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INFORMATION AS POSSIBLE

https://ntrs.nasa.gov/search.jsp?R=19800018905 2020-04-02T17:56:57+00:00Z

79-FM-23JSC-16026

(Supersedes 78.FM-29)(Supersedes JSC-14235)

STS-7

At Flight Feasibility ►Assessment4_1

(IIASA-TM-81091) SHUTTLE PROGRAM. STS-7 N80­ 2740bFEASIBIL ITY ASSESSMENI: IUS/IDRS - A (NASA)63 p HC A04/MF A01 CSCI 22A

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Mission and Planning Analysis Divisionpi R

July 1979

NnonNational Aeronautics and

Space Administration

-Lyndon B. Johnson Space CenterHouston, Texas

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SHUTTLE PROGRAM

STS -7 Flight Feasibility Assessment

IUS/TDRS -A

By Flight 'P lanning Branch

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ApprovedEdgar C. Lineberry, ChiefFlight Planning Branch

Approved: i0Ronald L. Berry, Chief T

Mission Planning and Analysis Divi

Mission Planning and Analysis Divisivk

National Aeronautics and Space Administration

Lyndon B. Johnson Space Center

Houston, Texas

July 1979

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ACKNOWLEDGMENT

The following Mission Planning and Analysis Division (MPAD) organizations and in-dividuals have contributed to the development and documentation of this STS -7Flight Feasibility Assessment document either directly or by providing input andconsultation.

Res onsibility Name Mail Phone

MPAD flight Jerome Bell FM2 4346eointegrator 'lax Kilbourne FM2 4401

TDRSS requirements and jzrome Bell FM2 4346interface

Propulsive consumables Don Nelson FM15 4907

Nonpropulsive consumables Harry Kolkhorst FM2 3278Chuck Pace FM2 3278

Deorbit Dallas Ives FM4 6347

Attitude and pointing Dave Scheffman FM2 3278time line Rocky Duncan FM2 3278

Brody McCafferty FM2 3278

Ground navigation W. Wollenhaupt FM8 9751J. C. Currie FM8 3921

Ascent performance John McNeely FM4 5954Gus Babb FM2 4401

Onorbit dispersion Allen Dupont FM2 5276

Entry R. Moore FM4 4401

In addition, MPAD wishes to express gratitude for contributions made by individ-uals in the following areas:

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Gonzalo Montoya (MDTSCO): Guidelines and groundrules, flight design annexTed Guillory (FOD): Crew activitiesDon Lewis (FOD): Crew activitiesSandy Price (MDTSCO): Documentation using FDS/Daconics systemBob Culpepper (MDTSCO): Documentation using FDS/Daconies systemLex Allen (MDTSCO): Document integration, IUS performance, andactivities time lineDick Paulson (MDTSCO): IUS performanceLarry Guderian (MDTSCO): Attitude and pointing time lineArshad Mian (MDT6 A): Attitude and pointing time lineRick Berry (MD7SCO): Propulsive consumablesSam Wilson (TRW): Separation maneuverCourtney Wright (Kentron) Launch window

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Virginia A. Duke (Kentron): Launch windowSherry Fries (Kentron): Launch windowTom Locke (MDTSCO): ReentryGary Zoerner (MDZACO): AscentM. Hazelrigs (MDTSCO): Crew activitiesLinda Langston (MDTSCO): Ascent

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CONTENTS

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Section Page

1 .0 INTRODUCTION . . . . . . . . . . . . . . . . . . . . 1

2 .0 ACRONYMS . . . . . . . . . . . . . . . . . . . . . 2

3.0 GUIDELINES AND GROUNDRULES . . . . . . . . . . . . . . . 5

3.1 GENERAL FLIGHT REQUIREMENTS . . . . . . . . . . . . . . . 5

4.0 FLIGHT DESIGN ASSESSMENT FOR STS-7 . . . . . . . . . 9

4.1 LAUNCH WINDOW j,.NALYSIS . . . . . . . . . . . . . . . . 9

4.1.1 TDRS-A Launch Window Requirements and Constraints . . . . 94.1.2 Orbiter Launch Window Requirements and Constraints . . . 94.1.3 Integrated STS-7 Launch Window . . . . . . . . . . . . 10

4.2 ASCENT PERFORMANCE . . . . . . . . . . . . . . . . . . . 11

4 .3 FLIGHT DURATION . . . . . . . . . . . . . . . . . . . . 12

4.3.1 Ascending Node Deployment Requirement . . . . . . . . . 124.3.2 Contingency Payload Bay Door Opening Requirement 124.3.3 Crew Work/Rest Cycle . . . . . . . . . . . . . . . . . 124.3.4 Onorbit Maneuver Requirements . . . . . . . . . . . . 13

4 .4 ATMOSPHERIC DESCENT . . . . . . . . . . . . . . . . . . . 14

4.5 SEPARATION SEQUENCE ASSESSMENT . . . . . . . . . . . 15

4.5.1 Performance. . . . . . . . 154.5.2 Evaluation of Potential Damage to Orbiter

Windows and Tiles . . . . . . . . . . . . . . . . . . . 15

4.6 ATTITUDE AND POINTING CONSTRAINTS . . . . . . . . 16

4.6.1 Compliance With TDRS Thermal Constraints . . . . . . . 17

4.6.1.1 Cumulative Time in Non-ZLV AttitudeWhile the Payload is Stowed . . . . . . . . . . . . . 17

4.6.1.2 Cumulative Time in Non-ZLV AttitudeWhile the IUS is Elevated . . . . . . . . . . . . . . 17

4.6.1.3 Compliance with Solar Constraints . . . . . . . . . 174.6.1.4 Compliance With Deployment Constraints . . . . . . . . . 184.6.1.5 Assessment of Backup Opportunities . . . . . . . . . . . 18

4.6.2 Compliance With TDRS RF Communication Requirements . . . 184.6.3 Compliance With IUS RF Requirements . . . . . . . . . . 19

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CONTENTS

Section Page

4.6.4 Attitude Compatibility Between TDRSand IUS RF Requirements . . . . , 19

` 4.7 IUS ATTITUDE INITIALIZATION REQUIREMENT . . . . . . . . . 20

4.7.1 Impact on TDRS RF Checkout Operations . . . . . . . . . . 20` 4.7.2 Impact on Backup Deployment 20

4.7.3 Impact on State Vector Initialization . . . . , . . . . . 204.7.4 Ascending Node Injection Requirement . . . . . . . . . . 21

4.8 ASCENDING NODE INJECTION REQUIREMENT . 1 . . . . 1 . . 22

4.8.1 Night Deployml=nt Requirements . . . . . . . v . . . . . . 22i4.8.2 Performance Requirements . . . . . . . . . . . . . . . . 22

4.8.2.1 OHS Requirements . . . . . . . . . . . . . , . . 224.8.2.2 RCS . . . . . . . . . . . . . . . . . . . . . . . . 23

4.8.3 Proximity Operations Requirement3. . . . . . . . . . . . 23j 4.8.4 TDRS/IUS Design . . . . . . . . . . . . . . . 23

4.9 ORBITER COMPATIBILITY . . . . . . . . . . . . . . . . . . 24

4.10 PROPULSIVE CONSUMABLES . . . . . . . . . . . . . . 25

4.11 NONPROPULSIVE CONSUMABLES . . . . . . . . . . . . . . . . 26

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4.12 NAVIGATION . . . . . . . . . . . . . . . . . . . • 27

' 4.12.1 Early Return Contingency . . . . . . . . . . . . : . . . 274 .12.2 IUS/TDRS Deployment . . . . . . . . . . . . . . . . . . . 274.12.3 Deorbit . . . . . . . . . . . . . . . . . . . . . . 27

5.0 REFERENCES . . . . . . . . . . . . . . . . . . . . . . 28

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TABLES

Table Page

I CURRENT SRM PARTICLE DAMAGE LIMITS FORSTANDARD SEPARATION SEQUENCE DESIGN . . . . . . .. . . . . 29

TI MINIMUM RCS PROPELLANT BUDGET . . . . . . . . . . . . . . 30

III OMS PROPELLANT BUDGET . . . . . . . . . . . . . . . . . . 31

IV ORBITER MASS PROPERTIES DURING THE MISSION . . . . . . . 32

V STS -7 TDRS COMPATIBILITY ASSESSMENT . . . . . . . . • . . 33

VI STS-7 COMPATIBILITY ASSESSMENT . . . . . . . . . . . 34

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FIGURES

Figure Page

1 STS-7 launch window

(a) TDRS descending node injection . . . . . . . 35(b) TDRS ascending node injection . . . . . . . . . . . 36(c) Orbiter abort Lighting . . . . . . . . . . . . . . . 37(d) Launch window composite . . . . . . . . . . . . . . 38

2 Deorbit performance requirements for 530west descending node injection . . . . . . . . . . . . . 39

3 Entry interface conditions for STS -7conceptual profile . . . . . . . . . . . . . . . . . . . 40

4 Effect of IUS pitch angle on hi-temp the erosion 41

5 Effect of IUS pitch angle on hi-temp the breakage . . 42

6 Attitude and pointing

(a) Deep space facing of the Orbiter payloadbay for different operations. 43

(b) Total deep space facing with raised tilttable is 44 minutes .

... . . . . . 44

(e) Position of Sun in Orbiter body blockage whilemaneuvering from -ZLV to deployment attitude . . . . 45

(d) Position of AGO in TDRS antenna beam for twodifferent inertial attitudes . . . . . . . . . . . . 46

(e) 450 field of view ...

. . . . . 47(f) Assumed hemispherical omni field of view . . . . . . 48(g) Position of HAW in IUS omni antenna FOV and

relative to the body blockage . . . . . . . . . . . 50(h) Overlap of FOV for TDRS and IUS omni antennas. . . . 51

7 Deorbit AV requirements for ascending nodeinjection opportunity . . . . . . . . . . . . . . . . . 52

8 Nonpropulsive consumables

(a) Flight 7 TDRS - potable H2O profile. . 53(b) Payload bay door opening time line for ascent . . 54(c) Shuttle payload bay closing time line descent . . . 55

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1.0 INTRODUCTION

The Space Transportation System (STS) Flight Assignment Manifest (ref. 1) hasscheduled the first tracking and data relay satellite system (TDRSS) spacecraft(TDRS-A) for a February 27, 1981 launch on STS-7.

This flight design document has been developed by the Mission Planning and Anal-ysis Division (MPAD) in support of the TDRS-A cargo integration review scheduledfor June 13, 1979. It is the companion document of the STS-7 ConceptualFlight Profile (CFP) (ref. 2).

This STS-7 Flight Feasibility Assessment (FFA), along with the STS-7 CFP,is intended to provide a base from which the various design, operation, andintegration elements associated with TDRS-A can perform mission planningand analysis. The STS-7 FFA identifies conflicts, issues and concerns associatedwith the integrated flight design requirements and constraints.

Questions concerning this document should be addressed to Jerome Bell, FlightPlanning Branch (FM2).

For questions relating to specific disciplines, the appropriate personnelidentified in the acknowledgement may be-contacted.

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2.0 ACRONYMS

AFSCF Air Force satellite control facility

AFO abort from orbit

AOA abort once around

AOS acquisition of signal

APU a,sx liary power units

ASE airborne support equipment

ATCS active thermal control system

CFP conceptual flight profile

e.g. center of gravity

CIR cargo integration review

At time increment

AV incremental velocity

DOD Department of Defense

EAFB Edwards Air Force Base

EDT eastern daylight time

EPDC electrical power distribution and control

ET external tank

EVA extravehicular activity

FCP fuel cell power plant

FTR flight test requirements

FWD Toward

GET ground elapsed time

GMT Greenwich mean time

GPC general purpose computer

GSTDN ground spaceflight tracking and data network

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ha

hp

IMU

IUS

JSC

KSC

LH

LOPT

aLOS

LV

LVLII

MECO

MPAD

MPS

NPC

OA

OMS

OMS-1

OMS-2

OP

PET

PI

PIP

PLHD

PROP

psf

apogee altitude

perigee altitude

inertial measurement unit

inertial upper stage

Johnson Space Center

Kennedy Space Center

local horizontal

landing opportunity

loss of signal

local vertical

local verticr^F..,'Iooal horizontal

main-engine cutoff

Mission Planning and Analysis Diviaion

main propulsion subsystem

rionpropulsive consumables

Orbiter after

orbital maneuvering system

first OHS maneuver

second OMS maneuver

Orbiter prior

phase elapsed time

payload integrator

payload integration plan

payload bay doors

propellant

pounds per square foot,

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a a

PTC passive thermal control

gmax maximum dynamic pressure

RCS reaction control system (primary)

RF radio Frequency

RMS remote manipulator system

RTLS return-to-launoh site

RTS remote tracking stations

SPIDPO Shuttle payload integration and development program office

SRB solid rocket booster

SRM-1 ZUS stage-1 solid rocket motor

SAM-2 IUS stage-2 solid rocket motor

SSME .space Shuttle main engine

ST star tracker

STS Space Transportation System

SV Shuttle vehicle

TBD to be determined

TCS thermal control system

TDRS tracking and data relay satellite

TDRS-A first TDRS spacecraft

TDRSS tracking and data relay satellite system

TVCS thrust vector control, system

VRCS vernier reaction control system

WTR Western Test Range

.-ZLV payloads-bay-to-Earth attitude

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3.0 GUIDELINES AND GROUNDRULES

3.1 GENERAL FLIGHT AEQU.IREMENTS

a. The launch date is February 27, 1981..b. Nominal and-of-mission shall be planned for 2 days.

c. The nominal post-Orbiter maneuvering system-2 (OMS-2) parking orbitis a 150 -n. mi. circular one.

d. At the time of deployment, the minimum parking orbit shall be the equiv-alent of a 150-n. mi. circular orbit.

e. The nominal parking orbit inclination is 28.48 degrees.

f. The launch and landing site is Kennedy Space Centex (KSC),

g. The payload complement consists of a tracking and data relay satellite(TDRS-A) spacecraft integrated on a Department of Defense (DOD) two-stage inertial upper stage (IUS), the IUS airborne support equipment(ASE), and the necessary Space Transportation System (STS) cargo-chargeable equipment required to interface the IUS vehicle with theOrbiter.

h. The crew size is four.

i, Orbiter vehicle 102 configuration per reference 2 will be used,.

J. The capability shall be provided to allow a return from orbit withouthaving to deploy the IUSITDRS.

k. Launch window shall be selected to prevent nominal end-of-mission orabort landings from occurring prior to sunrise or later than sunset.

1. Return-to-launch site (RTLS) and abort-once-around (AOA) landings willbe planned to be at KSC,

M. Provide the consumables loading to allow a landing within 7 hours GETfor an abort from orbit WO).

n. A backup landing opportunity will be provided one revolution after nomi-nal landing.

o. The maximum space Shuttle main-engine (SSME) thrust for nominal ascentis 100 percent; for aborts, the maximum thrust is 109 percent.

P. Lift-off, end-of-mission, and abort landing payload weights are per thePayload Data Annex to the TDRS Payload Integration Plan (PIP).

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q. The payload bay doors (PLBD) are to be opened as soon as operationallyconvenient after OMS-2; however, keeping the PLBD closed for up to 3hours postlaunch shall not preclude continuation of the mission.

r. The TDRS command and telemetry links must be checked out onorbit priorto deployment. The nominal path will be y : Ground spaceflight trackingand data network (GSTDN), Orbiter, payload interrogator, and TDRS.

s. One opportunity shall be provided for a direct TDRS to GSTDN radiofrequency (RF) check prior to deployment. This is a contingencyoperation.

t. When the PLBD are open, the Orbiter will fly a payload bay to Earth u(-ZLV) attitude except during the following activities:

(1) All Orbiter inertial measurement unit (IMU) alinements n

(2) TDRS/STDN direct RF check

(3) IUS attitude initialization

(4) IUS/TDRS deployment operation

(5) Preentry thermal conditioning, as required

u. The nominal geosynchronous placement is longitude 53 0 W.s

v. The maximum payload allowance will be based on two-sigma flight perfor-mance reserve loading for AOA. s

d

w. There will be four potable water tanks available for cooling using theflash evaporator. Also, one additional waste water tank can be usedfor additional cooling during aborts and contingencies. The potablewater tanks will be 95-percent full for normal entry.

x. For nonpropulsive consumables budgeting, the following contingencieswill be considered:

z

(1) A 24-hour hold without reservicen

(2) The worst ease of the following:

(a) Cabin puncture

(b) One extravehicular activity (EVA)9

(c) Last deorbit opportunity on mission extension day N

(d) One cabin repressurization

(e) Deorbit one orbit latetF

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Y. Computation and communications required to develop and transmit aground navigation state vector and Orbiter maneuver include the follow-ing:

(1) Tracking passes over at least three stations distributed duringone complete revolution are regaired to acquire enough data forcomputing an accurate ground navigation state vector.

(2) Two additional tracking passes are required to provide backup andmaintain navigation accuracy in the event of tracking station lossduring one of the passes in (1) above. These backup passes may belocated either before, after, or before and after the tracking in-terval in (1) above.

(3) All station passes should be above 3-degree ground station eleva-tion.

(4) Fifteen minutes are required for ground computation of statevector.

(5) Twenty minutes are to be allocated for computation of the Orbitermaneuver and uplink pads given the above state vector as input.

(6) One primary and one backup station pass are required for uplinkingthe state vector and/or maneuver data.

z. When possible, deorbit should be executed on a path that allows tPack-ing by a station between deorbit cutoff and entry interface. This sta-tion pass must be at a minimum of 14-degree elevation.

aa. Propellant loading for attitude control shall be planned on the basisof using primary RCS only. The resulting propellant loading will beneeded in the event of a failure of the vernier RCS.

bb. The IUS flight operations requirements and constraints are as pres-ented to the Shuttle Payload Antegration and Development Program Of-fice (SPIDPO) at the Johnson Space Center (JSC) April 17, 1979 anddocumented in a letter from Col. Shaffer (IUS Program Director) to G.Lunney (Manager) SPIDPO.

cc. The TDRS flight requirements and constraints are as defined in theTDRS PIP, April 19, 1979.

dd. The detailed TDRS/IUS data required for flight design implementationare as defined in the TDRS/IUS PIP annexes.

ee. The Orbiter separation sequence wily be designed in accordance withthe criteria and philosophy contained in formal briefings to STS man-agement (refs. 3 and 4).

ff. Nine hundred-n. mi. erossrange operational capability for landing willbe assumed.

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gg. The solid rocket booster (SRB) configuration is the TC-121-78 and usesthe Western Test Range (WTR) burn rate.

hh. There is no SRB ignition delay.

ii. SSME propellant line screens are assumed to be removed for thisflight.

jj. The abort decision lag time is zero.

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4.0 FLIGHT DESIGN ASSESSMENT FOR STS -7

4.1 LAUNCH WINDOW ANALYSIS

4. 1.1 TDRS-A Launch Window Requirements and Constraints

Figures 1(a) and 1(b) translate TDR4',-A '.:,uneh window requirements and con-straints, as defined in the TDRS PIP, into Orbiter lift-off time requirements.Each individual TDRS requirement and constraint are shown. Figure 1(a) is fordescending ar4 figure 1(b) for ascending node injection opportunities, re-spectively. tThe 25.5 degree constraint between the geosynchronous orbit planeand the ecliptic plane is not violated for a targel, orbit inclination of 2degrees or less. It does not impact the STS -7 launch window). Figures 1(a)and 1(b) show also that the requirement for a right. ascension of the ascendingnode between 270 and 290 degrees defines, for both injection opportunities,a TDRS launch window independent from the other launch window constraints.

The final TDRS-A placement longitude is required to be between longitudes 550Wand 990W. For the descending node case, a launch window that opens at 19:20:28GMT and closes at 20:16:33 GMT will provide three consecutive deployment opportu-nities. For the ascending node case, a launch window that opens at 19:34:35 GMTand closes at 20:32:32 GMT provides two consecutive deployment opportunities. Alaunch window compatible with all deployment opportunities under considerationwill require a lift-off between 19:34:35 and 20:16:33 GMT.

The maximum right ascension of the ascending node compatible with both des-cending and ascending injection opportunities for a February 27th launch isabout 245.5 degrees. This coincides with a lift-off at the closing of thelaunch window associated with the 25-degree nadir constraint for an ascendingnode injection.

4.1.2 Orbiter Launch Window Requirements and Constraints

The only STS-7 Orbiter constraint identified to date is for landing to occurduring daylight hours. A definition of the landing time and landing siteis required to convert this landing constraint into lift-off time require-ments. Figure 1(c) shows the Orbiter lift-off time requirements.

Landings resulting from launch aborts (RTLS and AOA) are assumed to occur atKSC. Launch abort landings are nominally constrained to occur between 30minutes after sunrise and 30 minutes prior to sunset. This limits launch tooccur between 12:19:24 and 21:21:12 GMT for a February 27th launch day. Ifrequired, the launch Window may be expanded by permitting landings as early assunrise or as late as sunset. Launch window open is defined by the RTLS landingtime while the AOA landing time defines the launch window close.

The capability must also be provided for return after OMS-2 in the event the pay-load bay doors cannot be opened. Navigation support requirements based on GSTDN(without an operational TDRS) and'the constraint that the Orbiter APU's cannot

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be restarted for 3 hours precludes landing opportunities on the second orbit;i.e., deorbit one revolution beyond OMS-2 with subsequent landing at about 3hours. For landings from the third orbit (about 4.9 hours GET, which coin-cides with the maximum water loading capability), the crossrange to KSC is about1000 n. mi. This exceeds the 900-n. mi. capability of the Orbiter, necessi-tating a landing at an alternate site. Figure 1(c) shows the daylight landinglaunch window constraints for landing at Northrup and Edwards. It is seen thatan EAFB landing provides for a launch time 46 minutes later in the day than theNorthrup landing site. Planning EAFB as the landing site for the post OMS-2abort maximizes the launch window and also accommodates the desirable landinglighting margin of 30 minutes. Use of Northrup would reduce the launch windowduration by a minimum of 12 minutes and would also entail a reduction in thedaylight remaining at the time of landing.

The nominal landing is not shown on figure 1(c) because, at present, it is notthought to be a factor. Nominal and backup landing opportunities can be selectedprior to 24 and 48 hours GET, which implies landing earlier in the day thanlaunch occurred. In fact, the daylight landing constraint is estimated to excludeonly one landing opportunity to KSC within the 2-day flight. The latest deorbitopportunity for landing during the first day in conjunction with a lift-off atthe close of the TDRS launch window will occur at night.

4.1.3 Integrated STS-7 Launch Window

The composite STS-7 launch window is shown in figure 1(d). Summarized on thisfigure is the acceptable launch window for TDRS-A ascending node injection, TDRSdescending node injection, Orbiter landing lighting, and the integrated STS-7launch window. The launch time (between 19:34:35 and 20:16:33 GMT on February27, 1981) satisfies all the launch window requirements and constraints.

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4.2 ASCENT PERFORMANCE

TBS

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4.3 FLIGHT DURATION

Landing opportunities, satisfying the Orbiter crossrange constraint occur on mul-tiples of every 22 through 24 hours following launch. The mission duration isdependent an how requirements for ascending node deployment, contingency payloadbay door opening, the crew work/rest cycle and onorbit maneuvering can be met.

4.3.1 Ascending Node Deployment Requirement

The flight ,aquirement to provide an ascending node backup deployment opportu-nity requires Orbiter activities associated with IUS/TDRS operations to continuethrough approximately 23 hours 42 minutes GET - IUS first-stage ignition. Thelatest planned first-day deorbit back to KSC, allowing for a backup deorbit op-portunity on the following orbit, occurs at approximately 24 hours 50 minutesGET. Assuming (a) 2 to 4 hours for deorbit preparation; (b) a quiescent coastphase prior to deorbit once the predeorbit tracking and state vector determina-tion operations are initiated; and (c) the deorbit tracking are is initiated aminimum of about 2.5 hours prior to deorbit to satisfy data acquisition, pro-cessing and uplink requirements, landing within approximately 24 hours afterlaunch is incompatible with an ascending node deployment. These constraintsnecessitate landings when the prime landing site is next accessible - during the48-hour GET time frame.

4.3.2 Contingency Payload Bay Door Opening Requirement

The requirement to accommodate a delayed payload bay door opening (contingencyor otherwise) for up to 3 hours past launch will deplete the Orbiter water forcooling to a level that could impact mission duration. Under the guideline thatthe four potable water tanks and the wastewater tank are to be full prior to nom-inal deorbit, fuel cell water will need to be generated onorbit to refill thetanks. The quantity of water produced and the rate at which it is replenishedis a function of when the payload doors are opened, the radiator performance,the power level being utilized, and the orbital attitude time line. It could,therefore, require more than 24 hours GET to produce sufficient water forcooling to support the nominal deorbit and contingency landing operationsrequired for the flight.

4.3.3 Crew Work/Rest Cycle

The TDRS longitude placement requirement provides a maximum, for a 1-day mis-sion, of about 15 hours between nominal (the first) deployment at 10 hours GETand deorbit. Requirements to provide backup deployment opportunities on subse-quent revolutions could reduce the available time interval to 12 hours. Withinthe 12 to 15 hours available, postdeployment proximity operations, crew sleep pe-riod, meals, and predeorbit activities must occur. The feasibility of a 1-dayflight must be assessed relative to the impact on the crew work/rest cycle re-quirements.

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4.3.4 Onorbit Maneuver Requirements

Deorbit opportunities associated with a 1-day flight could potentially requirean additional 30- to 50-fp3 AV. (fig. 2). This assessment is based on the as-sumption that crew work/rest cycle constraints will preclude the deorbit opportu-nity being selected based on minimum AV criteria. Instead, maximizing the timebetween deployment and deorbit for accomplishing all postdeployment/predeorbitactivities will dictate the selection. The deorbit time for a 2-day flight canpresumably be selected based on minimum AV deorbit requirements, with the crew-related activities worked around the deorbit time. The additional DV requirementfor a 1-day flight may be accommodated within required RCS loadings availableprovided the minimum loading (as defined by system constraints) adequately ex-ceed the nominal flight requirements.

Analysis is being done to assess modifications to the separation maneuver.These modifications will not require additional propellant loading nor violatethe criteria for selection of deorbit time.

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4.4 ATMOSPHERIC DESCENT

Based on an assumed nominal entry weight of 188 000 pounds, a center of gravity(e.g.) location (in nominal X, X, Z entry coordinates) of 1102.8, 0.5 1 and372.5, respectively, an entry weight of 230 000 pounds for an intact abort, andthe TPS design limits defined for STS-6, both the heavy and light weight atmos-pheric descents for STS -7 are feasible. A 900-n.mi. orossrange capability isavailable for a e.g. up to approximately 67 percent of the body length in theintact abort entry case. Shifts of the e.g. past that point would have to beaccommodated at the expense of orossrange capability because of TPS designlimits required for this flight.

Entry interface target lines for nominal and intact abort entries are presentedin figure 3. The indicated target points represent a Hohmann transfer from a150-n. mi. circular parking orbit for both the heavy and lightweight vehicle.The difference in the target line for the two entry configurations, represent a10-fps difference in deorbit AV; 280 fps are required for the 230 000-pound vehi-cle, and 270 fps are required for the 18d 000-pound vehicle.

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4.5 SEPARATION SEQUENCE ASSWSMENT

4,5.1 Performance

No performance penalty is incurred with the nominal separation sequence. Theabort from orbit following OMS-2 dictates the OMS propellant required for de-orbit. The quantity of propellant required to return with the TDRS /IUS onboardprovides a AV capability on the order of 50 fps greater than that required whenthe payload is deployed.In addition, abort from orbit requires approximately25 fps greater deorbit AV than the nominal mission. An abort from orbit withthe TDRS/IUS onboard, then, will require an additional OMs QV capability of ap-proximately 75 fps. The separation maneuver requires a AV of about 70 fps.Thus, the separation sequence can use the excess AV required for this abort casewithout impacting the OMS loading requirements,

4.5,2 Evaluation of Potential Damage to Orbiter Windows and Tiles

Based on the window and thermal tile damage budgets currently being used forstandard separation sequence design (table I), the potential of 49 breaks persquare foot for Orbiter windows exposed to the SRM exhaust particle flux wascalculated. This value is almost 700 times greater than the per-firing limitshown in table I, and five times greater, than the lifetime window breakagelimit. Therefore, it is imperative that the windows be shielded from the SRMexhaust plume. This is accomplished by pointing the Orbiter's undersideat the IUS prior to SRM-1 ignition and maintaining that attitude long enoughafter burnout to allow for the finite flight time of the impinging particles,

Other than the windows, only the Orbiter's high-temperature tiles will beexposed to a significant flux of exhaust particles during, and immediatelyafter, the SRM-1 burn.' The high=temperature tile breakage potential is 0.0008breaks per square foot for the nominal case, and the corresponding erosion poten-tial is 0,033 percent. These values are well within the current per-firinglimits. Figures 4 and 5 indicate that a significant deviation of the IUS atti-tude could cause the per- firing trajectory design limits on high-temperaturetile erosion and breakage, respectively, to be e.x3eeded by factors of about1.1 and 2.1. Such a deviation is not likely to occur, and even if it should,it is believed that the resulting damage would not be great enough to representa flight safety hazard or a very serious Orbiter maintenance problem.

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4.6 ATTITUDE AND POINTING CONSTRAINTS

The attitude and pointing profile for this Shuttle flight is unique because itsupports the deployment of the first tracking and data relay satellite (TDRS) tobe placed in geostationary orbit by an inertial upper stage (IUS). There are nu-merous Orbiter and payload/systems operational and hardware-related requirementsand constraints that must be considered in the design of this profile, and eaonrequirement and constraint must be implicitly satisfied by the attitude timeline. For this flight, both the TDRS and the attached IUS impose attitude re-quirements on the Orbiter that serve as drivers to the design of the STS-7 pro-file. There are three basic payload-related attitude profile design drivers.

First, the TDRS is thermally sensitive and must be protected from exposure tothe Sun as well as from long-duration exposure to deep space. Henoe, prior toand at deployment, the Orbiter attitudes must be selected to satisfy theseconstraints. The IUS is also thermally sensitive, but the TDRS thermal sensitiv-ity far exceeds that of the IUS.

Second, both the TDRS and IUS communications requirements must be met. Prior todeployment, the TDRS is required to communicate with a spacecraft tracking anddata network (STDN) ground station (either through the payload interrogator (PI)or directly) to verify the radio frequency (RF) link. Similarly, direct IUS com-muni.cations with an Air Force satellite/spaceoraft control facility (APSCF) re-mote tracking site (RTS) is required to verify the IUS command link. Afterdeployment, the IUS is required to perform autoantenna selection to maintain acommand link with the Orbiter, and the Orbiter must assume an attitude to pointits S-band payload antenna toward the IUS until the IUS is out of communication/tracking range.

Third, the IUS requires a very accurate initial attitude and state vector. Theinitial IUS attitude and state vector are obtained via the Orbiter/payloadinterface. That attitude is refined via star scan operations supported byOrbiter attitude maneuvers, and the state vector is transferred to the IUS fromthe Orbiter as soon as possible after uplink and attitude initialization. Toensure that the state is not degraded by uncompensated translational velocitychanges because of uncoupled control thruster firings, attitude maneuvers mustbe minimized after state vector transfer.

Another significant attitude constr,W.rlt is imposed on the Orbiter by the IUS.That is, the Orbiter must assume a "window protection attitude" during the IUSSRM-1 burn to avoid damage of the Orbiter windows by aluminum oxide and carbonparticles.

After IUS transfer, the Orbiter attitude constraints are not significantly dif-ferent from other STS flights. The attitude profile for the remainder ofonorbit stay time will consist of an Orbiter thermally benign attitude periodi-cally interrupted by routine inertial measurement unit (IMU) alinements.

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4.6.1 Compliance With TDRS Thermal Constraints

The base attitude profile selected for this flight consists of a »Z local verti-cal attitude hold (-ZLV), This attitude satisfies both Orbiter and payloadthermal. constraints. However, deviations from this base profile are requiredto support Orbiter IMU alinements and IUS IMU alinements via star scanningoperations/maneuvers. Other deviations from this base attitude are rsgviredto support TDRS/IUS communications, deployment and tracking.

4.6.1 .1 Cumulative Time in Non-ZLV Attitude Whiles the payload is Stowed

TDRS thermal constraints permit a maximum of 90 minutes of cumulative deep apacefacing, However, both IMU alinements and star scan operations require the pay-load bay (»Z axis) to be pointed towards deep space. Prior to deployment, twoOrbiter IMU alinements and two IUS star scan operations are re quired. The oumu-lative time required by these operations is shown in figure 6(a). As the figureillustrates, the total deep space pointing required by these operations is about87 minutes, assuming that a nominal IMU alinement requires 15 minutes, Thus^the90 minutes deep space pointing constraint is not violated for IUS stowed,

14.6.1.2 Cumulative Time in Non -ZLV Attitude; While the IUS is Elevated

Due to IUS power limitations, the Orbiter is not allowed to point the payloadbay towards deep space for more than 2.5 hours, while the SUS is elevated. Fig-ure 6(b) shows the impact of holding the deployment attitude on the deep spacefacing constraint. In doing so, the Orbiter -Z axis sweeps away from the nadirtowards the horizon in 28 minutes, after which it points towards deep space for14 14 minutes until the IUS is ejected. The total time in non-ZLV attitude whilethe IUS is elevated is 65 minutes. The figure indicates that this pointing con-straint is not violated.

14.6.1.3 Compliance With Solar Constraints

The primary Orbiter attitude requirement is to point the payload bay towardsthe center of the Earth (-ZLV + 2 degrees). This attitude satisfies theTDRS Sun-in-bay constraint when stowed. All attitude maneuvers (IMU alinements,star scan operations) are performed in darkness to ensure solar constraintsare met

When the IUS tilt table is elevated, the Orbiter maintains inertial holdin the deployment attitude, at which Orbiter body blockage shields tho TDRSfrom the Sun. If the TDRS RF check via the PI fails at HAW, then a maneuverto and from a backup RF-oheck attitude at AGO must be performed (in darkness).

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4.6.1.4 Compliance With Deployment Constraints

The deployment attitude is defined by pointing the Orbiter body vector

pitch (P) = 238 degreesyaw (Y) = 0 degrees

towards the Sun, Befure ejection of the IUS, the tilt table is elevated to 58degrees so that the ZUS -X axis points towards the Sun. The deployment attitude:thus provides the maximum solar shadowing that can be achieved from the Orbiterbody. Figure 6(c) shows the position of the Sun relative to the body blockagewhen the Orbiter maneuvers from -ZLV to deployment attitude and holds that atti-tude until ejection of IUS/TDRS. The figure illustrates that in deployment atti-tude the TDRS is always shielded from the Sun by the body blockage. Also, thedeployment attitude is compatible with the OMS prethrust attitude alinement forthe separation maneuver.

4.6.1.5 Assessment of Backup Opportunities

If the TDRS RF check via PI fails over HAW, then the Orbiter must maneuver to an-other attitude for the backup RF check via AGO. Maneuvering to and from this at-titude requires a substantial amount of time because of the large eigenangle be-tween the two attitudes, This maneuver could take much longer if there is a re-quirement to use vernier reaction control. system (VRCS) jets because of theelevated tilt table structural constraints. Also, degradation of state vectoraccuracy may be expeoted because of these maneuvers.

Cumulative deep space viewing of 87 minutes (stowed) and 44 minutes (elevated)is consumed by the nominal attitude profile. Therefore, no more IMU alinementsand star scan operations can be accommodated within the budget. However, theIUS can be kept elevated for another orbit to meet contingencies withoutviolating the 2.5 hours deep space pointing constraint.

4.6.2 Compliance With TDRS RF Communication Requirements

The antenna field of view (FOV) as stated in the PIP is +20 degrees about theTDRS +Z axis. This FOV is not capable of providing STDN tracking for more than3 minutes in a ZLV or inertial attitude under the best conditions. Thus, anySTDN tracking requirement in effect implies the use of the Earth` target-traokingoption of the universal pointing function,

The actual antenna pattern can be modeled as a cone with a 110 • by 79-degree el-liptical cross section. The centerline of the antenna is defined by the bodyvector (P 35 degrees Y = 325 degrees) when the tilt table is elevated 29degrees. This FOV can be utilized to provide approximately 5 minutes of track-ing coverage in an inertial attitude. The coverage time can be increased byusing a biased LVLH hold. Figure 6(d) shows how a STDN station (AGO) sweepsthrough the antenna FOV throughout the station pass, The antenna FOV shown isin the body coordinate system.

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4.6.3 Com fiance With IUS RF Requirements

The stated antenna FOV is + 45 degrees, The centerline of the antenna liesalong the body vector (P =119 deg, X a 33 drag) when the IUS is elevated to 29degrees.

For the IUS IF oheok, the Orbiter must maneuver to keep an RTS within the anten-na FOV throughout the station pass, Figure 6(e) shows the IUS RF communioationopportunities for the above FOV when the Orbiter is in a ZLV attitude. The fig-ure illustrates that the 145 degrees FOV is essentially incompatible with ZLV at-titude operations. Because the 45 degrees FOV IUS omni antenna is associatedwith RTS passes that occur in daylight, IUS RF communication requirements in gen-eral cannot be satisfied with this FOV.

The assumed antenna FOV is +80 degrees and can provide complete coverage in aninertial attitude. Figure 6(f) shows the coverage times available for differentRTS stations when th y+ Orbiter is in a ZLV attitude. The complete coverage pro-vided by the assumed FOV of +80 degrees is obvious from the figure. Figure 6(g)shows the assumed FOV in Orbiter body coordinate system. The position of HAW asit sweeps through the antenna FOV when the Orbiter Is in deployment attitude isalso illustrated.

14.6.4 Attitude Compatibility Between TDRS °nd IUS RF Requirements

The IUS and TARS antennas are incompatible for simultaneous tracking becausethere is a very small overlap of FOV as shown in figure 6(h). Thus, coveragefor one antenna practically implies no coverage for the other.

e

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4.7 IUS ATTITUDE INITIALIZATION REQUIREMENT

4.7.1 Impact on TDRS RF Checkout Operations

For the deployment time selected, constraining the IUS attitude initializationto occur 45 minutes prior to deployment is incompatible with TDRS RF checkoutand transmitter operati ins requirements. The TDRS RF activity requires thespacecraft to be el.evateO, while the star scan activity requires the IUS to bestowed. (",is is indepedent of attitude requirements.) These conflicting re-quirements dictate that the star scan operation be performed before turning onthe TDRS transmitter. Performing TDRS RF checks first would require the TARS tobe restowed after the checks, the transmitter turned off, and then repeatingthis TDRS event after the star scan activity has been completed and prior todeployment. In addition, the maneuver to the star scan attitude could not beginuntil the Orbiter is in darkness, (about 42 minutes prior to deployment at 9hours 23 minutes GET). Allowing 3 to 6 minutes for maneuvering to star scanacquisition attitude and 10 minutes for star scan activities, then 3 to 6minutes for maneuvering to the "correct RF checkout attitude" and 5 minutes forelevation of the tilt table to 29 degrees, 21 to 27 of the available 42 minutesof darkness are used. By this time Santiago may no longer be accessible to sup-port the Contingency RF activities (coverage of Santiago is between 9:41:41 and9:46:23 GET) via the payload interrogator. For direct GSTDN to TDRS RF commu-nications (press uming GSTDN coverage is available), an attitude maneuver will berequired from the contingency RF attitude to deployment attitude. These maneu-vers potentially require 3 to 6 minutes using the PRCS and a significantlylonger time if the vernier system is used when the TDRS/IUS is elevated.

4.7.2 Impact on Backup Deployment

The 45-minute attitude initialization constraint requires that a star scan opera-tion be performed prior to every deployment opportunity. This implies that no-go decisions based upon TDRS RF considerations (or any no-go decision made afterthe TDRS is elevated) will require restowage, turning off the TDRS transmitter,and possibly repeating the TDRS checkout operations after the IUS attitudeinitialization.

4.7.3 Impact on State Vector Initialization

The maneuvers_ associated with star scan operations may have a significant effecton the accuracy of the state vector transferred to the IUS, especially becausethere is an IUS requirement for state vector transfer after the attitudeinitialization operation. This concern originates from: potential uncoupled at-titude control from the Orbiter RCS that could be further aggravated by poten-tial requirements for contamination avoidance thruster inhibits; the potentialstate vector propagation requirements resulting from sparse ground-station track-ing coverage; the frequency with which the attitude initialization is inferredco be required for backup deployment; and potential longer thrust times toachieve attitude rates that minimize time line impacts.

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4.7.4 Ascending Node Infection Requirement

The 45-minute attitude initialization requirement appears incompatible (fromMRS thermal constraints) with an ascending node injection opportunity. For thelaunch window available, the 45 minutes prior to an ascending node deployment isin daylight, and the required star scan maneuvers may severely ,jeopardize theTDRS.

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4.8 ASCENDING NODE INJECTION REQUIREMENT

4.8.1 Night Deployment Requirements

With a nomimal 51-minute coast between deployment and IUS ignition, ascendingnode deployment would occur at approximately 22 hours 49 minutes GET (IUS igni-tion occurs at about 23 hours 40 minutes GET) for a 680 W geosynchronous place-ment, OMS ignition for separation would occur at 23 hours GET. The launch win-dow for the February 27, 1981 launch date results in onorbit sunset occurring be-tween about 22 hours 50 minutes and 22 hours 53 minutes GET (approximately 1 to4 minutes after deployment). Although general requirements have been statedthat the Orbiter shall be capable of deployment in Earth shadow, there is a con-cern whether this is compatible with the present Orbiter and IUS design. The ex-ecution of the various required proximity operations aetiviti.es during nightside passes from deployment until the OMS separation maneuver has not been fullyassessed. For the present, it is assumed desirable to have a minimum of 5minutes of daylight remaining following the OMS separation maneuver for visualsighting of the IUS/TDRS during the initial stage of the separation trajectory.Extending the postdeployment coast to 67 mintites (see. 4.8.2.1) would providefrom 6 to 9 minutes of daylight following the OMS separation burn. This wouldsatisfy the desired visibility requirement.

4.8.2 Performance Requirements

4.8.2.1 OMS Requirements

For the 51-minute coast between deployment and IUS ignition, the ascending nodeinjection opportunity requires deployment to occur approximately 6 minutes priorto arrival at the descending node. The subsequent postejection 11-minute coast,prior to the OMS separation maneuver, will result in apogee of the postsepara-tion orbit (approximately 188 n. mi.) to be in the northern hemisphere. Theresulting geometry between the orbital line of apsides orientation and the land-ing site at the time of deorbit, will result in an increase in the AV requiredfor deorbit (assuming landing occurs on the nominal revolution) of about 44 fpsabove the AV available. The AV requirement can be reduced by delaying deorbit.However, the requirement for a one-revolution backup deorbit opportunity in con-Junction with a daylight landing, limits the nominal deorbit to occur at 47:27GET, which is associated with about a 20-fps AV pe;.zalty. (For the nominaldescending node injection, the geometry is reversed; i.e., apogee is located inthe southern hemisphere, and the orbit is oriented in essentially near-optimum lo-cation for deorbit).

The QV penalty associated with the ascending node injection opportunity can beeliminated by increasing the coast time between deployment and IUS ignition toapproximately 67 minutes (16 minutes longer than nominal). Again, allowing forthe 11-minute coast time prior to OMS ignition for separation, the earlierdeployment time (relative to IUS ignition) permits a more optimum orientation ofthe orbit at the time of deorbit. In addition, because the overall time betweenOMS separation and IUS ignition will also be increased, a 10-fps AV savings is

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realized in the Orbi;:er/IUS separation maneuver sequence. As in the ease of the51-minute coast time between deployment and IUS ignition, the AV requirement forthe 67-minute case can also be reduced by delaying deorbit. However, thecombined AV requirements for the separation and deorbit maneuvers for the 67-minute postdeployment coast will match the available capability for planneddeorbit on the nominal revolution. Figure 7 illustrates the deorbit require-ments for the ascending node deployment opportunity.

4.8.2.2 RCS

Deployment for the ascending node injeetton opportunity will increase RCS propel-lant required for the separation phase by an additional 99 pounds above nominal.Of that quantity, 42 pounds is consumed from the forward RCS tank and 57 poundsfrom the aft tank. This additional usage is based on a 51-minute coast fromdeployment to SRM ignition. If a 67-minute coast is used, an additional 16pounds of RCS propellant (5 forward and 11 aft) is required as a result of thelonger coast time. The majority of the additional propellant required, is dueto the difference in deployment attitude and the maneuvers necessary for the OMSburn (assuming the Orbiter position at SRM-1 ignition remains fixed as behindand above). After deployment, the initial RCS translation maneuver is increasedfrom about 3 fps (nominal) to about 6.5 fps.

4.8.3 Proximity Operations Requirements

Providing for an ascending node injection opportunity will necessitate develop-ment of additional proximity operations trajectories and procedures, and poten-tially require additional training. This is required primarily because atTDRS/IUS deployment, the attitude must be such that the Sun is on the TDRS/IUS-X axis. For the nominal descending node injection opportunity, the Orbiterdeployment attitude requirements relative to local vertical are approximately-101 degrees pitch, -34 degrees yaw, and 200 degrees roll. For the ascendingnode opportunity (assuming deployment occurs 51 minutes prior to IUS ignition),the required deployment attitude is -47 degrees pitch, 3 degrees yaw, and 317 de-grees roll. This difference in deployment attitude requires modified proceduresin order to achieve the required attitude for the OMS maneuver, which is essen-tially fixed (within limits) with respect to the local vertical.

4.8.4 TDRS/IUS Design

Increasing the coast time between deployment and IUS ignition beyond 51 minutesmay impact the TDRS/IUS power requirements, and thus a thermal assessment wouldbe required.

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4.9 ORBITER COMPATIBILITY

There is concern whether the TDRS/IUS flight requirements and constraints willbe compatible with Orbiter capability existing at the time of the plannedflight, especially in light of an accelerated OFT program. In certain areas,such as entry, the flight profile generally assumed the availability of Orbiteroperational capability - the required capability having been demonstrated duringthe six-flight OFT program. A reduced OFT program prior to the TDRS/IUS flightmay result in placards being placed on Orbiter operations. An Orbiter/TDRS/IUScompatibility assessment with respect to the integrated operations needs to beperformed, particularly in the area of thermal requirements and constraints.This assessment needs to be made with respect to the proposed four-flight OFTprogram.

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4.10 PROPULSIVE CONSUMABLES

This document presents the OMS and RCS propellant budgets for the STS-7 mission.In addition, the initial mass properties of the Orbiter component loads atlift-off are given. Finally,'an Orbiter mass properties time history is shownat specific events. The tables presented are reproduced from the STS -7 CFP(ref. 2).

The RCS propellant budget shown in table II is the minimum RCS propellant usagefor the mission. The forward RCS tanks are loaded at the lowest capacityallowed and show a margin of 478.1 pounds. The aft RCS tanks are loaded atgreater than the minimum load allowed. Actually, the minimum total load couldbe reduced by the 15-pound margin shown. This was not done at the time of thestudy because this 15-pound difference was not considered significant in termsof making another study. However, this minimum RCS loading does not allow forany growth in the mission RCS maneuvers. Therefore, it is extremely doubtfulthat this minimum RCS loading philosophy will be used for the mission. It isrecommended that both the forward and aft RCS tanks be full loaded for the mis-sion.

The OMS propellant budget given in table III shows the propellant usages for twocases; Case I is the budget for the mission abort after the OMS-2 burn and CaseII is the budget for the nominal mission. The primary difference between thetwo is that in Case I the payload is not deployed. For Case I, the onorbit OMSusage is entirely for the OMS-2 burn. In Case II, the onorbit usage is the sumof the OMS 2 burn and the payload separation burn. The OMS-1 and OMS-2 burnsfor both cases have the same AVs. The OMS deorbit OVs are listed in the OMS pro-pellant budget for the two cases. Based on the OMS requirements for the mis-sion, the current OMS load is satisfactory.

Table IV shows the initial mass properties of the Orbiter component loads. Allof these values are subject to change. The RCS propellant load will probably beincreased as further FTR's and maneuvers are added to the mission. The nonpro-pulsive consumables load will be changed as mission requirements become moreclearly defined. Finally, the payload weight is subject to change.

Table V shows the Orbiter mass properties at important events during the missionfor the two cases discussed above. The Orbiter mass properties at entryinterface show that the X e.g. for each ease is well within the allowable limitsof 1083.2 and 1109.0 inches.

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4.11 NONNAOPULsin CONSOMAHLES

The nonpropulsive conaumables level A compatibility assessment for STS-7 hasbeen performed. The results of this assessment are presented in table V. Awater management profile is shown in figure 8, and an environmental controlincompatibility/solution table is presented in table VI.

a a

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4.12 NAVIGATION

4.12.1 Early Return Contingency

In the event that either the P LBD or staff s tracker doors do not open, the missionjoust be terminated. A navigation vector cannot be provided to support a 3:25GET landing at KSC. However, an adequate navigation vector is available to sup-port an EAFB landing at 4:40 GET.

4.12.2 IUS/TDRS Deployment

A navigation vector can be provided to support the nominal IUS/TDRS-A deploymenton orbit 8 (10:05 GET) and the backup opportunity on orbit 10. An adequate navi-gation vector cannot be provided to support the one-revolution later deploymenton orbit 9.

For the next day ascending case deployment, a navigation vector can be providedto support IUS/TDPS deployment (22:33 GET) provided DOD C-band support is avail-able.

4.12.3 Deorbit

Should a real-time decision be made to deorbit the first day, orbits 18 and 19can be used for deorbit provided DOD C-band support is available. Ground naviga-tion support is marginal for supporting a KSC landing on orbit 17. Adequateground navigation support cannot be provided to support earlier KSC landing op-portunities (orbits 15 and 16).

The nominal deorbit opportunity (orbit 31) has poor navigation support even withDOD C-band support. Very little tracking data are available above 3 degrees ele-vation and if Orral Valley (ORB) should fail during this critical period, navi-gation would not be able to provide an adequate navigation vector.

The remaining landing opportunities for second-day deorbit appear to be accept-able (from a navigation viewpoint) provided DOD C-band support is available. Ifthis C-band support is not available, landing.during orbits 33 and 34 would bepreferred (from a navigation viewpoint).

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5.0 REFERENCES

1. STS Flight Assignment Manifest. JSC-13000 -0-6P, Apr. 30 o 1979.

2. STS-7 Conceptual Flight Profile. JSC-15045, June 1979.

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I .

TABLE II.- MINIMUM RCS PROPELLANT BUDGET

;Propellant usage, lb Forward After Total

ET separation (4 fps) 57.2 114.5 171.7

Orbit trim maneuvers(15 fps) 4.5 406.3 410,8

PL separation (3.8 fps) 96.9 75.5 172.4

Additional prop forascending node PL sep 47.0 69.0 116.0

Attitude maneuvers 425.3 956.5 1381.8

Deorbit maneuvers .0 1181.2 1181.2

Total usable required 630.9 2803.0 3433.9

Trapped, display andcontrol 492.0 942.0 11134.0

Total required 1122.9 3745.0 4867.9

Total load 1601.0 3760.0 5361.0

Margina 478.1 15.0 493.1

aMbximum RCS load available = 7508 pounds.

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TABLE V. STS-7 TDRS COMPATIBILITY ASSESSMENT

Nonpropulsive consumables

Type ofConsumables requirement Quantity

Electrical power Power level 18.3 kw averageEnergy 1267, kwhVoltage/ourrent AcceptableFuel/cell reactants Two tank sets

Environmental (cabin) Air temperature NominalAir mixture NominalHumidity NominalWater loop temperature NominalWater storage A waste water dump is

required at <4 hr GET;A potable water dump isrequired at <28 hr GET

Active thermal control Freon temperature NominalRadiator Supplemental

flash evaporatorcooling required

Flash evaporator water AcceptableAmmonia cooling Acceptable

33

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TABLE VI.- STS:-7 COMPATIBILITY ASSESSMENT

Environmental control

Incompatibility Solution

Deployment of payload requires Mission planning musta period with no maneuvering; accommodate one waterit is assumed %nat this precludes dump between normal PLBDwater dumps

opening and T + 4-hr GETand two dumps a day duringflight

More detailed temperatureassessment should be per-formed once mission re-quirements have matured

If "no maneuvering periods' alsorequire inhibiting the flashevaporator operation

o Current ATCS freon temperaturewill have a heat sink outlettemperature >400

o Cabin and avionics baytemperature limits may beexueeded

o Payload freon loop temperaturemay exceed limits

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