroughout the 7-day sortie mission.
IV-43 and IV-44
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Table IV-11 NondepZoyed X-POP Configuration Weights [kg (Zb)]
Sortie Lab - 4.73 meters(15.5 ft) 5,755.1 (12,688)
Pallet - 13.6 meters(44.5 ft) 502.1 ( 1,107)
Stabilization System 713.5 ( 1,573)CMG (3) 571.5 (1260)CMG Inverter (3) 77.6 ( 171)IMU 6.8 ( 15)Control and Input Box 9.1 ( 20)Supports and Cabling 48.5 ( 107)
Mount and Gimbal Systems 2,189.1 ( 4,826)Forward Common Mount 509.4 (1112)Aft Common Mount 509.4 (1112)Ordnance 9.1 ( 20)Forward Gimbal System 585.6 (1291)Aft Gimbal System 585.6 (1291)
Electrical and Data System 43.1 ( 95)
Thermal Insulation 59.0 ( 130)
Solar Instruments 2,965.1 ( 6,537)Photoheliograph 997.0 (2200)Solar Group 1967.2 (4337)
Total Payload 12,227.0 (26,956)
d. NondepZoyed X-IOP or Z-POP Configuration - The final conceptinvestigated was the nondeployed X-IOP or Z-POP configuration.This configuration (Fig. IV-15) is essentially the same as thebaseline configuration but without the payload deployment require-ment. Because the payload is not deployed, it is not possible tomaintain an X-POP inertial attitude with the Shuttle, and conse-quently two other inertial attitudes were considered, they werethe X-IOP (Shuttle's longitudinal axis in the orbit plane) andthe Z-POP (Shuttle's longitudinal axis in the orbit plane and itsZ axis perpendicular to the orbit plane). Each of these inertialattitudes require the addition of one ATM type CMG to counteractthe gravity gradient torques.
IV-47 and IV-48
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The detailed weight statement forin Table IV-12.
Table IV-12 Nondeployed X-IOP or[kg (Zb)]
this configuration is presented
Z-POP Configuration Weights
Sortie Lab - 4.73 meters(15.5 ft)
Pallet - 13.2 meters(43.4 ft)
Stabilization SystemCMGs (4)CMG Inverter (4)IMUControl and Input BoxSupports and Cabling
Mount and Gimbal SystemsForward Common MountAft Common MountOrdnance PackageForward Gimbal SystemAft Gimbal System
Electrical and Data System
Thermal Insulation
Solar InstrumentsPhotoheliographSolar Group
Total Payload
762.0103.4
6.810.063.9
481.7481.79.1
585.6585.6
(1680)( 228)( 15)( 22)( 141)
(1062)(1062)( 20)(1291)(1291)
997.9 (2200)1967.2 (4337)
5,755.1 (12,688)
502.1
946.2
( 1,107)
( 2,086)
2,143.7 ( 4,726)
43.1
59.0
2,965.1
( 95)
( 130)
( 6,537)
12,414.3 (27,369)
The total payload weight of 12,414.3 kilograms (27,369 lb) in-cludes: the 4.73-meter (15.5-ft) Sortie Lab weight of 5755.1kilograms (12,688 lb); a lightweight nondeployable pallet 13.2meters (43.4 ft) in length and weighing 502.1 kilograms (1107 lb),and a stabilization system that has four CMGs.
This was the configuration that was selected as the preferred
solar payload concept as a result of analyses performed duringthis task.
IV-51
3. Inertial Attitudes
Because the inertial attitude of the Shuttle orbiter has a majorimpact on the configuration of the solar payload, three differ-ent inertial attitudes were investigated--X-POP, X-IOP, and Z-POP.Each of these inertial attitudes places different requirements onthe number of CMGs that are necessary to stabilize the Shuttle,and on the deployment and gimbal mechanisms that are required forsolar viewing.
The four configuration alternatives considered the differencesin the deployment and gimbal mechanisms that would be necessaryto view the sun for the indicated inertial attitudes. This sec-tion of the report summarizes the analyses performed to determinethe requirements and tradeoffs for pointing and control of theShuttle orbiter. Calculations were made to determine the momen-tum resource requirements, electrical power requirements, numbersand weights of the CMG system, and size and weights of alternativereaction control systems (RCS). Although the original 9-monthstudy did recommend the use of CMGs to stabilize the Shuttle or-biter, there is still sufficient interest in the use of an RCS toperform this function that it was desirable to show what the re-quirements would be for an RCS.
a. Assumptions - The Shuttle orbiter model used in this reportassumed the following moments of inertia and products of inertia:
I = 1.11 x 106 kg-m2 (0.821 x 106 slug-ft2)x
I = 9.66 x 106 kg-m2 (7.130 x 106 slug-ft2 )Y
I = 10.00 x 106 kg-m2 (7.376 x 106 slug-ft2 )z
I = I = I = 0.xy xz yz
The Shuttle orbiter was assumed to be stabilized in a 270-nautical-mile circular orbit.
Sizing of the CMG system was performed according to the followingground rules. The momentum requirements for stability were com-puted for the contingency that gravity-gradient bias and cyclictorques may result from an axis misalignment of as much as 1 de-gree. For sizing, it was further assumed that momentum accumula-tions would be desaturated once every orbit rather than twice perorbit, in order to obviate potential interference with scientific
IV-52
observations. The momentum requirements for maneuvering were com-puted to obtain at least 2 degrees/minute about the axis of great-est inertia, giving much greater maneuver capability about theX-axis. The transition from attitude control propulsion system(ACPS) control to CMG control requires sufficient CMG momentumcapability to absorb the maximum angular excursions possible whenthe ACPS control is terminated. Here, the assumption is madethat immediately after transition to CMG control, the accumulatedCMG momentum is removed by pulsing the ACPS in proper direction.
The CMG system was sized to handle the greater of these three re-quirements; one additional CMG was added for contingencies suchas failure or the situation where a maximum effort maneuver isdesired when accumulated momentum is near maximum. An importantfeature is that even with one CMG inoperative, sufficient capa-bility will exist in each case to complete the scientific pro-gram with perhaps somewhat reduced maneuvering capability.
The momentum requirements computed for RCS control apply for a7-day mission, and were computed on the basis of the same groundrules for stability, transition, and maneuvering as those usedfor CMGs. However, there are some differences in the way theymust be applied. For the RCS, the requirements for stability,transition, and maneuvering are additive, as the momentum ex-pended cannot be recovered. No additional fuel for other con-tingencies is carried, as the worst effect of running out of RCSfuel would be that the observation program would have to be ter-minated early. Because the location of the RCS thrusters on theShuttle orbiter are not known, calculations of impulse were madeusing an average moment-arm of 35 feet for torquing the vehicleabout all axes.
The momentum required for transition from the orbiter's ACPS tothe CMG or RCS is based on the following orbiter rates:
x = 3.470 mrad/sec (0.200 deg/sec);
X = 0.400 mrad/sec (0.023 deg/sec);y
w = 0.387 mrad/sec (0.022 deg/sec).z
IV-53
b. Control Requirements for X-IOP Inertial Attitude - This sec-tion discusses the control requirements for a Shuttle orbiterinertially stabilized in an X-IOP orientation, using a CMG sys-tem or an RCS.
1) CMG Momentum Requirements - The Shuttle orbiter externaltorque environment is assumed to be primarily gravity-gradientinduced torques with aerodynamic induced torques contributingonly slightly. The proposed CMG system must be capable of maneu-vering and stabilizing the vehicle. Considering the stabiliza-tion aspects, the CMG system must be capable of storing the total
resultant gravity-gradient induced angular momentum, HCMG )
which is composed of an accumulated momentum component, Ha, due
to constant axial torques, and a cyclic momentum component, Hc,
which varies periodically, depending specifically on twice theShuttle's orbital frequency, 2o . The CMG gravity-gradient mo-
mentum storage requirement is maximized by using the peak cyclicmomentum, H peak, and is given by
(STAB) - + *CMG = Ha +
p
These values were computed using the following assumptions:
1) The,Shuttle orbiter is in the X-IOP attitude and is initiallyoriented with its X and Y axes in the orbital plane;
2) The Shuttle orbiter is inertially held in its desired orien-tation throughout a complete orbit.
The magnitude of the accumulated momentum, 13 , for a complete
orbit, and the magnitude of the peak cyclic momentum componentsare found to be:
lHa |= 1740 N-m-sec (1280 ft-lb-sec)x
IHa| = Oa
HI = 0
IH I =. O
IV-54
iH I = 5050 N-m-sec (3730 ft-lb-sec)cy
IHcz| = 5100 N-m-sec (3760 ft-lb-sec).p
The total peak cyclic momentum is determined by
|H | = F | 2 + CZI 21c Hcylp
|HJ| = 7180 N-m-sec (5300 ft-lb-sec).I cp
The total CMG gravity-gradient momentum storage requirement forstabilization is therefore
(STAB) = - +HCMG I
= 8920 N-m-sec (6580 ft-lb-sec).
(STAB)These magnitudes are computed for the case where H G has itsCMG
maximum value obtained when the angle X, which subtends the or-biter's Z-axis and its projection onto the orbital plane, has avalue of X = 44.8 degrees.
The angle, X, can vary between 0 and 90 degrees while the vehicleremains in the X-IOP attitude. This will occur during stellarobservations in order to point the instrument at various targetsand will occur during solar observations because of variationsin the Beta angle and regression of the orbit. Because of thesepotential variations for the angle, A, two special cases havebeen computed--one for the X = 90 degrees (Z-POP, X-IOP) attitudeand one for the X = 0 degree, resulting in a Y-POP, X-IOP atti-tude. These represent the extreme cases away from the maximummomentum requirements.
First, consider the (Z-POP, X-IOP) attitude where X = 90 degrees.In this case, both the X and Y axes are in the orbital plane, anideal condition that results in zero accumulated momentum. Inaddition, the peak cyclic X and Y momentum components H , and
cx
H are also zero. The only term that contributes to H(STAB) iscy CMG
H = 7050 N-m-sec (5200 ft-lb-sec).cz
Also, consider the (Y-POP, X-IOP) attitude for this orbit. Inthis case, both the X and Z axes are in the orbital plane, andA = 0 degrees, a condition that also results in zero accumulated
IV-55
momentum. For this attitude, the peak cyclic X and Z momentumcomponents H and H are zero. The only term contributing to
H(STAB) is H = 7340 N-m-sec (5420 ft-lb-sec).CMG cy
These results (Table IV-13) show that the accumulated momentum
H , is zero for both the (Z-POP, X-IOP) attitude and the (Y-POP,
X-IOP) attitude. The (Z-POP, X-IOP) attitude also shows a lower
total peak cyclic momentum requirement than the (Y-POP, X-IOP)attitude (by approx 4%).
As previously stated, the X-IOP attitude, which exhibits the
greatest total stabilization momentum requirements, occurs forA = 44.8 degrees. This attitude also develops the largest accu-
mulated momentum per orbit and, therefore, requires desaturation.Table IV-13 contains two values for each component of the X-IOP,
X = 44.8-degree case. The smaller value is for an experimentalperiod of approximately 66% of an orbit, while the larger value
is for the total orbital period.
The table also shows that three ATM-type CMGs are sufficient tomeet X-IOP stabilization requirements, in the absence of a maneu-
vering requirement.
In addition to stabilization requirements, the angular momentumcapability for maneuvering must also be available at any time.
To compute the CMG maneuvering capability requirements, the fol-lowing assumptions are made:
1) Pointing the telescope to various targets is accomplished by
two distinct Shuttle orbiter maneuvers--first, a maneuverabout its Z-axis, and then one about its X-axis;
2) These maneuvers are performed at a maneuver rate of 2 degreesper minute.
Calculations show that for a maneuver rate of 2 degrees/minute,the X-axis and Z-axis angular momentums that must be imparted
to the orbiter by the CMGs are:
H( m) = 646 N-m-sec (477 ft-lb-sec);x
H(m) = 5820 N-m-sec (4300 ft-lb-sec).z
The proposed orbiter CMG system must be sized so that its momentum
maneuver capability H1 m) equals H the largest of the above
momentums. Therefore, H(m = 5820 N-m-sec (4300 ft-lb-sec) if theCMG
maneuver rate is to be 2 degrees/minute. This momentum capacitycould be provided by two Skylab ATM-type CMGs.
IV-56
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IV-57
A momentum absorption capability described as H(transition) is
also required. This is because of the fact that when the CMGsystem takes over control from the orbiter's ACPS, the residualmomentum left by the baseline system must be absorbed. The mag-nitude of this transition momentum is
H(trans) = 3320 N-m-sec (2450 ft-lb-sec).CMG
This event, however, occurs only once per mission and is a negli-gible factor in sizing the CMG system.
Aerodynamics torques also act on the orbiter and induce an angularmomentum that must be stored by the CMG system. It is assumedthat the aerodynamic induced momentum is
H(aero) = 130 N-m-sec (95 ft-lb-sec).CMG
Table IV-14 is a summary of the CMG momentum requirements.
TabZe IV-14 CMG Momentum Requirements, X-IOP
The momentum capacity calculated for stabilization alone is 8920N-m-sec (6580 ft-lb-sec) and includes the maximum expected accu-mulated momentum per orbit and the peak cyclic gravity gradientcomponent. Adding the aerodynamics-induced requirement to thisyields a subtotal of 9050 N-m-sec (6675 ft-lb-sec).
IV-58
H i = 1740 N-m-sec (1280 ft-lb-sec)
|H |= 7180 N-m-sec (5300 ft-lb-sec)
(aero
H( = 130 N-m-sec (95 ft-lb-sec)
(STAB+aero) = 9050 N-m-sec (6675 ft-lb-sec)CMG
H(man) = 5820 N-m-sec (4300 ft-lb-sec)CMG
(trans)H1
tCnG = 3320 N-m-sec (2450 ft-lb-sec)CMG
The momentum capacity calculated for maneuvering at 2 degrees/
minute is 5820 N-m-sec (4300 ft-lb-sec).
The momentum capacity calculated for the transition from ACPS toCMG control is 3320 N-m-sec (2450 ft-lb-sec).
In sizing a CMG system, it is only necessary to size to the larg-
est of the above momentum requirements, since the three momentumswill not occur simultaneously. For the X-IOP inertial attitude
the largest momentum requirement is caused by the stabilizationand aerodynamic momentum of 9050 N-m-sec (6675 ft-lb-sec).
The angular momentum that accumulates on the CMGs during each
orbital period causes a bias to build up. This bias, if notdumped, will eventually cause the CMGs to become saturated and
ineffective. The amount of angular momentum required to be
dumped per orbit, while maintaining the X-IOP configuration, is
IHal = 1740 N-m-sec/orbit (1280 ft-lb-sec/orbit).
2) RCS Momentum Requirements - The low-thrust RCS must be capa-
ble of delivering sufficient momentum to the orbiter to provide:
1) Stability to the vehicle, i.e., reaction against gravity
gradient torques [this term is designated HTAB)];
2) Reaction to aerodynamic torques [designated by HRCS er
3) Attitude maneuvering of the vehicle [designated by H(m a n )]
4) Residual momentum absorption left by the baseline ACPS when
the low-thrust RCS takes over control [designated by HRCans)].
The rectified gravity-gradient momentums H , accumulated during
one orbit are calculated on a per axis basis to be:
H = 1745 N-m-sec/orbit (1285 ft-lb-sec/orbit);grx
H = 40,600 N-m-sec/orbit (30,000 ft-lb-sec/orbit);gry
H = 40,600 N-m-sec/orbit (30,000 ft-lb-sec/orbit).grz
IV-59
The total rectified gravity-gradient momentum that the X-IOPShuttle orbiter stabilization RCS must absorb is
H = 83,000 N-m-sec/orbit (61,285 ft-lb-sec/orbit).gr
Aerodynamic momentum compensation requirements are assumed to be5% of H , as calculated for the X-POP attitude
gr
H(aero) | = 30,600 N-m-sec/mission (22,600 ft-lb-sec/mission)
TotalMission
The RCS momentum requirements for attitude maneuvers are calcu-lated on a per axis basis. The maneuver rate was taken as 2degrees/minute about each axis. To maintain the X-IOP attitudeconstraint, the momentum H must remain zero. Thus
Y
H(m) = 2(H + Hz)RCS
H(m) = 12,950 N-m-sec/maneuver (9554 ft-lb-sec/maneuver).RCS
Half of H (m is used to put the orbiter in motion; the other halfRCS
is required to stop it after it has passed through the desiredrotation angle.
In addition to the above momentum, it will also be necessary to(LC)
provide limit cycle momentum, HRCS ,for those periods of the or-
bit where the gravity gradient and aerodynamic torques are notlarge enough to prevent the RCS from operating in its deadbandlimits. The limit cycle momentum was estimated to be 32,200N-m-sec (23,800 ft-lb-sec) for the 7-day mission.
A summary of RCS angular momentum requirements is presented inTable IV-15. These values are then used in determining the RCSimpulse requirements. The total RCS impulse required for the7-day mission is summarized in Table IV-16. This impulse isbased on an average moment arm of 35 feet for each axis and as-sumes a total of 106 orbits and 28 maneuvers during the 7-daymission.
IV-60
RCS Momentum Requirements, X-IOP
H ( m )RCS
H(STAB)RCS
H (aero)RCS
H(trans)RCS
H(LC)RCS
(Total)RCS
TotalMission
= 362,000
= 8,800,000
= 30,600
= 5,780
= 32,200
= 9,230,580
N-m-sec
N-m-sec
N-m-sec
N-m-sec
N-m-sec
( 267,512
(6,496,210
( 22,600
( 4,275
( 23,800
N-m-sec (6,814,397
ft-lb-sec)
ft-lb-sec)
ft-lb-sec)
ft-lb-sec)
ft-lb-sec)
ft-lb-sec)
Note: Total mission requirements are based on 106 orbitsand 28 maneuvers per 7-day mission.
TabZe IV-16
EFAt I
=gg
ZFAtl =laero
FAt ILC
ZFAtman =
ZFAtltran
=
EFAtlos =
TotalRCS I
sp
RCS ImpuZse Requirements, X-IOP
822,000
2,900
3,030
102,500
542
0
N-sec
N-sec
N-sec
N-sec
N-sec
(185,000
( 650
( 680
( 23,000
( 122
( 0
930,972 N-sec/(209,452mission
lb-sec)
lb-sec)
lb-sec)
lb-sec)
lb-sec)
)
lb-sec/mission)
Note: Considers 106 orbits/mission and 28maneuvers/mission.
IV-61
TabZe IV-15
c. Control Requirements for Z-POP Inertial Attitude - This sec-tion discusses the momentum requirements for the Shuttle orbiterin a Z-POP attitude. The analysis assumes a misalignment, froma true Z-POP attitude, of 1 degree about the X and Y axes.
1) CMG Momentum Requirements - The stabilization momentum forthe Z-POP inertial attitude will consist of the accumulated mo-mentum and the peak cyclic momentum on a per-axis basis.
The accumulated momentum magnitudes are:
X-Axis--H = 60.3 N-m-sec/orbit (44.5 ft-lb-sec);ax
Y-Axis--H = 1600 N-m-sec/orbit (1180 ft-lb-sec);ay
Z-Axis--H = 0.az
The total accumulated momentum magnitude is given by
H | = {H2 + H2 + H2 iiHa ax ay az
= 1600 N-m-sec/orbit (1180 ft-lb-sec/orbit)
The cyclic momentum magnitudes are:
X-Axis--H = 3.55 (cos 2w t - sin 2wt);
Y-Axis--H = 94 (cos 2w t + sin 2wt);
Z-Axis--H = 5200 (cos 2w t)
The peak cyclic momentum magnitude occurs at w t = 27--o
IHcxl = 4.82 N-m-sec (3.55 ft-lb-sec)
p
Hcyl = 127.0 N-m-sec (94.0 ft-lb-sec)
pIHcz
j= 7050 N-m-sec (5200 ft-lb-sec).
P
IV-62
The total peak cyclic momentum magnitude is given by
H |a = (H + Hy + Hz)cipeak cx CY
= 7050 N-m-sec (5200 ft-lb-sec).
The CMG momentum storage requirement for stabilization in the
Z-POP attitude, allowing 1 degree errors in rotation about the
X and Y axes, becomes
(STAB) =-HCMG = + Hc
= 8650 N-m-sec/orbit (6380 ft-lb-sec).
Add 5% to this value for aerodynamic torque compensation, so that
H(aero) = 130 N-m-sec/orbit (95 ft-lb-sec/orbit),CMG
and obtain
H(STAB+aero) = 8780 N-m-sec/orbit (6475 ft-lb-sec/orbit).CMG
Total
This value can be satisfied with three ATM-type CMGs that would
provide 6900 ft-lb-sec; therefore, stabilization in the Z-POP
attitude for a complete orbit without desaturation can be obtained.
2) RCS Momentum Requirements - The accumulated rectified angular
momentums acting on the three vehicle axes are
X-Axis--H = 68.5 N-m-sec/orbit (50.5 ft-lb-sec);grx
Y-Axis--H = 1830 N-m-sec/orbit (1350 ft-lb-sec);gry
Z-Axis--H = 56,500 N-m-sec/orbit (41,600 ft-lb-sec).grz
IV-63
The total rectified gravity-gradient momentum that the Z-POPShuttle orbiter RCS must absorb during each orbit equals
H =H + Hgr grx gry grz
= 58,400 N-m-sec/orbit (43,000 ft-lb-sec/orbit).
For a 7-day mission containing 106 orbits
H(STAB)gr
= 6,200,000 N-m-sec (4,558,000 ft-lb-sec).
TotalMission
The additional Z-POP momentum requirements for maneuvering, limitcycle, aerodynamic drag compensation, and transition from Shuttlebaseline to RCS are similar to those calculated for the X-IOPattitude.
The angular momentum requirements for maneuvering depend on themoment of inertia about the Z-axis, I , and the maneuver rate,
X . If the Z-POP attitude is to be maintained during a maneuver,
then H and H = 0. The values calculated for X-IOP, however,x y
include an H component. Since the magnitude of H is only 10%
of Hz no appreciable error is included by using H(m) = 9554 ft-z RCS
lb-sec the X-IOP value of HCS
(i
The total Z-POP RCS momentum may then be obtained as follows:
H(m)RCS
TotalMission
H(LC)RCS
TotalMission
(aero)HRCS
Total
= 362,000 N-m-sec (267,512 ft-lb-sec);
= 32,200 N-m-sec (23,800 ft-lb-sec);
= 30,600 N-m-sec (22,600 ft-lb-sec);
1Mission
IV-64
(STAB)HR(S T) = 6,200,000 N-m-sec (4,558,000 ft-lb-sec);RCS
TotalMission
(trans) 4= 5800 N-m-sec (4275 ft-lb-sec).RCS
TotalMission
Therefore,
H(total)RCS(Z-POP)
= 6,600,000 N-m-sec (4,876,400 ft-lb-sec)(for a maneuvering rate of 2 deg/min).
TotalMission
d. Control Requirements for X-POP Inertial Attitude - This sec-tion discusses the momentum requirements for the Shuttle orbiterin the X-POP attitude. The analysis assumes a misalignment fromthe true X-POP attitude by two small Y and Z axis rotational er-rors of 1 degree.
1) CMG Momentum Requirements - The analysisdetermines the magnitudes of the accumulatedpeak cyclic momentum on a per axis basis.
for this sectionmomentum and the
The Shuttle orbiter external torque environment for this attitudeand an altitude of 270 nautical miles is again assumed to be pri-marily gravity-gradient induced torques with aerodynamic-inducedtorques contributing only slightly, and being of the same magni-tude as those calculated for the X-IOP attitude.
The values computed are based on the following assumptions:
1) The desired Shuttle orbiter attitude is the X-POP;
2) The actual attitude is misaligned from true X-POP by a 1degree rotation about both the Y and Z axes;
3) Noncyclic momentum is accumulated for one complete orbit,while the vehicle is inertially held in its desired orienta-tion.
IV-65
The magnitude of the accumulated momentum IHa I, for a complete
orbit, and the magnitude of the peak cyclic momentum componentsare:
| Ha = ( H2
+ H2+ H2 )a \ax ay az_
= 3rw S[(IZ X) X) ]
= 2240 N-m-sec/orbit (1655 ft-lb-sec/orbit);
IHcx = 282 N-m-sec (208 ft-lb-sec);
p
Hcy = 127 N-m-sec (94 ft-lb-sec);p
I HC = 171 N-m-sec (126 ft-lb-sec).
p
On a conservative basis, the total peak cyclic momentum can becomputed as
H = 1 + +lc yCl (Hcx2 + Hcyl 2 cz) I
= 352 N-m-sec (260 ft-lb-sec).
The CMG gravity-gradient momentum storage requirement, H( S T A B)
CMG,for the X-POP attitude with 1-degree errors about the Y and Zaxes is
(STAB) 4- 4-
CMG a + p
= 2600 N-m-sec (1915 ft-lb-sec).
Aerodynamic torques also act on the orbiter and induce an angularmomentum that must be stored by the CMG system. It is assumedthat the aerodynamic momentum storage requirement is the same asthat calculated for the X-IOP attitude. This value is
H(aero) = 130 N-m-sec (95 ft-lb-sec).CMG
IV-66
Therefore,
H(STAB) + H(aero) = 2720 N-m-sec (2010 ft-lb-sec)CMG CMG
2) RCS Momentum Requirements - The Shuttle orbiter is assumedto be stabilized in the X-POP attitude. The orbiter is also as-sumed to be misaligned from its true X-POP attitude by two smallY and Z axis rotational errors, E and £ , each of which was setequal to 1 degree. Y Z
The analysis for this section determines the magnitudes of therectified angular momentum axial components. Results indicate:
H = 2200 N-m-sec/orbit (1620 ft-lb-sec/orbit);grx
H = 1830 N-m-sec/orbit (1350 ft-lb-sec/orbit);gry
H = 1760 N-m-sec/orbit (1295 ft-lb-sec/orbit).grz
The total required RCS momentum per orbit for stabilization tocounteract the gravity-gradient torques is
(STAB)HRCS gg = 5800 N-m-sec/orbit (4265 ft-lb-sec/orbit).
For a complete 7-day mission consisting of 106 orbits, the totalRCS momentum requirement to counteract the gravity gradient torquesis
(STAB)HRCS gg = 611,000 N-m-sec (452,090 ft-lb-sec).
TotalMission
To determine the maneuvering momentum requirements the maneuverrate was taken as 2 degrees/minute as a maximum. Furthermore,the X-POP attitude was assumed to be maintained. Thus, only Hwill be allowed to be nonzero. Therefore,
H(m) 1290 N-m-sec (954 ft-lb-sec/maneuver).RCS Maneuver
IV-67
Considering there are to be 28 maneuvers per 7-day mission yields
(m) g
HRCSTotalMission
= 36,200 N-mi-sec= 36,200 mission (26,712 ft-lb-sec/mission).mission
The additional X-POP momentum requirements for aerodynamic drag,
limit cycle, and transition are as follows: H(aero) is 5% of
H(STA B ) , and H(LC) is 5% of [H( S T AB) + H(aer Therefore, insummary:
(aero)H(aero)R = 30,600 N-m-sec (22,600 ft-lb-sec);RCS
TotalMission
(trans)tRanS |= 5780 N-m-sec (4275 ft-lb-sec);
RCSTotalMission
(STAB)HRCS |= 612,000 N-m-sec (452,090 ft-lb-sec);
TotalMission
H(LC)RCS
TotalMission
H(m)RCS
= 32,200 N-m-sec (23,800 ft-lb-sec);
= 362,000 N-m-sec (267,512 ft-lb-sec);
TotalMission
(total)HRCS
TotalMission
= 1,042,580 N-m-sec (770,400 ft-lb-sec).
IV-68
e. Summary of Angular Momentum Requirements for CMG and RCSControl - The summary of angular momentum requirements discussedin the previous sections is presented for CMG data in Table IV-17;for RCS data in Table IV-18.
The results are tabulated for three major vehicle attitudes: theX-POP attitude with a 1 degree angular misalignment about the Yand Z axis; the Z-POP attitude with a 1 degree angular misalign-ment about the X and Y axis; and the X-IOP attitude calculatedfor maximum angular momentum requirements. For the X-POP andZ-POP cases, the 1-degree misalignment is considered conserva-tive since in reality the control system should be capable ofholding the desired attitude to within 0.5 degree.
The X-IOP attitude allows the largest freedom for maneuver, sincerotation can take place about the X axis and about the axis per-pendicular to the orbital plane. The angular momentum require-ments for maneuver are calculated on the basis of a 2 degree/
minute maneuver rate about the Z axis (the axis having the largest
moment of inertia). This same value of H (m a n ) is then considered
as the H(m a n ) requirement for both the Z-POP and the X-POP atti-
tudes. This is definately proper for the Z-POP attitude consid-eration since any maneuver must be only about the Z axis. Giventhe X-IOP maneuver capacity, the X-POP attitude maneuver ratewill be much larger since the moment of inertia about the X axisis 10 times smaller than about the Z axis.
The angular momentum requirement, H (trans), for absorbing theresidual momentum left by the baseline system when the CMG orRCS takes over control is independent of the final attitude as-
sumed. Therefore, H ( t r a n s ) has the same value for X-POP, Z-POP,and X-IOP for the given control system (i.e., CMG or RCS).
The magnitude of the peak cyclic momentum vector (Table IV-17)is the largest for the X-IOP attitude and the Z-POP attitude withvalues of 5300 ft-lb-sec and 5200 ft-lb-sec, respectively. TheX-POP attitude has comparatively small peak cyclic momentum com-
ponents. H(a e r
) is also comparatively smaller taken as 5% of
IH a + IHcp
IV-69
Table IV-17 Summary of Angular Momentum Requirements for CMG Control
X-POP 1-deg Misalign- Z-POP 1-deg Misalign- X-IOP Maximum atment about Y & Z ment about X & Y A = 44.8 deg
Momentum Source N-m-sec ft-lb-sec N-m-sec ft-lb-sec N-m-sec ft-lb-sec
ax 0 0 60.3 44.5 1740 1280ax
H 1580 1170 1600.0 1180.0 0 0ay
HaZ 1540 1140 0.0 0.0 0 0
Ia I = (Hax + 11 + Ia2) 2240 1655 1600.0 1180.0 1740 1280
H 282 208 4.82 3.55 0 0cxp
H 127 94 127.0 94.00 5050 3730cyp
H 171 126 7050.0 5200 5100 3760czp
IHI = (H2 + H2 + Hp)2 352 260 7050.0 5200 7180 5300cc xp cyp cZp
HXIaero) =.05°[Idl + iPp1
130 95 130 95 130 95
H(STAB aero) + + H(aer) 2720 2010 8770 6475 9040 6675HC. G = IHa + IRc p
H(maneuver) 2 deg/min 5820 4300 5820 4300 5820 4300Z-Axis
H(transition) 3320 2450 3320 2450 3320 2450
Table IV-18 Summary of Angular Momentum Requirements for RCS Control
X-POP 1-deg Misalign- Z-POP 1-deg Misalign- X-IOP Maximum atMomentum ment about Y & Z ment about X & Y X = 44.8 degSource N-m-sec ft-lb-sec N-m-sec ft-lb-sec N-m-sec ft-lb-sec
H (g g ) 612,000 452,090 6,200,000 4,558,000 8,800,000 6,496,210
H(aero) 30,600 22,600 30,600 22,600 30,600 22,600
H(LC) 32,200 23,800 32,200 23,800 32,200 23,800
H(m) 362,000 267,512 362,000 267,512 362,000 267,512
H(trans) 5,780 4,275 5,780 4,275 5,780 4,275
H ( T otal) 1,042,580 770,400 6,630,580 4,876,400 9,230,580 6,814,400
IV-70
The magnitude of the accumulated momentum per orbit (Table IV-17)for CMG control is largest for the X-POP attitude but is only ap-proximately 20% smaller for the Z-POP and X-IOP attitudes.
Assuming a single Skylab ATM-type CMG having an angular momentumcapacity of 2300 ft-lb-sec, the X-POP attitude requires two CMGs,the Z-POP attitude requires three CMGs, and the X-IOP attituderequires three CMGs. This is arrived at by assuming the requiredangular momentum capacity to be the larger of the stabilization
(STAB) (man)requirements, H( , or the maneuver requirement, H . Inthe X-IOP attitude, the calculated value is actually just slightlyabove that which could be handled by three CMGs. If one extraCMG is added to each described attitude case for contingency pur-poses, the totals are:
For X-POP--three CMGs;
For Z-POP--four CMGs;
For X-IOP--four CMGs.
The RCS angular momentum requirements are summarized in TableIV-18 for each attitude, and are largest for X-IOP and smallestfor X-POP.
The requirements for compensating aerodynamic drag, transition,maneuvering, and limit cycle allowance are assumed identical forall attitudes since in all cases they are smaller than the gravity-gradient compensation requirements.
f. Sizing CMG and a Low-Thrust RCS to ControZ the Orbiter - Asmentioned previously, CMGs were recommended as the preferred sys-tem for controlling the attitude of the Shuttle orbiter. Thisrecommendation was based primarily on the probable contaminationof astronomy instruments, should RCS firings be used for Shuttlecontrol. However, there is still enough resistance to the CMGsystem to justify sizing CMGs and a low-thrust RCS. Consequently,this section of the report will identify the weight and power forboth types of systems. For each of the systems, sufficient re-dundancy was included to obtain a high reliability (about 0.999)for a 7-day mission.
IV-71
1) Weight Comparison - The angular momentum requirements for CMGand RCS control are summarized in Tables IV-17 and IV-18. InFigure IV-16, plots are shown of total RCS weight as a functionof total impulse for cold gas, hydrazine monopropellant, and hy-drazine bipropellant using the specific impulse values noted.
(design yaw rate 2 deg/m
Impulse Required for
7 Days, Z-POP
nin) /
Impulse Required for7 Days, X-IOP
l
I .AOS'O9
Total Impulse, N-sec x 10 -5
l I I I I I I I0 4.5 9 13.5 18 22.5 27 31.5
Total Impulse, lb-sec x 10 -4
Figure I V-16 RCS Weight
In estimating the RCS weight it was assumed that thrusters, con-trol valves, and other critical components were designed in re-dundant pairs. The vertical lines on the figure indicate theimpulse required for RCS control of the Shuttle orbiter for 106orbits and 28 maneuvers during a 7-day mission. For missionslonger than 7 days, the impulse and weights can be estimated bya direct ratio of the orbits.
IV-72
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In sizing the CMG systems it was assumed that ATM-type CMGs wouldbe used to satisfy the momentum requirements. Each ATM CMG has amomentum capacity of 3120 N-m-sec (2300 ft-lb-sec). SufficientCMGs were provided to satisfy the maximum momentum requirementin terms of stabilization and aerodynamics, maneuvering, ortransition. In addition, an extra CMG was added for contingencypurposes. The total ATM CMGs for each inertial attitude were:
X-POP--three CMGs;
Z-POP--four CMGs;
X-IOP--four CMGs.
Table IV-19 is a summary of the weight requirements for a hydra-zine monopropellant RCS and a CMG system for each of the inertialattitudes considered. The CMG weight includes the cabling andmounting hardware plus all necessary electronic hardware. TheRCS shown is a hydrazine monpropellant system with a specificimpulse of 170 seconds. The weight difference between the RCSand CMG systems varies between 500 and 570 kilograms (1100 to1250 lb) with the CMG system weighing anywhere from 1.6 to 5times the RCS.
Table IV-19 Weight Summaries for CMG and RCS MonopropeZZantSystems
IV-73
X-POP-I* Z-POP-It X-IOP-I§
Weight CMG System 705 938 938Requirements, (1555) (2068) (2068)
kg (Ib) RCS 135 440 600
7 days (298) (972) (1324)I = 170 secsp
*Requires viewing with both telescopes parallel to the X-axiswith possible reflection interference from Shuttle surfaces.Requires deployment of telescope gimbals out of the payloadbay. Requires coarse gimbal pointing variations (28 or 14deg) in one plane during a 7-day mission.
tRequires coarse gimbal pointing variations (28 or 14 deg) inone plane during 7-day mission.
§No constraints on Y and Z axis. Does not require coarsegimbal freedom for pointing. Requires complex maneuvers todesaturate CMGs.
2) Electrical Power Comparison - In Table IV-20, the electricalpower requirements for both CMG control and low-thrust RCS con-trol are presented for the attitudes of X-POP, Z-POP, and aworst-case X-IOP. The values are conservative in the case ofpeak CMG power, as an extra CMG is carried to provide the requiredcapability with one CMG inoperative. The estimates apply to thecase of design maneuver rates of 2 degrees/minute about the axisof greatest inertia.
Table IV-20 Power Summaries for CMG Systems and RCS
X-POP-I Z-POP-I X-IOP-I l
The power requirements for the RCS system will not vary apprecia-bly whether the RCS uses cold gas, N2 H4 monopropellant, or a bi-propellant.
The power requirements for the CMG system include the power forthe CMGs and the CMG inverter assemblies. The RCS power is forthe operation of the RCS thruster control circuitry.
4. Recommended Solar Payload Configuration
The recommended configuration for the ASM solar payload is thenondeployed concept presented in Figure IV-15. The recommendedinertial attitude for the Shuttle orbiter during solar observa-tions is the Z-POP.
Table IV-21 is a summary of the pertinent characteristics for eachof the configurations and inertial attitudes considered in thisanalysis. As can be seen from the table, the baseline deployedX-POP and the modified deployed X-POP concepts of Figures IV-12and IV-13 are not compatible with the new ground rule that states"deployed payloads cannot exceed 13.72 meters (45 ft) in length."Therefore, these concepts are not acceptable. The nondeployedX-POP concept of Figure IV-4 satisfies the new ground rules, butdoes have the undesirable characteristic of receiving reflectionsfrom the Shuttle structure that could degrade the performance ofthe Coronagraphs. The final two concepts, nondeployed Z-POP andnondeployed X-IOP, both satisfy the revised ground rules and pro-vide excellent viewing potential for the instruments. There isa total weight penalty of approximately 187 kilograms (413 lb)but the viewing advantages are significant enough to recommendthe nondeployed concept shown in Figure IV-15.
IV-74
Power, CMG 144 Average 192 Average 192 Averagewatts System 576 Peak 768 Peak 768 Peak
RCS 15 Average 36 Average 45 AverageSystem 135 Peak 135 Peak 135 Peak
Comparison of SoZar PayZoad Concepts
It should be noted that although the Z-POP inertial attitude isrecommended for the solar payload, the CMG stabilization systemfor the Z-POP will also provide the capability for the X-IOP.The primary reason for selecting the Z-POP for the solar payloadwas to minimize the maneuvering that would be required to desat-urate the CMGs each orbit. With a Z-POP inertial attitude, themaneuvers would be very small angles each orbit that would re-sult in little, if any, loss in observation time.
IV-75
CHARACTERISTICS
Weight,Compatible with kg Compatible with
CONCEPT Stellar (lb) Viewing New Ground Rules Comment
Deployed X-POP Yes 13,068 Excellent No Unacceptable(baseline) (28,809)
Modified Deployed No 13,703 Excellent No UnacceptableX-POP (modified (30,209)baseline)
Nondeployed X-POP Yes 12,227 Shuttle Yes Undesirable(26,956) Reflec-
tions
Nondeployed Z-POP Yes 12,414 Excellent Yes Acceptable(27,369)
Nondeployed X-IOP Yes 12,414 Excellent Yes Acceptable(27,369)
TabZe IV-21
C. ON-ORBIT ACCESS
At the final review of the 9-month Astronomy Sortie Missionsstudy, conducted at MSFC in September 1972, the Sortie OpticalAstronomy Group requested that the follow-on study include inves-tigation of several alternative concepts to provide on-orbitshirtsleeve access to the focal planes of the stratoscope III,infrared, and photoheliograph telescopes. These alternative con-cepts were defined by Martin Marietta and discussed at a coordin-ation meeting held at MSFC in February, 1973. At this meeting,one of the alternatives was selected as the concept to be recom-mended for Astronomy Sortie missions. At the same meeting, theCOR NASA/I4SFC deleted the requirement for on-orbit access to theinfrared telescope, However, all data resulting from the inves-tigetion are included in this report.
1. Ground Rules and Assumptions
The following ground rules and assumptions were adopted for theon-orbit access studies.
1) The telescopes to be studied include--- The 100-centimeter photoheliograph being defined in theLarge Solar Observatory (LSO) study under contract toNASA/MSFC,
- The stratoscope III will be adopted from the 120-centi-meter design being done by Itek,
- The 100-centimeter IR telescope, based on the work performedby Martin Marietta and Itek earlier in this study;
2) The gas bearing concept defined for the AMES C141 IR tele-scope will be analyzed to determine applicability to Astron-omy Sortie missions;
3) Sortie Lab data were extracted from Reference 2 and updatedby Reference 7.
2, Selection of Alternatives
The advantages of providing shirtsleeve on-orbit access can be
grouped into scientific and engineering categories,
IV-76
Scientifically, it is advantageous to be able to make adjustmentsto the telescope and instruments (to ensure optimum performance ofthe equipment) before starting an observation. The adjustmentscan be done remotely, from the Sortie Lab. However, this requiresadditional hardware and adds to complexity of operation. Astron-omers are also used to monitor image quality during an observa-
tion, in order to judge such factors as "seeing" and atmosphericextinction. These factors, of course, do not apply in space,but others, such as telescope performance or guide errors, couldstill be monitored. Again, remote observation is possible butmore complex.
Engineering advantages for providing manned access to the equip-ment include the capability for inflight repairs, and avoiding
some automation, such as remote filter selection or moving a new
instrument into the telescope optical axis by remote control.
When these scientific and engineering motivations for manned
access are analyzed, it is apparent that only one calls foraccess during an observation. All except the monitoring of imagequality take place between observations; thus, if remote moni-
toring and some corrective functions such as focus adjust areprovided, schemes may be considered that provide access onlybetween observations.
The baseline accommodation mode (Ref 1) does not incorporateprovisions for on-orbit access. Telescopes are located onhemispherical viewing mounts, attached to the pallet. FigureIV-17 shows the baseline concept and the on-orbit access con-
cepts discussed below.
Two concepts for providing manned access to the telescope focalplane were studied; the airlock/hangar concept, and the foldedoptics concept.
The airlock/hangar concept allows the telescope to operate onits hemispherical viewing mount (located on the pallet) duringobservations. It is brought back to the launch position foraccess from the Sortie Lab. This scheme does not allow accessduring observations, therefore, monitoring and some adjustmentsmust be done remotely. Also, the Sortie Lab airlock must beoperated each time access is required. The advantages of this
scheme are that it minimizes the impact on the telescope designand allows the telescope to be pointed over a large part of thesky without reorienting the Shuttle orbiter.
IV-77
The second concept for providing shirtsleeve access brings thetelescope image into the Sortie Lab. Two variations of the foldedoptics concept were developed; one uses a gas bearing; the otherincorporates a mechanical gimbal for supporting the telescope(this scheme allows continuous access to the instruments, evenduring observations). Bringing the focal surface into the SortieLab involves impacts on the telescope optics, as discussed inthe following paragraphs.
3. Telescope Definitions
The three intermediate class telescopes selected for the on-orbitaccess study are defined in Reference 1. These baseline designsdo not include provisions for on-orbit shirtsleeve access. Briefdescriptions of the baseline telescopes are included below, alongwith descriptions of the modified telescopes, to illustrate themodifications required to achieve compatibility with shirtsleeveon-orbit access concepts.
a. Baseline
1) Photoheliograph - The baseline 100-centimeter, f/3.8-f/50photoheliograph is a scaled-up version of the 65-centimeterinstrument defined by BBRC under NASA contract NAS8-30190. Thebaseline design (Fig. IV-18) has an overall length of 4.58 meters(180 in), which represents the maximum allowable length of thetelescope when integrated into the baseline solar payload (whichincludes two telescope groups). The instruments are located ina housing lying along the side of the telescope tube.
2) Stratoscope III - Figure IV-19 shows the baseline 120-centi-meter, f/2.2-f/12 stratoscope as defined by Itek. The opticalsystem was modeled after the Itek LST, and the telescope is in-tended to serve as an engineering model of the LST. Overalllength of the telescope is 4.22 meters (166 in).
3) Infrared Telescope - The baseline 100-centimeter, f/1.5-f/10IR telescope was defined by Martin Marietta and Itek (Fig. IV-20).The design is based on an integral tank approach, using liquidneon as the coolant. Coolant is circulated by capillary forces,rather than a pump system. This method avoids the disturbancescaused by mounting rotating machinery and fluid flow control com-ponents on the telescope. The design shown allows either of twoinstruments to be remotely positioned at the telescope focal plane.
IV-78
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1) Photoheliograph - Itek, under separate contract to NASA/MSFC,has studied a 1.5 meter photoheliograph for the Large Solar Ob-servatory, a 1.0-meter telescope for balloon use, and another1.0-meter version for Shuttle Sortie flights (Fig. IV-21). Thefigure shows the results of repackaging the instruments andlocating them at the aft end of the telescope. The entire in-strument compartment of this concept must be pressurizable forshirtsleeve access. A small, remotely operated door is used toclose the image hole in the pressure bulkhead during pressuriza-tion. The compartment structure is sized to allow shirtsleeveaccess to the entire instrument complement. The interface tothe Sortie Lab airlock is located at the aft end of the struc-ture. A thermal/meteoroid barrier door covers the opening atthe interface while the telescope is in operating mode.
2) Stratoscope III - The airlock/hangar concept of the stratoscopeIII (Fig. IV-22) represents a modification of the baseline tele-scope. The pressurizable instrument compartment shown is essen-tially the same as shown on the photoheliograph. However, witha Cassegrain system, the image is brought to the instrumentsthrough a hole in the secondary mirror, which requires using adifferent door assembly to seal the image hole in the pressurebulkhead.
3) Infrared Telescope - Extensive modifications were made tothe baseline telescope to achieve compatibility with the airlock/hangar access concept. For the baseline design, the instrumentswere located within the very cold confines of the wraparoundtank structure. Since the temperatures involved with this ap-proach are far too low for man to tolerate in shirtsleeves, theinstrument compartment was relocated to the outside of the tank,as shown in Figure IV-23. The cylindrical structure of thepressurizable compartment is made of fiberglass laminate inorder to minimize heat losses and the resulting high neon boil-off rates. Sealed foam insulation is attached to the inside sur-faces of the compartment to prevent condensation of moisture onthe cold surfaces.
No provision is included for remote positioning of instruments atthe focal plane, since this would be accomplished during theaccess periods. Because the instruments are no longer operatingin a very cold environment, a cryostat was added to providecooling.
IV-87
The instruments, although cryogenically cooled, can be packagedto be handled by a man in shirtsleeves. However, a major prob-lem that needs further study is associated with condensationand freezing of moisture on the cold surfaces in the vicinityof the door in the pressure bulkhead. Another problem involvesthe need for disconnecting and reconnecting the cryogen linesbetween the cryostat and the instrument.
c. FoZded Optics Concepts - The folded optics configurationsrepresent modifications to the airlock/hangar designs describedabove. The major change was relocating the instrument comple-ments inside of the Sortie Lab. A configuration for the IRtelescope in the mechanical gimbal is not shown (Fig. IV-23),since the requirement for on-orbit access to this telescope wasdeleted before definition of the mechanical gimbal.
1) Photoheliograph - This telescope incorporates a second foldmirror (located in the light tube between the first fold mirrorand the instruments) that deflects the image through the bearinginto the Sortie Lab. In the gas bearing concept (Fig. IV-24).this mirror is fixed, because the telescope and instrumentpackage move together as a unit. However, the mechanical gimbalconcept (Fig. IV-25) requires that this mirror translate alongthe axis of the light tube, and tilt to keep the light beam fixedwith respect to the instruments. Internal image motion compensa-tion (IMC) is required for both the gas bearing and mechanicalgimbal concepts. The primary differences in the IMC are due tothe fact that the gas bearing will provide a base stability ofapproximately 0.1 sec, while the mechanical gimbal will providea base stability of approximately 1.0 min.
In both concepts, the instrument complement is rigidly mountedto the telescope support structure. The support structure andthe removable cover allow the instruments to operate in avacuum without having a window in the path of the image beam.If desired, the instruments may be operated under pressure bymoving a pressure window (located in the support structure) intoposition. With the window in position, the cover may be removedfor access to the instruments. Various instrument controls andmonitoring devices may be mounted on the housing to provide"hands-on" operations.
IV-88
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2) Stratoscope III - The fold mirror in this telescope is locatedon the centerline of the telescope and remains fixed for the gasbearing concept (Fig. IV-26). For the mechanical gimbal concept(Fig. IV-27), the mirror need not translate, but must be tilted.Since the mirror is located at the intersection of the gimbalaxes, a 45.8-centimeter (18-in) long slot must be provided throughthe telescope structure to allow passage of the image to theinstruments. As in the case of the photoheliograph, IMC is re-quired for both concepts. Instrument mounting and access fea-tures are identical to those described for the photoheliograph.
3) Infrared Telescope - The fold mirror of this telescope islocated on the centerline of the telescope and remains fixed forthe gas bearing concept (Fig. IV-28). No IMC is required forthis telescope on the gas bearing. Since the instruments forthis telescope must operate at cryogenic temperatures, a cryostatis provided. No instrument cover is provided since only oneinstrument will be operating at any time. The instrument casemust be pressure-tight to allow operation of the detector in avacuum. When instrument replacement is desired, the pressurewindow located in the support structure is moved into position.This configuration involves the need to disconnect and reconnectthe cryogen lines between the cryostat and the instrument.
d. Mass Properties - Summary weight statements for the telescopeconfigurations developed for on-orbit access are presented inTables IV-22 and IV-23. The weight statements are presented inthe form of removals and additions to the baseline telescopes.For the gas bearing and mechanical gimbal concepts, the telescopeinstruments are included in the weight statements (para 5) forthe concepts.
4. Telescope Performance Analyses
Telescope performance analyses were conducted to provide a com-parison between the baseline telescopes and the telescopes de-signed for the gas bearing access concept, which involves bringingthe telescope image into the Sortie Lab. Performance of the tele-scopes in the airlock/hangar concept is the same as the base-line telescopes. Similarly, performance of the photoheliographand stratoscope III in the mechanical gimbal concept will be es-sentially the same as for the gas bearing concept. The require-ment for on-orbit access to the IR telescope instruments wasdeleted during the study, but results of the investigation con-ducted up to this point.in time, are included.
IV-99 and IV-100
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Table IV-22 Summary of Telescope Weights for Hangar Concept
IR Telescope [kg (lb)]
Baseline IR Telescope (Ref 1)
Remove From Baseline:InstrumentsInstrument MechStructure and InsulationNeon
Add:Structure and InsulatiuonHeaviest InstrumentNeon
Modified Telescope
Alternative Instrument (In Sortie Lab)
Modified Telescope with Alternative
103.922.7
320.7349.3
200.561.7
158.7
1988.1 (4383)
-796.5 (-1756)
+420.9 (+928)
(229)(50)
(707)(770)
(442)(136)(350)
1612.5 (3555)
+42.2 (+93)
1654.7 (3648)
Photoheliograph [kg (lb)]
Baseline Photoheliograph (Ref 1)
Remove From baseline:Instrument HousingTrunion Ring
Add:HangarForward BulkheadAft BulkheadCYLInsulationMeteoroid BumperTrunionInternal FrameDoor and MechIns and M/P on Door
Modified Telescope
133.8 (295)9.1 (20)
52.261.7127.041.734.510.493.927.22.3
997.9 (2200)
-142.9 (-315)
+450.9 (+994)(115.0)(136.0)(280.0)(92.0)(76.0)(23.0)(207.0)(60.0)(5.0)
1305.9 (2879)
Stratoscope III [kg (lb)]
Baseline Stratoscope III (Ref 1)
Remove From Baseline:Instrument Compartment
CompartmentInsulationMisc.
1792.1 (3951)
-158.3 (-349)
124.710.922.7
(275)(24)(50)
Add:Hangar (Same as PHG)
Modified Telescope
1633.8 (3602)
+450.9 (+994)
2084.7 (4596)
IV-107
Table IV-23 Summary of TeZescope Weights for Gas Bearing and MechanicaZ GimbaZ Concepts
IR Telescope [kg (lb)]
Baseline IR Telescope (Ref 1) 1988.1 (4383)
Remove From Baseline: -947.5 (-2089)
Instruments 103.9 (229)
Instrument Mech 22.7 (50)
Structure and Insulation 320.7 (707)
Neon 349.3 (770)Cooler 61.2 (135)
Helium 10.0 (22)Frame 66.2 (146)
Meteoroid Frames 13.6 (30)
Add: +383.8 (+844)
Main Frame 111.1 (245)
Bulkhead 84.4 (186)Foam Insulation 3.6 (8)
Super Insulation 4.0 (9)Tank Screen 1.4 (3)Heat Ex. Tubes 2.3 (5)
Neon 158.7 (350)Fold Mirror 9.1 (20)
Fold Mirror Supp. 8.2 (18)
Modified Telescope 1423.4 (3138)
Photoheliograph [kg (lb)]
Baseline Photoheliograph (Ref 1) 997.9 (2200)
Remove From Baseline:Instruments and Housing -351.5 (-775)
Add:Attach Structure +22.7 (+50)
Modified Telescope 669.1 (1475)
Stratoscope III [kg (lb)]
Baseline Stratoscope III (Ref 1) 1792.1 (3951)
Remove From Baseline: -919.4 (-2027)
Instrument Compartment 124.7 (275)Insulation 10.9 (24)
Electrical 63.9 (141)Instruments 699.0 (1541)
Misc. 20.9 (46)
Add: +77.1 (+170)
Bulkhead 17.2 (38)Insulation 5.4 (12)Attach Structure 9.1 (20)Fold Mirror Install. 45.4 (100)
Modified Telescope 949.8 (2094)
IV-108
a. Wavefront Error Effects - One extra mirror is needed to con-vert the conventional photoheliograph, stratoscope III, and IRtelescope optics so that their focal surfaces are inside theSortie Lab. If this mirror was absolutely flat and smooth, thenit would not affect wavefront error and so would not affect op-tical resolution. Unfortunately, some residual curvature inevit-ably is found in a "flat'--this curvature introduces some astig-matism in the folded image.
1) Photoheliograph - The 1-meter wavefront error budget for thephotoheliograph (developed for NASA/MSFC) came to 0.10 A rms at0.63 pm. A fold mirror to bring the focal surface into the SortieLab will have an error of approximately 0.01 A rms. A smallernumber might be achieved, but this is already near the limits oftest accuracy. When the 0.10 A rms photoheliograph is rss'd withthe 0.01 A rms fold mirror, the result is an increase of about0.0005 A rms in the system. Figure IV-29 shows how system modu-lation transfer (MTF) varies with wavefront error (a 30% obscura-tion, and zero motion, were assumed). From this figure it is
possible to relate angular resolution to wavefront error (Fig.IV-30). Resolution was defined at 30% modulation because somany astronomical scenes are of low contrast. The flat is seento introduce less than 1% loss in resolution due to wavefronterror.
2) Infrared Telescope - The baseline IR telescope wavefronterror has been estimated at 0.64 A rms at 0.63 pm (or 0.1 Xrms at the design wavelength of 4 vm). This large error wasassumed because of the severe thermal environment. In thecase of a small fold mirror, however, it should be possibleto keep thermal gradients under control so that the system per-formance should not be further degraded.
3) Stratoscope III - The stratoscope III wavefront error budgetsums to 0.050 A rms. This is an improvement over the photo-heliograph because the optics are simpler, the thermal environ-ment is less rigorous, and a desire to duplicate the LST budgetwas assumed. If it is again assumed that a fold mirror intro-duces 0.01 X rms, then the system error increases 0.001 A rms.From Figure IV-30, the resolution loss is less than 2%.
Adding a fold mirror to any of the three baseline telescopesdoes not introduce enough wavefront error to degrade resolutionsignificantly.
IV-109
e = 30%
No Image Motion
w = 0.06 X rms
w = 0.08 X rms
----- w = 0.10 X rms
0.2
00 0.2 0.4 0.6
Angular Frequency (normalized)
Figure IV-29 Effect of Wavefront Error on MTF
0.8 1.0
0.5 r
0.4
0.3 F-
0.2
0.1
MTF = 30%
e = 30%
No Image Motion
01 1 1 1
0 0.05 0.10 0.15
Wavefront Error, X rms
Figure IV-30 Normalized Angular Resolution UsWavefront Error
1.0
0.8
0.6
low
E.
a
0
a
4..
0'-4
'00
0.4
'0
-41-4
0.r
U-c
0~
v
wp
to
nr
b. Obscuration Effects - Having determined that the fold mirrorused to bring the focal surface into the Sortie Lab introduces anegligible wavefront error, the obscuration of the aperture andthe resulting system performance must now be considered. Angularresolution is degraded by wavefront error, obscuration, and imagemotion. It was assumed that the fold mirror does not affectimage motion.
In this analysis a gas bearing gimbal and a fold distance of 1.75meters (68.9 in) (Fig. IV-31) were used. Due to the addition ofa cold stop to the IR instruments, this distance was later in-creased to 2.0 meters (78.7 in). Performance was not recalculatedbecause on-orbit access requirements for the IR telescope weredeleted, and the photoheliograph and stratoscope III are notsensitive to this change.
Gimbal at System cg
Instruments
1.75 meters
Figure IV-31 Gas Bearing GimbaZ at System Center of Gravity
IV-111
l
To achieve highly stable pointing of the telescope, the systemmust be gimbaled at its center of gravity (cg). This dictatesthat the fold mirror must be at the telescope cg, since the sys-tem cg is on a line between the cg of the telescope and the in-struments. If the fold mirror were at some other location, thenthe tunnel through the gimbal would have to be off center andthe gimbal would be enlarged.
With the location of the fold mirror fixed at the telescope cg,and the fold distance fixed by considerations of telescope diam-eter, gimbal size, and Sortie Lab requirements, the vertex backfocus is no longer a variable (Fig. IV-32). However, vertexback focus can influence obscuration and, therefore, telescoperesolution. If too much resolution is lost by folding thefocal surface to the side, the fold mirror must be moved andcounterweights added to the telescope, or changes must be madeto basic parameters, such as the focal lengths of the primarymirror and telescope. These changes will impact field of view,plate scale, manufacturability, and tolerance to misalignment.
D t "l
__ S
I ILhi
I-h l
B= g-h
S= (f -B)/(m + 1)
=D (B + P)/(f + P) . es D/Dp
= (B + P)/(f + P)
Where:
e = Obscuration of secondary mirror;
D D = Dimeters of primary and secondary mirrors;
B = Vertex back focus;
f = Focal length of system (negative in aGregorian telescope);
m = Magnification _ f/p;
P = Focal length of primary mirror;
g = Fold distance;
h = Location of fold mirror.
Figure IV-32 Basic Two-Mirror TeZescopeRelations
IV-112
Figure IV-22 shows how obscuration by the secondary mirror isrelated to vertex back focus. If there were negligible fieldof view, then this would be the system obscuration as well. How-ever, the photoheliograph, IR telescope and stratoscope III allhave high-quality data fields of view on the order of 1.45 mradto 1.75 mrad (5 to 6 min) in diameter, so that a baffle system isneeded, and the obscuration will be increased a few percent. Ifoff-set guiding is to be used on stratoscope III, the field ofview must be about 8.72 mrad (30 min) and the obscuration isincreased substantially. Unfortunately, there are no simpleformulae for calculating baffle obscuration so that educatedguesses, graphical analyses, or computer programs are required.
The fold mirror itself can contribute to obscuration if it isplaced so far forward that it blocks light converging from theprimary toward the secondary mirror (Fig. IV-33). This problemhas been investigated by graphical analysis.
a
b
c
Secondary Fold Primary
Forward
Note: 1. Rays a and b can pass to the focus.2. Ray c is intercepted by the fold mirror, which is too far forward.3. The obscuration is therefore b/a rather than c/a. The solution
is to move the fold closer to the primary mirror.
Figure IV-33 Obscuration by FoZd Mirror
IV-113
Obscuration and resolution changes for each of the three telescopesare discussed in the following paragraphs.
1) Photoheliograph - The airlock/hangar/photoheliograph is the1.0-meter f/3-f/40 system shown in Figure IV-21. The field ofview is 1.75 mrad (6 mTn). Because this is a Gregorian telescope,the system focal length and magnification are negative (Fig. IV-32).
The fold mirror could be arranged to take the light out the same
side as the off-axis light tunnel or out the opposite side (Fig.
IV-34). The vertex back focus (and therefore obscuration) andtube length are less in the first case, but the mounting to thegimbal becomes more complicated, both structurally and dimension-ally. Both fold options were investigated, the results are shown
in Figure IV-35. Because the cg will be near h=l.0 meter(39.4 in), the fold mirror appears to introduce some extra ob-scuration. However, the obscuration is in fact set by the struc-ture supporting the secondary mirror, and is approximately 30%.
The conclusions reached are that the fold mirror can be movedover a wide range without introducing excess obscuration, and
that, as far as image quality is concerned, the fold may be awayfrom or through the tube. Some other factors affecting choice
of fold direction will be discussed later.
2) Infrared Telescope - The airlock/hangar IR telescope is the
1.0-meter f/1.5-f/10 system described in Reference 1. The cgof the folded version has been estimated to be at h=0.6 meter(23.6 in) in front of the primary. Guiding will be done withan auxiliary telescope, so that the field of view is only 1.45mrad (5 m'n).
Figure IV-36 shows how the location of the fold mirror affectsobscuration. The 25% f/1.5-f/10 telescope obscuration is in-creased to 35% with the mirror at the cg h=0.6 meter (23.6 in).From Figure IV-37, the resulting resolution loss is about 6%.This figure was derived from MTF analysis, and includes theeffects of 0.1 X rms wavefront error at 4.0 pm, 0.0015 mrad(0.4 sec) rms guide error, and a 0.1-millimeter (0.0039-in) sen-sor. The optics themselves suffer a 30% resolution loss, butthe system loses only 6% because it is sensor limited. For thisreason the folded system would not benefit as much from improve-ments in the sensor state of the art as would the unfolded sys-tem.
IV-114
/Heat Dump
PrimarySecondary /
b
t Heat Dump 1
Figure IV-34 Two Fold Options for Photoheliograph
30 - 30Baseline Baseline
a 20 a 20
a i 7 = r Secolldary/mirr4r Secondary Mirro
10 u -l
a t~ ~~~~~ - Fold Mirror
Fold mirror t t0 cg i | Icg
0i I 0 I I0 0.5 1.0 1.5 2.0 0.5 1.0 1.5 2.
Fold Mirror Distance h, meters Fold Mirror Distance h, meters
a) Fold away from tube (Option a) b) Fold through tube (Option b)
Figure IV-35 Obscuration Analysis for Folded PhotoheZiograph (1-meter f/1.5 - f/10. g = 1.75 meters)
IV-115
0
- - Baseline
Secondary Mirror
0.2 0.4 0.6
Fold Mirror Location h, meters
a) f/1.5 - f/10
0.8 1.0
Baseline_ _ _ _ _
'-- Baffle
t -- Secondary Mirror
cgi i I
0 0.2 0.4 0.6 0.8
Fold Mirror Location h, meters
b) f/1.5 - f/15
Figure IV-36 Obscuration Analysis for Folded IR Telescope
10
0
-10
-20
-3020 30 40 50
Obscuration e, percent
IV-37 Change in Resolution of IR Telescope vsObscuration
60
60
U4
w0
jU0a
0go
0
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4.U
40
20
0
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0aIJ
0.a 20
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UM
o 0
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a-40a
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toa.0
Figure
IV-116
cg
Figure IV-36 shows that an f/1.5-f/15 telescope could maintainthe baseline obscuration of 25%. However, this causes a 33%loss in field of view, and the tolerance to misalignment of thesecondary mirror is 3.4 times worse. It is not possible toachieve 25% obscuration in a folded f/10 system by changing pri-mary f/number, or by moving the cg or locating the fold mirroroff the cg.
An investigation was made to determine the possibility of re-ducing obscuration to 25% by changing the fold distance from1.75 meters (68.9 in) (g in Fig. IV-32). Figure IV-38 showsthat the distance must be reduced to about 1.3 meters (51.2 in).This is insufficient to include the telescope radius, the gasbearing, and space required in the Sortie Lab.
70
60
r.a)o4-IS
a)
Q.00H4-i
u
00O0
a
50
40
30
201 2 3 4
Fold Distance g, meters
Figure I V-38 Effect of Fold Distance onBaseline IR TeZescope
IV-117
This leaves two feasible alternatives: an f/1.5-f/10 telescopewith 6% loss of resolution; or an f/1.5-f/15 telescope with 33%loss of field of view, and tighter tolerancing. The tighter tol-erances may make it impossible to maintain image quality atreasonable cost; therefore, it was concluded that the f/1.5-f/10should be retained as baseline for the folded system.
3) Stratoscope III - The baseline stratoscope III is the 1.2-meter f/2.2-f/12 scaled-down LST. Guiding will be done withoff-set sensors and a 13.08 mrad (45 min) guide field will berequired to ensure high probability of finding adequately brightstars in any part of the sky. LST needs only 6.97 mrad (24 min)but has a larger aperture so it can see objects that are twomagnitudes fainter. Baseline obscuration is 34% (Currently 30%in LST).
From Figure IV-39, it can be seen that the fold mirror can beat the center of gravity without increasing obscuration. The cgis estimated to be approximately 0.5 meter (19.69 in.) in frontof the primary mirror.
40
Baffle for 45 m'n
vJ Baseline
c) _ ~
30
g)
30 . I .I I I l
0 Second0.2 04 06 0.8 1.0
Fold Mirror Distance h, meters
-
' 20
Co
10 I I I0 0.2 0.4 0.6 0.8 1.0
Fold Mirror Distance h, meters
Figure IV-39 Obscuration Analysis of Folded Stratoscope III
[1-meter, f/2.2 - f/12. 45-m~n (13.1 mrad field.g = 1.75 meters.)]
IV-118
c. Other Considerations - Considerations that could impact thedecision to fold the focal surface into the Sortie Lab are dis-cussed in the following paragraphs.
One problem common to both the IR telescope and the stratoscopeIII is thatof diffuse reflections off the outside of the baffle(Fig. IV-40). (In the case of the photoheliograph there is noconic baffle because of the field stop at the primary focus).This diffuse reflection can perhaps be controlled with knife edgeson the baffle. (The University of Arizona is developing methodsfor analyzing such baffle structures.)
Secondary Primary
Stray Light
Figure IV-40 Diffuse RefZection off Bafflein Folded System
Two options for folding the focal surface of the photoheliographwere previously discussed (Fig. IV-34). If the fold is throughthe tube (option b) then the heat dump (solar image at the pri-mary focus) is towards the Sortie Lab and clearance must be al-lowed for the external mirror shown in Figure IV-21. If optiona is chosen, the external mirror can perhaps be eliminated--itis needed only if the heat dump could pose a problem to the Shut-tle or other instruments,
IV-119
j
The cryogenically-cooled IR telescope has a special problem whenit is reconfigured to bring the focal surface into the SortieLab. This is because any warm structure that can be seen by thesensor, directly or by reflection, will generate photoelectronsthat degrade the signal-to-noise ratio. In the unfolded IRtelescope, the image is brought through a warm gimbal that couldget in the detector's field of view.
Figure IV-41 shows'how the problem comes about.
Sensor
Note: For l-meter, f/1.5 - f/10 IR Telescope with
D = (.0951 + .0292/S) X -0.0146.x
Figure IV-41 ProbZem of Warm Structures in IR TeZescopeSensor FieZd of View
No structure is allowed in cone aa if vignetting is to be avoided.In most systems this would be the only consideration; however,in the IR telescope there must be no warm structures within conebb. The diameter, D , is a function of s (which should be as
xlarge as possible) and x (which should be as small as possible).In the case of the l.0-meter f/1.5-f/10 IR telescope with 1.30mrad (5 imfn) field of view, if the cold stop is at s = 0.2 meter
(7.88 ino) and the gimbal is at x = 0.35 meter (13.78 ino), the
IV-120
hole in the gimbal must be Dx = 0.07-meter (2.75 in). If the
telescope mounting to the gimbal at x - 1.0-meter (39.37 in) in-volves warm structures, then it must have a clear opening inexcess of 0.26-meter (10.24 in) diameter.
There are at least two methods of avoiding this problem; however,each method presents difficulties of its own. One method wouldbe to extend a cryogenic sheath into the gimbal, and thereforeeffectively extend the cold stop distance, s. This would probablybe difficult to service and could create a large thermal load.The other method is to use relay optics as a field lens. Thesemust be all reflecting so that several mirrors are needed, and someextra wavefront error will result.
d. Summary - In the preceding paragraphs, some of the problemsthat arise when the focal plane of a telescope is folded intothe Sortie Lab, were analyzed. The fold gives maximum mannedaccess to the focal instruments, but no access to the telescope.In these analyses no compelling optical arguments against thisreconfiguration of the telescopes were found, though problemswere found that will result in slightly poorer performance, alittle more difficulty in achievement, and a slightly highercost.
5. Configuration Analysis
The telescope designs for the concepts discussed below are de-scribed and analyzed in the preceding sections of this report.This section describes the major features and problem areas ofthe shirtsleeve on-orbit access accommodation concepts, and pre-sents mass properties data for the concepts. A summary of thebaseline concept is included for comparison. Sketches of thethree on-orbit access concepts are shown in Figure IV-17, alongwith the baseline concept, which does not provide shirtsleeveon-orbit access.
a. Baseline Concept - The baseline concept (Ref 1) uses theshort Sortie Lab with a single telescope mount attached to thepallet. The baseline includes two mounts; however, only the for-ward mount is shown. This allows a valid comparison to be madebetween the baseline and the on-orbit access concepts. Themount shown deploys the telescope out of the payload bay and pro-vides hemispherical viewing capabilit as well as fine pointingcapability with a-stability of 0.1 sec. Four CMGs, mounted tothe pallet, are used for Shuttle orbiter orientation and stabiliza-tion. Because no shirtsleeve on-orbit access to the telescope is
IV-121
provided activities such as monitoring, adjustments, filter se-lection, and instrument changes are accomplished by remote controlfrom the C&D panel in the Sortie Lab. Inflight maintenance andrepair cannot be accomplished.
bo. Airlock/Hangar Concept - This concept provides for periodicon-orbit access to the instrument compartments of the telescopes.Figure IV-42 shows the airlock mounted to a bulkhead attached tothe aft end of the Sortie Lab. Since this concept does not in-trude into the Sortie Lab, a short Sortie Lab, 4.72 meters (186in.) in length may be used. The telescopes are shown on the pal-let, supported by a mount and gimbal assembly very similar to thecommon mount and gimbal assembly defined in Reference 1. Whenoperating, the telescope is deployed out of the payload bay;when access is desired, it is returned to the launch positionshown. The telescoping airlock is extended to interface with thepressurizable instrument housing or hanger, on the telescope, andthe latches are engaged across the interface. At this time, thedoor in the telescope pressure bulkhead is closed by remote con-trol from the Sortie Lab and pressurization begins. As the pres-surization proceeds, the screwjacks are actuated to provide ten-sion across the interface, which alleviates the large loads thatwould otherwise act on the telescope mount. These loads areconsiderably higher than those imposed on the mount during Shuttlelaunch and return to Earth. The telescoping airlock design isbased on an expandable structure designed, fabricated, and groundtested by Martin Marietta under contract to the Air Force MaterialsLaboratory (Ref 9).
The major anticipated problem area in this concept involves theenvironment to which the instruments are exposed. Pressurizationof the instrument compartment for shirtsleeve access requiresintroduction of the warm and relatively moist atmosphere of theSortie Lab, which could result in condensation of cool surfaces.When the telescope is in the operating position, the thermal en-vironment may be quite different, causing the optical assembliesto change geometry. A special problem is associated with the IRtelescope, which has cryogen-cooled instruments. The instrumentscan be designed with acceptable touch temperatures for handling,and it is assumed that an instrument would only be removed whenit is to be replaced with another unit. However, the replacementunit should be cooled to near operating temperature to minimizetelescope downtime. This requires exclusion of moisture from thecold surfaces of the unit until the telescope is again ready tooperate. The instrument mounting surfaces and the door assembly
on the pressure bulkhead must be kept warm until the atmospherehas been released. Another contamination problem will occur ifthis atmosphere is dumped overboard.
IV-122
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While the problems highlighted above should not prove to be insur-mountable, each one will require careful consideration during theearly stages of development of this concept.
c. Gas Bearing Concept - Bringing the telescope image into theSortie Lab provides continuous access to the instruments for moni-toring and adjustment, even during observations. However, whenthe instruments operate in a vacuum in order to prevent passingthe image through a window, direct access for instrument replace-ment requires suspending observations. Figures IV-24, IV-25, andIV-27 show the three intermediates class telescopes mounted on agas bearing similar to the one used in the Ames C141 IR TelescopeProgram. This bearing is capable of providing a base stabilityof approximately 0.1 sec, but gimbal travel is limited to 0.0436radian (± 2.5 deg) about the three orthogonal axes of the bearing.The centerline of the bearing is offset below the centerline ofthe Sortie Lab in order to accommodate any of the telescopeswithin the payload envelope of the Shuttle orbiter. Counterbalanc-ing is necessary in all cases since the cg of the entire gimbaledmass must lie at the center of the spherical gas bearing. Inaddition, the IR telescope requires a movable counterweight tocompensate for neon boiloff from the telescope cryogen tank. Along Sortie Lab, 7.77 meters (306 in.) in length, is required toaccommodate the large instrument package within the Lab.
Several problem areas are inherent with the gas bearing concept.The first of these is the very limited gimbaling capability of thebearing, which requires maneuvering the Shuttle orbiter when it isdesired to move from target to target. The thermal environmentimposed on the instruments in this concept should be more constantthan for the airlock/hangar concept, but active thermal controlmay still be required. The instruments for the IR telescopeagain cause special problems, similar to problems anticipated inthe airlock/hangar concept. The movable counterweight system,required only for the IR telescope, will add complexity to thisconcept.
The gas supply and scavenging system required for this concept,is of special interest. Two types of systems were investigated:a blow down system that captures and holds the used gas; and arecycling system that filters and reuses the gas. Both systemsretain the "used" gas rather than releasing it where it may be-come a source of contamination. Also, both systems require aground-based gas supply and scavenging system if the bearing is tobe operated at 1 g on the ground. Design parameters for the
IV-125
systems were: helium gas; 0.0018-inch bearing gap; 1.6 scfm,flow rate; Shuttle g level of 0.05; 2000 psi storage pressure(max); and a system mechanical efficiency of 20%.
The blowdown system regulates the 2000 psi storage tank pressuredown to the proper level for the bearing. Pumps are used totake the scavenged gas to a 2000 psi holding tank. No filteringis required since the gas is not reused. This system requiresapproximately 3125 Watts of electrical power, and includes twolarge tanks, each capable of storing 95 cubic feet of helium at2000 psi.
The recycling system stores makeup gas in a 5-cubic-foot tankat 2000 psi. From this tank, the helium passes through a regu-lator into a 3-cubic-foot plenum tank that feeds the gas to thebearing at a pressure of 45 psi. The scavenged gas is pumpedthrough filters back to the plenum tank, and then reused. Anestimated 1600 Watts of electrical power is required for thissystem.
Several variations to these systems were considered, but wererejected because of anticipated deficiencies. A blowdown sys-tem without scavenging capability would be a simpler systemthan the one described above, but dumping of the large amountof helium may cause a contamination problem. Reducing the gasbearing gap would reduce the gas flow rates, but would resultin a more difficult bearing to fabricate.
The recycling system is recommended for this application, due tothe lower power requirement, and the need for one relativelysmall low pressure tank rather than the two large tanks requiredfor the blowdown system with scavenging capability.
d. Mechanical GimbaZ Concept - This concept provides the sameon-orbit access as the gas bearing concept, and the instrumentsare accommodated in the same manner as shown in Figures IV-24and IV-26. The major difference lies in the fact that the gasbearing is replaced with a two-axis gimbal, capable of largergimbal angles. As shown in Figure IV-12, ± 0.26 radian (± 15 deg)is provided in the pitch direction, and ± 0.87 radian (± 50 deg)is possible in roll. This means an important reduction in Shuttlemaneuvering requirements. The mechanical gimbal will have a basestability of approximately 1.0 min. Along with eliminating theneed for the gas supply and scavenging system, this conceptsimplifies the counterbalancing problem. The only requirement isto keep the telescope cg on the telescope pitch axis. A longSortie Lab is required to accommodate this concept.
IV-126
The instrument accommodations are the same for this concept as forthe gas bearing approach, and the same problems can be anticipated.Some added complexity is incurred in the design of the telescopesfor this concept. As discussed previously, the fold mirrors mustbe articulated and the IMC systems must operate from a less stablebase than the gas bearing provides.
e. Mass Properties - Summary weight statements are shown inTables IV-24, IV-25, and IV-26 for the on-orbit access schemes.Weights of the telescopes were taken from Tables IV-22 and IV-23.A summary weight statement for the baseline no-access concept isshown in Table IV-27. Figure IV-43 locates the cg of the variouspayloads on a plot of the Shuttle orbiter cg constraints. Theseweights and cg locations include only the items required whenflying a single telescope. The figure indicates that all ofthese payload cg's are very far forward, and that several exceedthe limits of the Shuttle orbiter cg constraint envelope.
40 r-- -Nominal MaximumLanded PayloadWeight
PHG & S-III G/B -7/
30 --
IR & S-III M/G
____--- IR B/L
S-III B/L
*-- PHG M/G
CG constraint curveextracted from SpaceShuttle RFP No.9-BC421-67-2-40P
20 30
Legend:
PHG - PhotoheliographSIII - Stratoscope IIIIRT - Infrared Telescope
G/B - Gas BearingM/G - Mechanical GimbalB/L - BaselineA/H - Airlock/Hangar
40 50 60 70.
IV-127
PHG
O-4
-I
M.4
3l
0
20 -
10 -
010
Distance from Forward Payload Bay Wall, feet
Figure IV-43 Concept CGs vs Shuttle Orbiter CG Constraints
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IV-1
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6. Operations Analyses
The gas bearing, mechanical gimbal, and hangar/airlock accommoda-tions modes provide the capability for on-orbit servicing opera-tions requiring crew involvement and reducing time available forscientific observation. To assess the effect of these operations,timelines of the functions were prepared and net reductions indata acquisition durations were estimated. As shown in TableIV-28, time requirements per access are not excessive for eithermode, and are therefore feasible for sortie missions.
TabZe IV-28 Access Operations
GAS BEARING ORMECHANICAL GIMBAL HANGAR/AIRLOCK
Access Reduction Access ReductionDuration, in Observa- Duration, in Observa-
TELESCOPE hr: min tion Time, % hr: min tion Time, %
Photoheliograph 2:30 1.6 5:40 2.5
Stratoscope III 2:00 1.0 5:40 2.7
IR Telescope 2:52 1.8 10:00 4.6
In addition to the times shown in the table, the pointing con-straints of the concepts were considered. The gas bearing con-cept imposes a limit on gimbaling freedom of ± 0.0436 radian(± 2.5 deg) about all axes. The mechanical gimbal provides± 0.262 radian (± 15 deg) about the Y-axis and ± 0.833 radian(+± 50 deg) about the X-axis. The baseline mission (does not pro-vide on-orbit access to the telescopes) uses the common mount,which provides hemispherical gimbaling freedom, and results inthe operations durations and mission effectiveness shown inTable IV-29.
The hangar/airlock concept also uses the common mount for thetelescope. Thus, for this mode, only the use of the accesscapability reduces time in the mission available for data acquisi-tion. Each access reduces mission effectiveness (from the base-line in Table IV-29) by the values shown in Table IV-28 for eachtelescope. This is also true for the mechanical gimbal concept,although it should be noted that target selection may be con-strained by the gimbaling limits (this is not a severe constraint).
IV-132
Table IV-29 Operations Effectiveness
*Operations time for the gas bearing concept were based on oneaccess to the telescope during the mission, which requires thedurations shown in Table IV-14.
For the gas bearing mode, the gimbaling freedom of only 0.043radian (± 2.5 deg) requires orienting the orbiter toward a sec-ond target unless one object may be viewed throughout an orbit.Viewing constraints and orbit inclination permit continuous viewingfor the photoheliograph and stratoscope III, but not the IR tele-scope. Resulting times available for data collecting observations,and percentages of total mission times, are shown in Table IV-29.(Note the severe reduction for the IR telescope).
From these timelines and operations analyses, it may be observedthat the mechanical gimbal and hangar accommodations modes aresatisfactory for all telescopes, and that the gas bearing may beused for the photoheliograph and stratoscope III. Using the gasbearing with the IR telescope reduces mission effectiveness toonly 32.6%.
7. Cost Analyses
Estimates were made of the costs of accommodating the photo-heliograph and stratoscope III using the baseline, hangar/airlock,gas bearing, and mechanical gimbal concepts. Table IV-30 is asummary of the DDT&E and production costs for each of the con-cepts investigated.
IV-133
BASELINE MISSION GAS BEARING
Operations OperationsTime, hr: Percent of Time, hr: Percent of
TELESCOPE min Mission min* Mission
Photoheliograph 123:36 74 122:38 73.4
Stratoscope III 143:00 86 131:26 79
IR Telescope 128:23 77 54:28 32.6
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34
The DDT&E costs were based on: design, development, test, andevaluation of the flight hardware; design, development, andproduction of GSE; systems support and integration; programmanagement; and a profit of 10%. Cost estimating relationships(CERs) were used to derive the flight.hardware design, develop-ment, test, and evaluation costs. This cost then served as thereference cost to which cost ratios were applied to determinethe costs for GSE, system support and integration, program manage-ment, and profit. The cost ratios used were:
DDT&EHARDWARE = CER
GSEDESIGN & DEVELOPMENT = (29%) (DDT&EHARDWARE)
GSEPRODUCTION = (116%) (t ARTICLEHARDWARE)
SYSTEM SUPPORT= (11%) (DDT&EHARDWARE + GSEDESIGN & DEVELOPMENT
+ GSEPRODUCTION)
PROGRAM MANAGEMENT = (7.6%) (SYS SUPPORT + DDT&EHARDW ARE
+ GSEDESIGN & DEVELOPMENT + GSEPRODUCTION)
PROFIT =, 10% (PROG MANAGEMENT + SYS SUPPORT + DDT&EHARDWARE
+ GSEDESIGN & DEVELOPMENT +GSEPRODUCTION)
DDT&EToTAL
= PROFIT + PROG MGMT + SYS SUPPORT + DDT&EHARDWARE
+ GSEDESIGN & DEVELOPMENT + GSEPRODUCTION
The production costs were based on developing the cost of thefirst article using CERs, and then using this cost as a referencefor determining the cost of spares, system support and integration,program management, and profit using cost ratios. The cost ratiosused were:
IV-135
First article = CER
Initial spares = (5%) (1St article)
System support = (11%) (1st article + initial spares)
Program management = (7.6%) (system support + first article+ initial spares)
Production costs = profit + program management + system support+ initial spares + first article.
The costs shown for the baseline concept reflect the cost of thecommon telescope mount and the inertial reference system. Thecommon telescope mount is made up of a wide-angle azimuth andelevation mount that provides hemispherical viewing for thetelescopes, and a three-axis gimbal with flexible pivots thatprovides-the fine pointing and stabilization for the stratoscopeIII and photoheliograph. The reference assembly is made up ofa set of four precision star trackers and an inertial measure-ment unit that provides the inertial reference to the telescopemount.
The hangar/airlock costs include the same common telescope mountand reference assembly required for the baseline concept, and theadditional costs associated with providing a pressurized environ-ment for the on-orbit access to the stratoscope III and photo-heliograph. The Sortie Lab modification is the cost to developthe new aft bulkhead and support structure necessary for the air-lock. The cost of the expandable airlock includes the airlockstructure, insulation, meteoroid protection, mechanisms, seals,pressurization system, and monitoring equipment. The instrumentmodification is the cost to provide a pressure shell that en-closes the instrument section of the telescopes. The productioncosts shown are for the production of two instrument pressureshells, one for the stratoscope III and one for the photoheliograph.
The gas bearing concept costs include the reference assembly andSortie Lab modifications required for the hangar/airlock concept,plus the cost of the gas bearing system and the additional costs forthe tertiary mirror that brings the light into the Sortie Lab.The gas bearing costs include the gas bearing, actuators, supportstructure, airlock, gas system, and gas scavenging system. Thecost of installing the tertiary mirror on both the stratoscope IIIand the photoheliograph are included under instrument modifications.
IV-136
The mechanical gimbal concept is a variation of the baselineconcept, where a limited wide-angle gimbal provides 2 degrees ofcoarse pointing for the telescopes. Fine pointing and stabiliza-tion for the telescopes is provided by IMC systems that areinternal to the telescopes. The reference assembly and SortieLab modifications required for the hangar/airlock are also requiredfor the mechanical gimbal concept. The costs shown reflect atertiary mirror and an IMC system for the stratoscope III and thephotoheliograph.
In comparing the costs of the four alternatives, it is importantto note that the costs shown only reflect the cost of the pointingand stabilization systems necessary for the telescopes, and thecosts associated with providing the on-orbit access. The costsof the telescopes themselves and the additional ancillary hardwarenecessary to outfit a payload are not included.
It is also important to note that the baseline costs will alwaysbe required, since the IR telescope and the additional solarinstruments will still require this equipment. The cost of pro-viding the on-orbit access for the photoheliograph and- strato-scope III will be a delta cost to this baseline.
8. Recommended Approach
The recommended concept, selected at a coordination meeting heldat MSFC in February 1973, is the mechanical gimbal concept.
A comparison was made between the three on-orbit access conceptsin order to select the recommended concept for the AstronomySortie Missions. The major characteristics that were considered,and the evaluation of each concept, are presented in Table IV-31.The baseline concept, which does not provide on-orbit access, isincluded for comparative purposes.
The airlock/hangar concept was eliminated first, for severalreasons. The most important of these was that only periodicaccess can be provided to the instruments, while the other twoconcepts provide continuous access for monitoring and adjustments.In addition, costs associated with this concept are considerablyhigher than for the competing concepts. Favorable characteristicsinclude the capability of using the short Sortie Lab, and the useof the hemispherical viewing mount; however, these features areoutweighed by the negative factors mentioned above.
IV-137
The comparison between the gas bearing and mechanical gimbal con-cepts shows significant differences in weights and costs, withthe mechanical gimbal being superior in both respects. The largergimbaling capability of the mechanical gimbal allows faster slewingrates than the Shuttle can provide, thus affording greater flexi-bility of on-orbit operations. The impact on the telescope isgreater for the mechanical gimbal concept since the fold mirrorsmust be articulated, and the IMC system must operate over a largerrange.
The mechanical gimbal concept was selected over the gasbecause of its lower costs and weights, and its greaterflexibility.
bearingoperational
TabZe IV-31 Concept Conrparison
CONCEPTS
Characteristic Gas Bearing Airlock/Hangar Mechanical Gimbal Baseline
Access toInstruments Continuous Periodic Continuous None
Weight, kg(lb)
Photoheliograph 10272.9 (22648) 10100.4 (22268) 9468.2 (20874) 9220.1 (20327)
Spectroscope III 10982.3 (24212) 10879.3 (23985) 10180.8 (22445) 10014.2 (22078)
Costs ($106)
DDT&E 35.723 50.448 23.970 26.019
Production 5.877 8.993 4.221 4.287
Impact on Long Sortie Short Sortie Long Sortie Short SortieSortie Lab Lab Lab Lab Lab
Impact onTelescope Moderate Moderate Greatest --
OperationalEfficiency,%
PHG 73.4 71.5 73.4 74
SIII 79.0 83.3 79.0 86
IV-138
D. PLANNING DATA
1. Project Planning Requirements
The gross planning requirments for the Astronomy Sortie MissionsDefinition Study (Ref 1) are generally unchanged as a result ofthe follow-on study, except for specific numbers of flights ofthe payloads and the project duration of 12 years. In the studycontinuation, a flight sequence and the project duration werenot defined, but program development requirements were generatedfor individual telescopes, arrays, and instruments. The objectivewas to provide planning data that could be used to synthesizeworkable programs having a range of durations, numbers of flights,and costs.
Although the number of flights and program duration were notspecified, several flights of each instrument over a period ofyears are required to satisfy scientific objectives. Therefore,the gross planning requirements presented previously for engineer-ing and manufacturing, testing, quality and reliability assurance,facilities, and project management are applicable to any programderived from the elements analyzed in this study.
2. Schedules
Schedules were developed for each flight hardware item that wasconsidered for inclusion in an Astronomy Sortie Mission project.Since no program was defined, a program milestone schedule andnetwork could not be developed. However, it is expected thatwhen the program is defined, consideration of long-lead-time (LLT)developments and early authorization-to-proceed will be requiredfor some elements.
a. Operations - The operational concept involving the PayloadIntegration Center (PIC) at MSFC, a Launch and Landing Site, anda Space Astronomy Control Facility (SACF) remains the baseline.Initial integration and test for the first flight is expected torequire 6 months. The totally integrated Sortie Lab and pallet,with subsystems and experiments, is then loaded in the SuperGuppy and delivered to the launch site for 3 months of prelaunchoperation before the first flight.
IV-139
For the small UV instruments to be flown on flights when the op-portunity is presented, integration into the Sortie Lab and pallet
may not be required. Prelaunch operations for these instrumentswill be compatible with the primary payload.
The payloads are returned after each flight. The Sortie Lab andpallet with experiments are then shipped to the PIC at MSFC, the
experiments are removed, and the Sortie Lab and pallet are re-furbished to receive other experiments. After integration of
other telescopes and arrays, the integrated payload is tested,accepted as flight-ready, and shipped to the launch site for an-
other sortie flight.
Flight payloads do not pass through the SACF. Film exposed inmissions will be returned to the SACF for processing, and scien-
tists there must have communications with the flight (throughmission control), but the payload handling capabilities necessary
at the PIC-MSFC and at the launch site are not needed at SACF.The overall turnaround time for a payload to complete one flight
and be refurbished for another flight is 10 weeks (Fig. IV-44).
Weeks
O 1 2 3 4 5 6 7 8 9 10*I I i , . . I I I I
42 Pack, Transfer to Launch Site
55 Receipt to Load in Orbiter
12 ] Load in Orbiter and Verify
123 Prelaunch Operations
Launch
Numbers X 168 X Ascent, On-Orbit, DeorbitAre WorkHours Land
22 Safe, Remove Payload, Service
Pack, Transfer to Payload 54 PIC at MSFC
Refurbish and Test 135
Figure IV-44 Turnaround Schedule
IV-140
b. Facilities and GSE - The primary facilities to be activatedare the PIC-MSFC, the launch and landing site, and the SACF.These facilities will be outfitted with GSE developed concurrentlywith the flight hardware by the hardware manufacturer, and pay-load handling and support equipment GSE developed by the programcontractor.
Lead times for activation of facilities, GSE development, andtelescopes and arrays support equipment are:
Lead Time, MonthsItem before First Flight
Facilities
PIC at MSFC 9
Launch and Landing Site 3
SACF 3
GSE Development
Payload Support Equipment 9
Delivery to MSFC 6
Delivery to Launch Site 6
Delivery to SACF 6
Telescopes and Arrays Support Equipment 3 to 6
c. Telescopes, Arrays, and UV Instruments - Experiment hardwarewill be built using a single article approach, in which the flightunit will be subjected to any development tests, and then be mod-ified and serviced for flight acceptance. It is expected thateach telescope and array will be individually contracted for byNASA and provided to this project.
Development and production of experiment hardware should be sched-uled so that delivery will lead expected first flight by 9 months(Fig. IV-45).
IV-141
TELESCOPE, ARRAY, YearOR INSTRUMENT 5 4- 3 2 2 0
_...
First Flight
IR Telescope
Monitor (IR)
Stratoscope III
Photoheliograph
Solar Group
X-Ray Telescope
XUV Spectroheliograph
Coronagraphs
Monitor (Solar Group)
Wide Coverage X-Ray Detector (Group A)
Narrow Band Spect/Polar (Group B)
X-Ray Spectrometer | (Group C)
Low Bkgrnd X-Ray Detector I
Large Modulation Collimator (Group D)
Large Area X-Ray Detector
Collimated PC Spectrometer (Group E)
Six-Inch Survey Cameras (Tifft)
All-Reflective Spect (Morton)
10-cm Image Conv Spect (Carruthers)
15-cm Image Conv Spect (Carruthers)
40-cm Internal Grating Spect (Carruthers)
40-cm Imaging Camera (Carruthers)
Echelle Spectrograph (Morton)
Scanning Spectrometer (Kondo)
C
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Figure IV-45 ScheduZe for TeZescopes, Arrays and Instruments
IV-142
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d. Subsystems for Intermediate Telescopes and Arrays - Test andflight articles are required for the pointing and control, struc-tures, electronic, and thermal control subsystems of the SortieLab and pallet.
The test article will be used by the project contractor to conductsystems tests (structural loading, dynamic, natural environments,and limited function) at the contractor's plant. The flight unitwill undergo preintegration acceptance tests before shipment tothe PIC-MSFC, where integration with the Sortie Lab and pallet is
performed. The pointing and control system is the long-lead timedevelopment item (Fig. IV-46) requiring 3¼ years for DDT&E andinitial production. All subsystems must be delivered 1 year be-fore first flight launch date.
Figure IV-46 ScheduZe for Subsystems Used withIntermediate TeZescopes and Arrays
IV-143
YearSUBSYSTEM 5 3 2 1 0
Launch First Flight
System Preintegration Acceptance at Contractor
Integration & Launch Operations
Pointing & Control System
CMG
Actuators
Reference Assembly
Structures
Common Mount Assembly
Telescope Gimbal Assembly
Support Equipment Set V
Array Platform Assembly
Solar Telescope Housing Assembly
Electronic
Controls & Displays
Electrical Distribution
Data System
Thermal Control
Coating
Insulation
Legend:
O ATP
V Delivery Flight Hardware
DDT&E
'-d Production
e. Subsystems for SmaZZ UV Instruments - Test and flight articlesare required for the pointing and control, structures, and elec-tronic subsystems of the small UV instruments on flights of op-
portunity. Integration of the flight hardware into the carriermay be at the launch site or at a prime payload integrating con-tractor's facility. Schedules are shown in Figure IV-47; theperiod provided for integration of the subsystems is 6 months
for the first flight.
YEAR
SUBSYSTEM 2 1
Launch 1st Flight
Integration & Launch Operations
Pointing & Control
UV Gyro Package (M 7
UV Outer Gimbal Mount Actuators o wm w~7
UV Reference Assembly 7
Structures
UV Instrument Mount Assembly v p7
UV Instrument Platform
Electronic
UV Instrument Control & Display (7
UV Instrument Electrical - 7
UV Instrument Data ( -7
Legend:
O ATP
V Delivery Flight Hardware
DDT&E
r/~/~ Production
Figure IV-47 Schedule for Subsystems Used withSmall UV Instruments
3. Supporting Research and Technology
No supporting research and technology (SRT), in addition to thatidentified in Reference 1, was identified in this continuing study.It should be noted that the schedules shown in the final reportfor SRT items were based on the flight schedule requirements ofthe project as then defined. Changes in the project schedules(to be generated with consideration for the results of this follow-on study) will effect the SRT schedule dates. The span timesshown for these items in the final report have not been revised.
IV-144
V. COST ESTIMATES
A. COSTING APPROACH, METHODOLOGY, AND RATIONALE
The estimates provided in this section describe the cost of devel-opment, first unit production, and operational support for variousitems of hardware and operations that are candidates for an Astron-omy Sortie Mission Project. The estimates do not describe thetotal cost of a particular program of numerous flights, since nospecific program has been defined. Costs are presented in a mannerthat permits combining elements in desired variations to synthesizeastronomy programs for Shuttle sortie missions.
Parametric costing was used to develop the prices for hardware andoperations using cost estimating relationships (CERs), cost ratios,and factors in conjunction with detailed estimates. The CERs, costratios, and factors used were developed from Martin Marietta's ex-perience on programs such as Skylab and Viking, and from data bankinformation of other industry programs. Generally, subsystem hard-ware was priced using CERs; the telescopes, arrays, and instrumentswere priced using detailed estimates and CERs. Software elements,such as project management and systems support, were estimated usingcost ratios and factors. Operations elements were priced usingdetailed estimates and CERs. Estimates were derived for design,development, test and evaluation (DDT&E), first unit production,and operations.
A Work Breakdown Structure (WBS) dictionary, listing in numericalorder the elements considered for this project [as defined by thedata requirements (DR)], is presented in Section E. The elementsof the WBS were followed to present costs for non-recurring (DDT&E),recurring (production), and recurring (operations) (Section B).Floating items costs are included in these prices at the level atwhich they occur. The costs are not accumulated to a project level,since no project is defined and, therefore, there is no basis forcombination.
The general ground rules and assumptions used in this pricing are:
1) Fiscal 1973 constant dollars;
2) Sortie Lab and pallet are GFE -- costs are not included herein;
V-1
3) Shuttle operations cost is excluded. The proton flux detectorwas considered part of Shuttle operations cost;
4) NASA center cost is excluded -- costs are not included herein;
5) Experiment and subsystem configurations remain fixed;
6) No contingencies or discounts are included;
7) Costs include 10% profit;
8) Facility costs are excluded, as they are program dependent,and estimates provided do not describe a program;
9) A prime contractor will be responsible for developing thesubsystems, producing them, and supporting their integrationinto the Sortie Lab and pallet at the PIC at MSFC;
10) The telescopes, arrays, and instruments will be developed byvarious industry or university associate contractors and willbe purchased directly by NASA.
B. COST ESTIMATES BY WBS ELEMENT
Estimates were prepared for non-recurring (DDT&E) on Cost DataForm - A(1), for recurring (production) on Cost Data Form - A(2),and for recurring (operations) on Cost Data Form - A(3). (Theforms are included in this section.)
1. Cost Data Form - A(1) Non-Recurring (DDT&E)
A description of the contents of each column of Form - A(l) follows:
1) Identification Number - The 13-digit WBS number of the item ofcost;
2) WBS Identification - The alphanumeric nomenclature of the itemfrom the WBS;
3) WBS Level - The level at which the element is carried;
4) Expected Cost - The cost estimate for the WBS item;
V-2
5) Confidence Rating - A value of 1 through 4 representing theestimator's confidence in the estimate shown in the WBS itemcost column. The values were obtained by reviewing the cri-teria presented in the Phase A and Phase B Studies Cost Esti-mates Document, DRD No. MF-030A, Table 1, "Confidence LevelGroups for Cost Estimates", and selecting the value mostapplicable;
6) Td - The duration of the costs of the DDT&E activity in months;
7) T - The lead time (months) measured from the start of cost
accrual for the item to the launch milestone;
8) Spread Function - An index number representing a cost distri-bution curve that the estimator recommends for the time phas-ing of costs over Td. Standard distributions from Figure 8
of the Phase A and Phase B Studies Cost Estimates Document,No. MF-030A, were used where applicable.
a. DDT&E for TeZescopes, Arrays, and Instruments - Development ofeach telescope, array, and instrument requires producing one unit,using the unit for tests and checkouts, and updating the same unitfor flight. Costs for these items are provided herein for infor-mation purposes, and do not include prime contractor G & A or profitin addition to that of the associate contractors, since they willbe procured directly by NASA.
"Expected Cost" (column 4) of Form - A(1) for telescopes, arrays,and instruments is defined as follows:
EXPECTED COST = BASIC DDT&E + SYSTEMS SUPPORT.
Where:
BASIC DDT&E includes engineering analyses, design specificationpreparation, tooling design and fabrication, test criteria, devel-opment tests, but no test article; and SYSTEMS SUPPORT = 11% OFBASIC DDT&E.
b. DDT&E for Subsystems - The subsystems consist of assembliesand components that are either qualified and are available off theshelf, or require extensive development. The inherent projectadvantages of returning each payload from orbit, and the relativelyshort on-orbit stay times for each flight, were considered in
V-3
recommending limited development tests requiring only one testarticle for the subsystems. Conformance with orbiter certifi-cation level requirements for qualification is planned for all"critical" components.
"Expected Cost" (column 4) of Form - A(1) for the subsystemsassemblies and components is defined as follows:
EXPECTED COST = A + B + C + D.
Where:
A = BASIC DDT&E + DEVELOPMENT UNIT + GSE DEVELOPMENT, and BASICDDT&E includes engineering enalysis, design, specificationpreparation, tooling design and fabrication, test criteria, anddevelopment tests;
DEVELOPMENT UNIT is cost of unit used for testing;
GSE DEVELOPMENT is 29% of BASIC DDT&E;
B = GSE PRODUCTION, which was estimated as 116% (first flightunit + initial spares);
C = SYSTEMS SUPPORT, which was estimated as 11% (A + B);
D = PROJECT MANAGEMENT, which was estimated as 7.6% (A + B + C).
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4
2. Cost Data Form - A(2) Recurring (Production)
A description of the contents of each column of Formi- A(2) follows:
1) Identification Number - The 13-digit WBS number of the itemof cost;
2) WBS Identification - The alphanumeric nomenclature of theitem from the WBS;
3) WBS Level - The level at which the element is carried;
4) Number of Units - The quantity of units for each WBS itemused in the production phase for one flight or mission with-out allowance for spares;
5) First Unit Cost T1- The cost to produce theoretical first
item (this is the intercept of the learning curve on a log-log plot);
6) Expected Cost - The cost estimate for the WBS item (refer tospecific definitions for telescopes, arrays, and instrumentsand for subsystems defined in later paragraphs);
7) Reference Unit - The production sequence number of the firstunit that is used in the recurring phase of the program;
8) Confidence Rating - A value of 1 through 4 representing theestimator's confidence in the estimate shown in the WBS itemcost column. The values were obtained by reviewing the cri-teria presented in the Phase A and Phase B Studies Cost Esti-mated Document, DRD No. MF-030A, Table 1, "Confidence LevelGroups for Cost Estimates," and selecting the value mostapplicable;
9) Td
- The duration (months) of costs of producing the flight
article;
10) T - The lead time (months) measured from the start of cost
accrual for the item to the launch milestone;
V-15
11) Spread Function - An index number representing a cost distri-bution curve that the estimator recommends for the time phas-ing of costs over Td. Standard distributions from Figure 8
of the Phase A and Phase B Studies Cost Estimates Document,No. MF-030A, were used where applicable;
12) Learning Index - A numerical index of a learning rate relatedto the recurring cost.
a. Recurring (Production) for Telescopes, Arrays, and Instruments -Each telescope, array, or instrument unit that was produced fordeveloping the item, will be updated as required for flight. The"Expected Cost" shown in Form - A(2) for telescopes, arrays, andinstruments is defined as follows:
EXPECTED COST = BASIC PRODUCTION + INITIAL SPARES + SYSTEMS SUPPORT.
Where:
BASIC PRODUCTION is the first UNIT COST T1
(column 5) for items
requiring one unit (column 4) [for items having more than one unit,BASIC PRODUCTION is derived by applying the Learning Index (column13) to the number of units and the first Unit Cost T1];
INITIAL SPARES is derived by applying a percentage, which is vari-able depending on the item and the expected failure or wearoutrate, to the BASIC PRODUCTION cost;
SYSTEMS SUPPORT = 11% of (BASIC PRODUCTION + INITIAL SPARES).
b. Recurring (Production) for Subsystems - Production costs forthe flight hardware were estimated by using CERs. Costs wereestimated at the component or subassembly level. This was re-quired to enable a consistant application of a 90% improvementcurve due to the variable quantity of each item required for DDT&E,production, and operations (spares). The "Expected Cost" shownin Form - A(2) for the subsystems, assemblies, and components isdefined as follows:
EXPECTED COST = A + B + C.
Where:
A = BASIC PRODUCTION + INITIAL SPARES and BASIC PRODUCTION is thefirst UNIT COST T1 (column 5) for items requiring one unit (column
4). [For items having more than one unit, BASIC PRODUCTION is
V-16
derived by applying the Learning Index (column 13) to the numberof units and the 1st UNIT COST T1];
INITIAL SPARES is derived by applying a percentage, which isvariable depending on the item and the expected failure or wear-out rate, to the BASIC PRODUCTION cost;
B = SYSTEMS SUPPORT, which was estimated as 11% of (A);
and C = PROJECT MANAGEMENT, which was estimated as 7.6% of (A +B).
V-17
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V-24
3. Cost Data Form - A(3) Recurring (Operations)
Two distinct elements of costs make up recurring (operations)costs: spares hardware; and operations manpower. The two havebeen estimated separately and are presented on Form - A(3) inseparate groups.
The first element (spares) has been estimated for each candidateWBS item for a payload, WBS Identification Numbers 01-001-05-XX-XX-XX [pp 1 thru 7 of Form - A(3)].
The second element (operations) consists of four separate opera-tions activities [pp 8 thru 11 of Form - A(3)] --
WBS Identification WBS Identification Number
Launch 01-001-06-XX-XX-XX
Mission 01-001-07-XX-XX-XX
Support 01-001-08-XX-XX-XX
Recovery & Refurbishment 01-001-09-XX-XX-XX
A description of the contents of each column of Form - A(3) follows:
1) Identification Number - The 13-digit WBS number of the itemof cost;
2) WBS Identification - The alphanumeric nomenclature of the itemfrom the WBS;
3) WBS Level - The level at which the element is carried;
4) Number of Units - The quantity of units is not shown. A per-centage (variable depending on the item and the expected failureor wearout rate) was applied to the production quantity toestimate spares per flight or mission (pp 1 thru 7). A quantityis not applicable for operations costs shown on pages 8 through11;
5) Expected Cost - The cost estimate for the WBS item;
6) Reference Unit - Not applicable;
7) Reference Unit Cost - Not applicable;
V-25
8) Confidence Rating - A value of 1 through 4 representing theestimator's confidence in the estimate shown in WBS item costcolumn. The values were obtained by reviewing the criteriapresented in the Phase A and Phase B Studies Cost EstimatesDocument, DRD No. MF-030A, Table 1, "Confidence Level Groupsfor Cost Estimates," selecting the value most applicable;
9) Td - For recurring (operations), Td is undefined since no
flight schedule is defined;
10) T - For recurring (operations), Ts is undefined since no
flight schedule is defined;
11) Spread Function - An index number representing a cost distri-bution curve that the estimator recommends for the time phas-ing of costs over Td. Standard distributions from Figure 8
of the Phase A and Phase B Studies Cost Estimates Document,No. MF-030A, were used where applicable;
12) Learning Index - A numerical index of a learning rate relatedto the recurring cost.
a. Recurring (Operations) Hardware Spares (Pages 1 through 7) -Costs of spares for hardware elements for one flight or missionwere estimated. The "Expected Cost" shown in Form - A(3) forspares is defined as follows:
EXPECTED COST = A + B + C + D.
Where:
A = SPARES HARDWARE, which was estimated as a percentage, depend-ing on the item and the expected failure or wear-out rate, andapplied to the FIRST UNIT COST T1;
B = GSE, which was estimated as 25% of (A);
C = SUBSYSTEM SUPPORT, which was estimated as 11% of (A + B);
D = PROJECT MANAGEMENT, which was estimated as 7.6% of (A + B + C).
b. Recurring (Operations) for Launch Mission Support, and Recoveryand Refurbishment (pp 8 thru 11) - Cost Estimates for the launch,mission, support, and recovery/refurbishment functions were basedon the engineering, technical, and scientific manpower levels
V-26
required to perform these operations. Associated (but still direct)support from laboratory personnel at the PIC and for interfacingwith the Sortie Lab and pallet are included. Total direct man-power requirements were then adjusted for systems support and proj-ect management.
These cost estimates assume the functions are accomplished by thecontractor at three separate locations. They further assume thatgovernment facilities such as buildings, utilities, office equip-ment, transportation, gases, and fluids are available at eachlocation at no cost to the contractor.
"Expected Cost" as shown in Form - A(3) for these operations isdefined as follows:
EXPECTED COST = A + B + C.
Where:
A = TOTAL DIRECT MANPOWER COST, which was estimated as the man-months required considering skill levels;
B = SYSTEMS SUPPORT, which was estimated as 11% of (A);
C = PROJECT MANAGEMENT, which was estimated as 7.6% of (A + B).
V-27
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C. COST ESTIMATES BY COMMON ITEMS, TELESCOPE, ARRAY GROUPS, ANDINSTRUMENTS
Costs for DDT&E, recurring (production), and recurring (operations)have been tabulated for each WBS item, grouping the items accordingto application. In this grouping by use, the intermediate tele-scopes and array groups with their subsystems are considered sepa-rately from the small UV instruments and the subsystems requiredfor them. The small UV instruments are unique and may be flownseparately when an opportunity is available.
For the intermediate class telescopes and the array groups, itemsof subsystems equipment having some common use were assigned toone of four categories of commonality. These categories includeportions of the various subsystems, but have the distinguishingcharacteristic of some use that is common to the entire payload,or to two or more Astronomy Sortie telescopes, or to telescopesand array groups, or to two or more array groups. Subsystemequipment that has no common application (i.e., is peculiar toa particular telescope or to one array group) was excluded fromany of the "common" categories. The four "common" categories forthe intermediate telescopes and array groups are as follows:
Category Description
I Common to Payloads
II - Common to Telescopes and Arrays
III Common to Telescopes
IV Common to Array Groups B, C, D,&E
The equipment lists, with costs, are shown in Table V-1 for thesecategories of common equipment.
The subsystem components and assemblies included in each of the"common" categories for the telescopes and arrays are as follows:
1) Category I - Equipment "Common to Payloads" includes the sub-system components and assemblies that must be provided in aSortie Lab and pallet for any astronomy sortie mission. Thisequipment supports either an intermediate telescope or anarray group or both. The distinguishing characteristic is
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that the equipment is necessary for any of the AstronomySortie missions, regardless of telescope and array groupcomplement.
2) Category II - "Common to Telescopes and Arrays" includesthe subsystem equipment that must be provided for each in-termediate telescope and/or for each array group. Thisequipment is mainly the common mount and peripherial equip-ment. Equipment that is specifically for the telescopes isexcluded here and is Category III. Equipment that is spe-cifically for the array groups is also excluded and isCategory IV.
3) Category III - "Common to Telescopes" includes the subsystemequipment that must be provided for any of the intermediatetelescopes (or the solar group). This is mainly equipmentused on, or in conjunction with, the common mount to accom-modate the telescopes or to support telescope operations.Specifically excluded from this "common" grouping is equip-ment required only for array groups B, C, D, and E.
4) Category IV - "Common to Array Groups B, C, D, and E" includesthe subsystem equipment that must be provided for any of thearray groups B, C, D, and E. This is mainly equipment usedon, or in conjunction with, the common mount to accommodatethe array groups, or to support array operations. ArrayGroup A equipment is not common to the other array groupsand is listed separately on summary sheets as peculiarequipment.
Subsystem equipment for the small UV instruments is common to allof the instruments considered and to the two groups of these in-struments. The two groups are the three ST-100 instruments (GroupA) and the two complementary 40-centimeter Carruthers instruments(Group B). The common equipment for these instruments was groupedas Category V "Common to UV Instruments." The equipment lists,with costs, are shown in Table V-2.
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Tabulations were then prepared from the appropriate "common" cate-gories of equipment (adding any peculiar equipment) for each in-termediate telescope, each array group, and the UV instruments.In these tabulations, the costs for each telescope, array group,and UV instrument are summarized, and the detail of DDT&E, Recurring
(Production), and Recurring (Operations) is presented. For the
total cost of the first flight, DDT&E plus Recurring (Production)are required. For follow-on flights only Recurring (Operations)costs are incurred.
These data are presented in Table V-3 for the intermediate tele-
scopes, Table V-4 for the array groups, and Table V-5 for the smallUV instruments.
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D. TECHNICAL CHARACTERISTICS DATA
Technical Characteristics Data (TCD) Form B sheets were preparedfor each of the telescopes, arrays, and subsystems at WBS levels5 and 6. The TCDs are arranged here in order of increasing num-ber. The data were used in estimating costs of each item. Adescription of the information in each column of the TCD form isas follows:
1) WBS Identification Number - The 13-digit WBS code number ofthe item;
2) WBS Identification - The alphanumeric nomenclature of the itemfrom the WBS;
3) Quantity or Value - The numerical quantity or value of the char-acteristic (Column 5) under consideration. Where no character-istic is identified, quantity refers to the WBS item and iden-tification number;
4) Units of Measure - The identification of the units of measureassociated with the characteristics (Column 5) under consid-eration. Where no characteristic is identified, units ofmeasure applies to the WBS item and identification number;
5) Characteristics - The identification of the technical propertyunder consideration;
6) Notes - Comments or explanations to clarify any of the informa-tion presented.
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12
E. WORK BREAKDOWN STRUCTURE (WBS) DICTIONARY AND DIAGRAM
The final WBS dictionary and diagram are presented in this section.This WBS is compatible with the breakdown that was used in esti-mating costs.
The dictionary (Table V-6) is a listing in numerical order of thetelescopes, arrays, and instruments that are candidates for anAstronomy Sortie Missions project. Included with these items arethe subsystems and their components (to level 7 when appropriate)necessary for the flight of the telescopes, arrays, and instruments.
Figure V-1 shows the relationship of these elements in a familytree. This diagram shows the payload (a level 4 item) elementsto level 5 and level 6. Operations, including Launch, Mission,Support, and Recovery and Refurbishment (level 4 items) are shownto level 5.
Project elements at level 4, other than payloads and operations,are shown on the diagram, but were not priced. These level 4items (including Project Management, System Support and Integra-tion, Facilities, and GSE) depend on the number of payloads,number of flights, and duration of the project, which were notselected for this study.
V-113
Table V-6 WBS Dictionary
WBS IDENTIFICATION NUMBER
LEVEL 21 LEVEL 3 LEVEL 4 LEVEL 5 LEVEL 6 LEVEL 7 WBS IDENTIFICATION
0 1 0 0 0 0 0 0 0 0 0 0 0 SHUTTLE SORTIE MISSION PROGRAM
O 1 0 0 1 0 0 0 0 0 0 0 0 ASTRONOMY SORTIE MISSION PROJECT
O 1 0 O 1 0 1 0 0 0 0 0 0 PROJECT MANAGEMENT
0 1 0 O 1 0 1 0 1 0 0 0 0 PROGRAM CONTROL
O 1 0 O 1 0 1 0 2 0 0 0 0 CONFIGURATION MANAGEMENT
O 1 0 0 1 0 1 0 3 0 0 0 0 CONTRACTUAL DATA MANAGEMENT
O 1 0 0 1 0 2 0 0 0 0 0 0 SYSTEM SUPPORT & INTEGRATION
O 1 0 O 1 0 2 0 1 0 0 0 0 SYSTEMS ANALYSIS
O 1 0 O 1 0 2 0 2 0 0 0 0 PAYLOAD INTEGRATION
O 1 0 O 1 0 2 0 3 0 0 0 0 PROGRAM INTEGRATION
O 1 0 O 1 0 2 0 4 0 0 0 0 SAFETY AND RELIABILITY
0 1 0 O 1 0 2 0 5 0 0 0 0 QUALITY ASSURANCE
O 1 0 O 1 0 3 0 0 0 0 0 0 FACILITIES
O 1 0 O 1 0 3 0 1 0 0 0 0 CONTRACTOR
O 1 0 O 1 0 3 0 2 0 0 0 0 GOVERNMENT
O 1 0 O 1 0 4 0 0 0 0 0 0 GROUND SUPPORT EQUIPMENT
0 1 0 O 1 0 4 0 1 0 0 0 0 ELECTRICAL & ELECTRONIC
O 1 0 O 1 0 4 0 2 0 0 0 0 STRUCTURAL & MECHANICAL
0 1 0 O 1 0 4 0 3 0 0 0 0 OPTICAL
O 1 0 O 1 0 4 0 4 0 0 0 0 TRANSPORT & HANDLING
O 1 0 0 1 0 5 0 0 0 0 0 0 PAYLOADS
O 1 0 O 1 0 5 0 1 0 0 0 0 TELESCOPES
O 1 0 O 1 0 5 0 1 0 1 0 0 IR TELESCOPE
O 1 0 0 1 0 5 0 1 0 2 0 0 STRATOSCOPE III
O 1 0 O 1 0 5 0 1 0 3 0 0 PHOTOHELIOGRAPH
O 1 0 O 1 0 5 0 1 0 4 0 0 X-RAY TELESCOPE
O 1 0 O 1 0 5 0 1 0 5 0 0 XUV SPECTROHELIOGRAPH
O 1 0 O 1 0 5 0 1 0 6 0 0 CORONAGRAPHS
O 1 0 O 1 0 5 0 1 0 7 0 0 MONITORS
O 1 0 O 1 0 5 0 2 0 0 0 0 ARRAYS
O 1 0 0 1 0 5 0 2 0 1 0 0 LARGE AREA X-RAY DETECTOR
O 1 0 0 1 0 5 0 2 0 2 0 0 WIDE COVERAGE X-RAY DETECTOR
O 1 0 0 1 0 5 0 2 0 3 0 0 LARGE MODULATION COLLIMATOR
O 1 0 O 1 0 5 0 2 0 4 0 0 NARROW BAND SPECTRO/POLARIM
O 1 0 O 1 0 5 0 2 0 5 0 0 COLLIMATED PC SPECTROMETER
O 1 0 O 1 0 5 0 2 0 6 0 0 GAMMA-RAY SPECTROMETER
O 1 0 0 1 0 5 0 2 0 7 0 0 LOW BACKGROUND y-RAY DETECTOR
O 1 0 O 1 0 5 0 2 0 8 0 0 PROTON FLUX DETECTOR
V-114
TabZe V-6 (cont)
WBS IDENTIFICATION NUMBER
LEVEL 2 LEVEL 3 LEVEL 4 LEVEL 5 LEVEL 6 LEVEL 7 WBS IDENTIFICATION
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 1
0 0 1
001
001
001
001
001
001
001
001
001
001
001
001
001
001
001
001
001
001
001
001
001
001
001
001
001
001
001
001
001
001
001
0 0 1
001
001
0 0 1
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
1 0 0 1
5
5
5
5
5
5
5
5
5
5
5
5
5
5
5
5
5
5
5
5
5
5
5
5
5
5
5
5
5
5
5
5
5
5
5
5
0 3
O 3
O 3
0 3
0 3
O 3
0 3
O 3
0 3
O 3
O 3
0 3
0 3
0 3
0 3
0 3
0 3
0 3
0 3
O 3
0 3
0 3
0 3
0 3
0 3
0 3
0 3
0 3
0 3
0 3
0 4
0 4
0 4
O 4
O 4
0 4
5 10
0 0
0 1
0 1
0 1
0 1
O 2
O 2
0 2
0 3
O 3
O 3
O 3
O 3
O 4
O 4
O 5
0 5
O 5
0 5
O 5
0 5
O 6
O 7
O 7
O 7
O 8
O 8
0 8
0 8
O 8
O O
0 1
0 1
0 1
0 1
O i
4 10
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
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0
1 0
0
0
1
2
3
0
1
2
0
1
2
3
4
0
1
0
1
2
3
4
5
0
0
1
2
0
1
2
3
4
0
0
1
2
3
4
5
POINTING & CONTROL SYSTEM
CMG ASSEMBLY
DOUBLE GIMBAL CMGs
INVERTERS
IMU
COMMON MOUNT ACTUATORS
AZIMUTH POINTING
DEPLOYMENT
TELESCOPE GIMBAL ACTUATORS
ELEVATION POINTING & STABILITY
AZIMUTH STABILITY
ROLL
PITCH & YAW (CORONAGRAPHS)
ARRAY PLATFORM ACTUATOR
ELEVATION POINTING
REFERENCE ASSEMBLY
STRAPDOWN STAR TRACKERS
TELESCOPE IMU
FINE SUN SENSOR
BORESIGHTED STAR TRACKER-PRECISION
CORRELATION TRACKER
UV GYRO PACKAGE
UV OUTER GIMBAL MOUNT ACTUATORS
ELEVATION POINTING & STABILITY
AZIMUTH STABILITY
UV REFERENCE ASSEMBLY
VIDICON CAMERA
REFERENCE CAMERA
SUN-EARTH SENSOR
RADIATION DETECTOR
STRUCTURES
COMMON MOUNT ASSEMBLY
AZIMUTH TABLE
AZIMUTH YOKE
DEPLOYMENT YOKE
DEPLOYMENT GEARMOTORS &
LAUNCH LOCKS
JETTISON EQUIPMENT
V-115
0
Table V-6 (cont)
WBS IDENTIFICATION NUMBER
LEVEL 21 LEVEL 3 LEVEL 41 LEVEL 51 LEVEL 6 LEVEL 7 WBS IDENTIFICATION
0 1 0 0 1 0 5 0 4 0 2 0 0 TELESCOPE GIMBAL ASSEMBLY
O 1 0 0 1 0 5 0 4 0 2 0 1 OUTER GIMBAL RING
0 1 0 0 1 0 5 0 4 0 2 0 2 OUTER ROLL RING
0 1 0 0 1 0 5 0 4 0 2 0 3 INNER ROLL RING
0 1 0 0 1 0 5 0 4 0 2 0 4 ROLL GEAR
0 1 0 0 1 0 5 0 4 0 2 0 5 TELESCOPE P&C PLATFORM
0 1 0 0 1 0 5 0 4 0 2 0 6 GIMBAL GEARMOTORS & LAUNCH LOCKS
0 1 0 0 1 0 5 0 4 O0 3 0 0 ARRAY PLATFORM ASSEMBLY
0 1 0 0 1 0 5 0 4 0 3 0 1 ARRAY MOUNT
0 1 0 0 1 0 5 0 4 0 3 0 2 PLATFORM GEARMOTORS & LAUNCH LOCKS
0 1 0 0 1 0 5 0 4 0 4 0 0 SUPPORT EQUIPMENT SET
0 1 0 0 1 0 5 0 4 0 4 0 1 CMG SUPPORT STRUCTURES
0 1 0 0 1 0 5 0 4 0 4 0 2 WC X-RAY DETECTOR MOUNT
0 1 0 0 1 0 5 0 4 0 4 0 3 y-RAY SPECT HOUSING & EXT MECH
0 1 0 0 1 0 5 0 4 0 5 0 0 SOLAR TELESCOPE HOUSING ASSY
0 1 0 0 1 0 5 0 4 0 5 0 1 TUBULAR STRUCTURE
0 1 0 0 1 0 5 0 4 0 5 0 2 BULKHEADS
0 1 0 0 1 0 5 0 4 0 5 0 3 SUNSHIELD-FIBERGLASS
0 1 0 0 1 0 5 0 4 0 5 0 4 APERTURE DOORS
0 1 0 0 1 0 5 0 4 0 5 0 5 DOOR ACTUATORS
0 1 0 0 1 0 5 0 4 0 6 0 0 UV INSTRUMENT MOUNT ASSEMBLY
0 1 0 0 1 0 5 0 4 0 · 6 0 1 AZIMUTH TABLE
0 1 0 0 1 0 5 0 4 0 6 0 2 AZIMUTH YOKE
O 1 0 0 1 0 5 0 4 0 6 0 3 LAUNCH LOCKS
0 1 0 0 1 0 5 0 4 0 6 0 4 JETTISON EQUIPMENT
O 1 0 0 1 0 5 0 4 0 6 0 5 AZIMUTH GEARMOTORS
0 1 0 0 1 0 5 0 4 0 6 0 6 FITTINGS & FIXTURES
0 1 0 0 1 0 5 0 4 0 7 0 0 UV INSTRUMENT PLATFORM
0 1 0 0 1 0 5 0 4 0 7 0 1 EQUIPMENT PLATFORM
0 1 0 0 1 0 5 0 4 0 7 0 2 GIMBAL RING
0 1 0 0 1 0 5 0 4 0 7 0 3 OUTER ROLL RING
0 1 0 0 1 0 5 0 4 0 7 0 4 INNER ROLL RING
0 1 0 0 1 0 5 0 4 0 7 0 5 PLATFORM GEARMOTORS
0 1 0 0 1 0 5 0 4 0 7 0 6 FITTINGS & FIXTURES
0 1 0 0 1 0 5 0 5 0 0 0 0 ELECTRONIC
O 1 0 0 1 0 5 0 5 0 1 0 0 CONTROL & DISPLAY
0 1 0 0 1 0 5 .O 5 0 1 0 1 CB/DISTRIBUTOR PANEL
0 1 0 0 1 O 5 0 5 0 1 0 2 MULTIPURPOSE CRT
V-116
Table V-6 (cont)
WBS IDENTIFICATION NUMBER
LEVEL 2 LEVEL 3 LEVEL 4 LEVEL 5 LEVEL 6 LEVEL 7 WBS IDENTIFICATION
ELECTRONIC (cont)
0 1 0 0 1 0 5 .0 5 0 1 0 3 SYMBOL GENERATOR
O 1 0 0 1 0 5 0 5 0 1 0 4 FUNCTION KEYBOARD
O 1 0 0 1 0 5 0 5 0 1 0 5 ALPHANUMERIC KEYBOARD
0 1 0 0 1 0 5 0 5 0 1 0 6 KEYBOARD ENCODER
0 1 0 0 1 0 5 0 5 0 1 0 7 MICROFILM VIEWER
0 1 0 0 1 0 5 0 5 0 1 0 8 EVENT TIMER
0 1 0 0 1 0 5 0 5 0 1 0 9 MISSION TIMER
0 1 0 0 1 0 5 0 5 0 1 1 0 THREE-AXIS CONTROLLER
0 1 0 0 1 0 5 0 5 0 1 1 1 ANNUNCIATOR BANK
0 1 0 0 1 0 5 0 5 0 1 1 2 RECORDER
0 1 O 0 1 0 5 0 5 0 2 0 0 ELECTRICAL
O 1 0 0 1 0 5 0 5 0 2 0 1 LOAD CENTER SWITCH
0 1 0 0 1 0 5 0 5 0 2 0 2 FEEDER CABLES
0 1 0 0 1 0 5 0 5 0 2 0 3 JUNCTION
0 1 0 0 1 0 5 0 5 0 3 0 0 DATA
0 1 O 0 1 0 5 0 5 0 3 0 1 DATA BUS INTERFACE UNIT
O 1 0 0 1 0 5 0 5 0 3 0 2 COAX DATA BUS
O 1 0 0 1 0 5 0 5 0 3 0 3 PALLET INSTRUMENTATION BOX
0 1 O 0 1 0 5 0 5 0 3 0 4 DATA PROCESSOR
O 1 0 0 1 0 5 0 5 0 4 0 0 UV INSTRUMENT CONTROL & DISPLAY
O 1 0 0 1 0 5 0 5 0 5 0 0 UV INSTRUMENT ELECTRICAL
O 1 0 0 1 0 5 0 5 0 5 0 1 LOAD CENTER SWITCH
O 1 O 0 1 0 5 0 5 0 5 0 2 FEEDER CABLES
O 1 0 O 1 0 5 0 5 0 5 0 3 JUNCTION BOX
O 1 O 0 1 0 5 0 5 0 6 0 0 UV INSTRUMENT DATA
0 1 0 0 1 0 5 0 5 0 6 0 1 TELEMETRY
0 1 0 0 1 0 5 0 5 0 6 0 .2 PROGRAMMER
0 1 0 0 1 0 5 0 5 0 6 0 3 MINI-COMPUTER
O 1 0 0 1 0 5 0 6 0 0 0 0 THERMAL CONTROL
O 1 0 0 1 0 5 0 6 0 1 0 0 THERMAL COATING
0 1 O 0 1 0 5 0 6 0 2 0 0 MULTILAYER INSULATION
O 1 O 0 1 0 5 0 7 0 0 0 0 SMALL UV INSTRUMENTS-
O 1 0 0 1 0 5 0 7 0 1 0 0 SIX-INCH SURVEY CAMERAS (TIFFT)
O 1 0 0 1 0 5 0 7 0 2 0 0 ALL-REFLECTIVE SPECTROGRAPH (MORTON)
0 1 0 0 1 0 5' 0 7 0 3 0 0 10 CM IMAGE CONVERTING SPECTROGRAPH
(CARRUTHERS)
0 1 0 0 1 0 5 0 7 0 4 0 0 15 CM IMAGE CONVERTING SPECTROGRAPH
(CARRUTHERS)
V-117
Table V-6 (cont)
WBS IDENTIFICATION NUMBER
LEVEL 2 LEVEL 3 LEVEL 4 LEVL 51 LEVEL 6 LEVEL 7 WBS IDENTIFICATION I1SMALL UV INSTRUMENTS (cont)
0 1 0 0 1 0 5 0 7 0 5 0 0 40 CM INTERNAL GRATING SPECTROGRAPH(CARRUTHERS)
O 1 O 0 1 0 5 0 7 0 6 0 0 40 CM IMAGING CAMERA (CARRUTHERS)
0 1 0 0 1 0 5 0 7 0 7 0 0 ECHELLE SPECTROGRAPH (MORTON)
0 1 O 0 1 0 5 0 7 0 8 0 0 SCANNING SPECTROMETER (KONDO)
0 1 O 0 1 0 6 0 0 0 0 0 0 LAUNCH OPERATIONS
O 1 O 0 1 0 6 0 1 0 0 0 0 IR TELESCOPE
0 1 O 0 1 0 6 0 2 0 0 0 0 STRATOSCOPE III
0 1 O 0 1 0 6 0 3 0 0 0 0 PHOTOHELIOGRAPH
0 1 O 0 1 0 4 0 0 0 0 0 0 SOLAR GROUP
0 1 0 0 1 0 6 0 5 0 0 0 0 LARGE AREA X-RAY DETECTOR & COLLIMATEDPC SPECT (GROUP E)
0 1 0 0 1 0 6 0 6 0 0 0 0 WIDE COVERAGE X-RAY DETECTOR (GROUP A)
0 1 0 0 1 0 6 0 7 0 0 0 0 LARGE MODULATION COLLIMATOR (GROUP D)
0 1 0 0 1 0 6 0 8 0 O 0 0 NARROW BAND SPECTRO/POLARIM (GROUP B)
O 1 O 0 1 0 6 0 9 0 0 0 0 GAMMA-RAY SPECTROMETER AND LOW BACK-GROUND y-RAY DETECTOR (GROUP C)
0 1 0 0 1 0 6 1 0 0 0 0 UV INSTRUMENT
0 1 O 0' 1 0 6 1 1 0 0 0 0 INTERMEDIATE TELESCOPES SUBSYSTEMS
O 1 O 0 1 0 6 1 2 0 0 0 0 ARRAY GROUPS SUBSYSTEMS
O 1 O 0 1 0 6 1 3 0 0 0 0 UV INSTRUMENT SUBSYSTEMS
0 1 0 0 1 0 7 0 0 0 0 0 0 MISSION OPERATIONS
0 1 0 0 1 0 7 0 1 0 0 0 0 IR TELESCOPE
O 1 O 0 1 0 7 0 2 0 0 0 0 STRATOSCOPE III
0 1 0 0 1 0 7 0 3 0 0 0 0 PHOTOHELIOGRAPH
0 1 O 0 1 O 7 0 4 0 0 0 0 SOLAR GROUP
0 1 0 0 1 0 7 0 5 0 0 0 0 LARGE AREA X-RAY DETECTOR & COLLIMATEDPC SPECT (GROUP E)
0 1 0 0 1 0 7 0 6 0 0 0 0 WIDE COVERAGE X-RAY DETECTOR (GROUP A)
0 1 0 0 1 0 7 0 7 0 0 0 0 LARGE MODULATION COLLIMATOR (GROUP D)
0 1 0 0 1 0 7 0 8 0 0 0 0 NARROW BAND SPECTRO/POLARIM (GROUP B)
0 1 0 0 1 0 7 0 9 0 0 0 0 GAMMA-RAY SPECTROMETER AND LOW BACK-GROUND y-RAY DETECTOR (GROUP C)
0 1 0 0 1 0 7 1 0 0 0 0 0 UV INSTRUMENT
O 1 0 0 1 0 7 1 1 0 0 0 0 INTERMEDIATE TELESCOPES SUBSYSTEMS
0 1 0 0 1 0 7 1 2 0 0 0 0 ARRAY GROUPS SUBSYSTEMS
0 1 0 0 1 0 7 1 3 0 0 0 0 UV INSTRUMENT SUBSYSTEMS
V-118
r-
I
Table V-6 (concZ)
WBS IDENTIFICATION NUMBER
LEVEL 2 LEVEL 3 LEVEL 4 LEVEL 5 LEVEL 6 LEVEL 7 WBS IDENTIFICATION
0 1 0 0 1 0 8 0 0 0 0 0 0 SUPPORT OPERATIONS
0 1 0 0 1 0 8 0 1 0 0 0 0 IR TELESCOPE
0 1 0 0 1 0 8 0 2 0 0 0 0 STRATOSCOPE III
0 1 0 0 1 0 8 0 3 0 0 0 0 PHOTOHELIOGRAPH
0 1 0 0 1 0 8 0 4 0 0 0 0 SOLAR GROUP
0 1 0 0 1 0 8 0 5 0 0 0 0 LARGE AREA X-RAY DETECTOR & COLLIMATEDPC SPECT (GROUP E)
0 1 0 0 1 0 8 0 6 0 0 0 0 WIDE COVERAGE X-RAY DETECTOR (GROUP A)
0 1 0 0 1 0 8 0 7 0 0 0 0 LARGE MODULATION COLLIMATOR (GROUP D)
0 1 0 0 1 0 8 0 8 0 0 0 0 NARROW BAND SPECTRO/POLARIM (GROUP B)
0 1 0 0 1 0 8 0 9 0 0 0 0 GAMMA-RAY SPECTROMETER AND LOW BACK-GROUND y-RAY DETECTOR (GROUP C)
0 1 0 0 1 0 8 1 0 0 0 0 0 UV INSTRUMENT
0 1 0 0 1 0 8 1 1 0 0 0 0 INTERMEDIATE TELESCOPES SUBSYSTEMS
0 1 0 0 1 0 8 1 2 0 0 0 0 ARRAY GROUPS SUBSYSTEMS
0 1 0 0 1 0 8 1 3 0 0 0 0 UV INSTRUMENT SUBSYSTEMS
0 1 0 0 1 0 9 0 0 0 0 0 0 RECOVERY AND REFURBISHMENT OPERATIONS
0 1 0 0 1 0 9 0 1 0 0 0 0 IR TELESCOPE
0 1 0 0 1 0 9 0 2 0 0 0 0 STRATOSCOPE III
0 1 0 0 1 0 9 0 3 0 0 0 0 PHOTOHELIOGRAPH
O 1 0 0 1 0 9 0 4 0 0 0 0 SOLAR GROUP
0 1 0 0 1 0 9 0 5 0 0 0 0 LARGE AREA X-RAY DETECTOR & COLLIMATEDPC SPECT (GROUP E)
0 1 0 0 1 0 9 0 6 0 0 0 0 WIDE COVERAGE X-RAY DETECTOR (GROUP A)
0 1 0 0 1 0 9 0 7 0 0 0 0 LARGE MODULATION COLLIMATOR (GROUP D)
0 1 0 0 1 0 9 0 8 0 0 0 0 NARROW BAND SPECTRO/POLARIM (GROUP B)
O 1 0 0 1 0 9 0 9 0 0 0 0 GAMMA-RAY SPECTROMETER AND LOW BACK-GROUND y-RAY DETECTOR (GROUP C)
0 1 0 0 1 0 9 1 0 0 0 0 0 UV INSTRUMENT
0 1 0 0 1 0 9 1 1 0 0 0 0 INTERMEDIATE TELESCOPES SUBSYSTEMS
0 1 0 0 1 0 9 1 2 0 0 0 0 ARRAY GROUPS SUBSYSTEMS
0 1 0 0 1 0 9 1 3 0 0 0 0 UV INSTRUMENT SUBSYSTEMS
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VI. SUGGESTED ADDITIONAL EFFORT
In addition to the additional effort recommended during the orig-inal 9-month study (Vol I of Ref 1), the effort described belowshould be performed before deciding on a recommended AstronomySortie Mission Program.
Integration of Astronomy Working Group Requirements - The work-ing groups for Astronomy Sortie Missions are presently definingthe objectives and instruments that should be accommodated bythe Shuttle sortie missions. The objectives and instruments be-ing defined by the working groups differ significantly from thoseinstruments considered by this study. It is recommended that ad-ditional effort be expended to convert the scientific objectivesand instrument requirements into engineering requirements thatwould be used to identify the ancillary hardware required, inter-face requirements on the Shuttle, and the operations necessaryto support the instruments. In addition, planning data should bedeveloped for the various working groups to provide NASA with abroader base for planning purposes.
Integration of Sortie Mission Requirements - There are a numberof studies and working groups that are defining the ancillaryhardware, interfaces, and operations that are necessary for aparticular sortie payload. To date, these studies have beenworked independent of each other, and no effort has been made tointegrate the results. It is suggested that a study be initiatedto integrate the results of the sortie working groups and relatedstudy activities, thereby providing definition of the ancillaryhardware, interfaces, and operations that would satisfy the sortiemode of operation in the most cost-efficient manner.
Low-Cost Payload Study - Recognizing the importance of low-cost
payloads for the early sortie missions, it is recommended that astudy be initiated to investigate the possibility of using exist-ing astronomy instruments for the early sortie flights. Thisstudy would survey governmental, institutional, and industrial
sources to determine astronomy instruments that are in existenceand candidates for the sortie mode of operation. The results ofthis study would be: definition of existing instruments for sortiemissions; determination of ancillary hardware necessary to accom-modate the instruments; and definition of the operations required.In addition, detailed planning data should be provided for eachof the candidate instruments and their accommodation requir ments.
VI\1
Contamination Study - Since contamination of astronomy instru-ments is of major concern, it is recommended that a detailedstudy on the cause and effect of contamination be undertaken.The purpose of the study would be to identify the possible sources
of contamination for the sortie mode of operation and to deter-mine the effects of the contamination on the astronomy instru-ments. Typical contamination sources that should be investigatedinclude materials outgassing, atmosphere leakage, overboard dumps
of_water and waste, thruster firings, gas leakages, etc. The ef-fects on the astronomy instruments would include the degradationof the instruments due to absorption, deposition, scattering, etc.
VI-2
VII. REFERENCES
1. Astronomy Sortie Missions Definition Study Final Report.Contract NAS8-28144, Martin Marietta Corp. Denver, Colorado,Sept 1972.
2. Sortie Can Conceptual Design. NASA/MSFC, ASR-PD-DO-72-2,March 1, 1972.
3. Space Shuttle Baseline Accommodations for Payloads, NASA/MSC,MSC-06900, June 27, 1972.
4. Final Technical Report, 6-Inch ApoZZo Ultraviolet Camera.NAS8-20651, June 1, 1971.
5. Reference Earth OrbitaZ Research and Applications Investigation.NHB 7150.1 (Blue Book), Jan 15, 1971.
6. ShuttZe/RAM Deployment Mechanism Conceptual Design. GeneralDynamics Corp, July 28, 1972.
7. Sortie Lab Design Requirements. NASA/MSFC, S&E-S/P-T, Dec 1,1972.
8. Sortie Lab Briefing to Solar Physics Working Group. NASA/MSFC,Nov 14, 1972.
9. Final Report, Expandable Space Structure. AF 33(657)-9733,Martin Marietta Corporation, Denver, Colorado, September 1964.
VII-1