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Updated October 2, 2000

DISCOVERY:

DELIVERING 

FRAMEWORK

FOR FUTURE

SPACE STATION

GROWTH

STS-92
STS-92

Table of Contents

Mission Overview .........................................................................................................1

Mission Index/Objectives .............................................................................................2

Crew ...............................................................................................................................8

Flight Day Summary Timeline ...................................................................................12

Rendezvous ................................................................................................................13

EVAs

STS-92 EVA ................................................................................................................15

Payloads

Pressurized Mating Adapter 3 .....................................................................................31

Z1 Integrated Truss Segment ......................................................................................33

DTO/DSO/RME

DTO 675 Incapacitated EVA Crewmember Translation ...............................................37

DTO 689 Safer Flight Demonstration ..........................................................................38

DTO 700-14 Single-String Global Positioning System .................................................40

DTO 847 Solid-State Star Tracker (SSST) Size Limitations ........................................41

DTO 257 Structural Dynamics Model Verification ........................................................42
STS-92

Shuttle Reference and Data 

Shuttle Abort Modes ....................................................................................................43

Shuttle External Tank ..................................................................................................48

Space Shuttle Rendezvous Maneuvers .......................................................................51

Space Shuttle Solid Rocket Boosters ..........................................................................52

Space Shuttle Super-Lightweight Tank (SLWT) ..........................................................60

Acronyms and Abbreviations .......................................................................................61

Mission Benefits ........................................................................................................66

Media Assistance ......................................................................................................69

Media Contacts ..........................................................................................................71
STS-92

Mission Overview

DISCOVERY DELIVERS FRAMEWORK FOR FUTURE SPACE 

STATION GROWTH

Space Shuttle Discovery is poised to deliver the next in a series of major hardware 

components to the International Space Station (ISS) during the STS-92 mission as 

the new international facility receives its first framework structure to house 

communications and motion control equipment. 

Launch of Discovery is scheduled to occur no earlier than 9:38 p.m. EDT on Oct. 5 on 

the 28th flight of the orbiter and the 100th mission in Shuttle Program history.

A seven-person crew will be commanded by Brian Duffy, (Col., USAF), who will be 

making his fourth flight into space. Duffy will be joined on the forward flight deck by 

Pilot Pam Melroy (Lt. Col., USAF), who will be will making her first flight into space as 

the third female shuttle pilot in history, following in the footsteps of Eileen Collins and 

Susan Kilrain.

Over the course of four scheduled spacewalks, two teams of space walkers and an 

experienced robot arm operator will collaborate to install the so-called Z1 (Z for zenith 

port) truss structure on top of the U.S. Unity connecting node on the growing station 

and to deliver the third Pressurized Mating Adapter (PMA 3) to the ISS for the future 

berthing of new station components and to accommodate shuttle dockings. 

The Z1 truss will be the first permanent lattice-work structure for the ISS, very much 

like a girder, setting the stage for the future addition of the station's major trusses or 

backbones. The Z1 fixture will also serve as the platform on which the huge U.S. 

solar arrays will be mounted on the next shuttle assembly flight, STS-97.

The Z1 contains four large gyroscopic devices, called Control Moment Gyros 

(CMGs), which will be used to maneuver the ISS into the proper orientation on orbit 

once they are activated following the installation of the U.S. laboratory.

1
STS-92

Discovery OV103

Launch: Thursday, October 05, 2000 

 9:38 PM (eastern time)

Mission Objectives

For Mission 3A, logistics transfer refers to the movement of items between the Space 

Shuttle and the ISS. These tasks are based on experience moving items between the 

Shuttle and the Russian Mir space station. The transfer tasks begin with the unbuckling 

of launch restraints (or opening lockers) and continue through the transfer and stowage. 

If required, transfer time also includes the installation and power-up of items.

The transfer of items between the Space Shuttle and the ISS is orchestrated by the 

loadmaster. The rest of the crew members do the actual moving. The loadmaster is the 

lead for all transfer tasks, organizes the orbiter unloading, supervises ISS stowage, is 

the single point of contact between the crew members and the ground for all transfer 

operations, and is responsible for keeping the transfer lists up to date. Since only a 

small number of items are to be transferred on Mission 3A, there is only one 

loadmaster.

Transfer Overview

STS-92 has two opportunities for transfer. As planned, ISS entry can occur on FD4 and 

FD9. Items can be transferred on both days, but some equipment can be transferred only 

on FD9 because of the mission design. The latter category includes WIS cables and EVA 

equipment that is to be stowed on board the ISS (e.g., EMU parts and EVA tools).

2
Stowage Accommodations

Crew transfer bags (CTBs) come in four sizes: half, single, double, and triple. These 

sizes relate to middeck locker equivalents (MLEs). Mission plans call for the transfer of 

single CTBs only, but the crew may need to maneuver all four sizes, depending on the 

stowage layout of the ISS. 

The carrying capacity and dimensions of the CTBs are as follows:

· Half bag: 30 pounds maximum carrying capacity, 16-3/4 by 9-3/4 by 9-1/4 inches

· Single bag: 60 pounds maximum carrying capacity, 16-3/4 by 19-3/4 by 9-1/4 inches

· Double bag: 120 pounds maximum carrying capacity, 18-3/4 by 19-3/4 by 18-3/4 

inches

· Triple bag: 180 pounds maximum carrying capacity, 18-3/4 by 19-3/4 by 28 inches

3
Content label information, in English, is as follows:

· Location code

· General description of what is in the bag (e.g., clothing, tools)

· Detailed equipment list and quantity

· Toxicity level of the most toxic item in the bag

Each item in the bags is labeled as follows:

· Bar code label

· Operations nomenclature

· Toxicity level

For a quick visual indication of stowage location, labels are color-coded as follows:

· Salmon--stowed in Node 1

· Buff/tan--stowed in Functional Cargo Block

· Green--"go home"

· White--bag label changes (For bags packed on orbit, the white cards are used to detail 

items inside the bag.)

Node 1 contains four bays for stowage. The deck bay (NOD1D4) contains the hard-shell 

stowage rack. The starboard bay (NOD1S4) contains the early communications system 

and can be used for stowing large items. The overhead bay (NOD1O4) contains a softstowage rack insert for bag stowage. The port bay (NOD1P4) also contains a softstowage rack insert.

4
Some items are planned to be stowed in Zarya. Due to thermal, ventilation, and crew 

access concerns, there are areas of Zarya, the ceiling and starboard side, where 

stowage is not permitted. 

The loadmaster uses transfer lists to track transfers. For Mission 3A, the transfer lists 

are located in the Assembly Ops book. Key information includes

· Items to be transferred

· Item location

· How the item is transferred (bag or stand-alone transfer)

· Items transferred and any special notes, handling, or reference to installation 

procedures

The loadmaster keeps the transfer lists current. Before or during a presleep period, the 

loadmaster transmits the number of items transferred during that day to Mission Control.

Resupply transfer lists for STS-92 include items that are transferred from the orbiter 

middeck to the ISS. These items include contingency water containers that are filled from 

the orbiter galley and new Station Operations Data File books for the Increment 1 crew.

Items for return include used harmful contaminant filters from the Zvezda and common 

berthing mechanism (CBM) controller assemblies from the Z1 vestibule.

General Transfer and Stowage Constraints

Transfer and stowage constraints for STS-92 include the following:

· Because Node 1 is not continuously thermally conditioned before STS-97, items stowed 

inside the node must be able to withstand the -35°F temperatures that could occur.

· Zarya's corridor must not be obstructed because it would hinder the ability of the crew to 

leave the ISS quickly.

· Panels that require crew interface (fire holes, caution and warning panels, etc.) must not 

be blocked by transfer items.

· Care must be taken to prevent items stored temporarily from blocking air circulation vents. 

The transfer lists prohibit the permanent stowage of items over these vents.

Items to be transferred are securely restrained with Velcro straps and Russian bungees. 

The flight rules (both generic and flight-specific rules) provide details on ISS transfer and 

stowage constraints.

5
Crew

Commander: Brian Duffy

Pilot: Pamela A. Melroy

Mission Specialist 1: Leroy Chiao

Mission Specialist 2: William "Bill" S. McArthur

Mission Specialist 3: Peter "Jeff" J.K. Wisoff

Mission Specialist 4: Michael E. Lopez-Alegria

Mission Specialist 5: Koichi Wakata

Launch

Orbiter: Discovery OV103

Launch Site: Kennedy Space Center, Launch Pad 39A

Launch Window: Between 2.5 and 5 Minutes

Altitude: 173 Nautical Miles

Inclination: 51.6 Degrees

Duration: 10 Days 19 Hrs. 4 Min.

Vehicle Data

 

Shuttle Liftoff Weight: 4,520,596 lbs.

Orbiter/Payload Liftoff Weight: 253,807 lbs.

Orbiter/Payload Landing Weight:

 

204,455 lbs.

Software Version: OI-27

Space Shuttle Main Engines

SSME 1: 2045 SSME 2: 2053 SSME 3: 2048

External Tank: ET-104A ( Super Light Weight Tank)

SRB Set: BI-104PF

Shuttle Aborts

Abort Landing Sites

 RTLS: Kennedy Space Center Shuttle Landing Facility

 TAL: Zaragoza

 AOA: Edwards Air Force Base, California

 

6
Landing Date: 10/16/00

Landing Time: 4:42 PM (eastern time)

Primary Landing Site: Kennedy Space Center Shuttle Landing Facility

Payloads

Cargo Bay

Z1 Integrated Truss Segment

Pressurized Mating Adapter 3

7
STS-92

Crew Profile Menu

Commander: Brian Duffy

Brian Duffy (Col., USAF), 47, is Discovery’s Commander for the STS-92 mission. This is 

Duffy’s fourth flight into space, having served as the Pilot on the STS-45 and STS-57 

missions and Commander of the STS-72 mission. Duffy is responsible for the overall 

success and safety of the flight and will be in charge of Discovery’s rendezvous and 

docking to the International Space Station. Before being assigned to this flight, Duffy 

served as the Acting Deputy Director of the Johnson Space Center.

STS-45 and STS-57: Pilot STS-72 mission: Commander

Ascent Seating: Flight Deck - Port Forward

Entry Seating: Flight Deck - Port Forward

Pilot: Pamela A. Melroy

Pam Melroy (Lt. Col., USAF), 39, is Discovery’s Pilot for the STS-92 mission. Flying on 

her first mission, Melroy becomes the third female Pilot in Shuttle program history. 

Melroy will assist Commander Brian Duffy in the rendezvous and docking to the 

International Space Station and will undock Discovery from the ISS at the end of docked 

operations before conducting a flyaround of the complex. Melroy will prepare the berthing 

port mechanisms for the mating of the Z1 truss to the U.S. Unity connecting node, then 

will join crewmate Jeff Wisoff in the newly Z1 truss structure to outfit the truss with 

grounding straps and other logistical hardware. 

She will also operate a 3D IMAX movie camera during the mission to collect footage for a movie being 

produced about the assembly of the new Station.

STS-92 is Melroy's first flight.

Ascent Seating: Flight Deck - Starboard Forward

Entry Seating: Flight Deck - Starboard Forward

8
Mission Specialist 1: Leroy Chiao

Dr. Leroy Chiao, 40, serves as Mission Specialist 1 (MS 1) on his third flight into space on 

STS-92. 

Chiao is primarily responsible for Z1 truss systems and will serve as Extravehicular crew 

member 1 (EV 1) during the first and third space walks of the flight. He will wear the 

space suit bering the solid red stripes. Chiao and Bill McArthur will hook up umbilicals 

between the newly installed Z1 truss and other International Space Station components 

and will help to hook up electrical connections between the third Pressurized Mating 

Adapter and the Unity connecting node. 

During the second and fourth space walks by Jeff Wisoff and Mike Lopez-Alegria, Chiao will serve as the 

intravehicular crew member (IV), responsible for the proper choreography of the space walk.

Chiao previously flew on the STS-65 and the STS-72 missions. 

Ascent Seating: Mid Deck - Starboard

Entry Seating: Mid Deck - Starboard

EV1

Mission Specialist 2: William "Bill" S. McArthur

Bill McArthur (Col., USA), 49, is making his third flight into space on STS-92, having 

previously flown on STS-58 and STS-74, which was the second docking mission to the 

Russian Mir Space Station. 

McArthur will serve as flight engineer during Discovery’s launch and landing and is 

designated Mission Specialist 2 (MS 2). He will also be designated as Extravehicular 

crew member 2 (EV 2) during the flight, and will be teamed with Leroy Chiao during the 

first and third space walks of the flight to hook up umbilicals between the newly installed 

Z1 truss and other International Space Station components and electrical connections 

between the Station’s third Pressurized Mating Adapter and Unity. McArthur will wear the white space suit 

bearing no stripes.

McArthur will also serve as the backup robot arm operator during the second and fourth spacewalks, joining 

Koichi Wakata on the flight deck to maneuver other spacewalkers around the International Space Station.

STS-58 and STS-74.

Ascent Seating: Flight Deck - Center Aft

Entry Seating: Flight Deck - Center Aft

EV2

9
Mission Specialist 3: Peter "Jeff" J.K. Wisoff

Dr. Jeff Wisoff, 42, is making his fourth flight into space on STS-92, having flown 

previously on STS-57, STS-68 and STS-81 to the Russian Mir Space Station. 

Wisoff will serve as Mission Specialist 3 (MS 3) during the flight, and will be designated 

as Extravehicular crew member 3 (EV 3), conducting the second and fourth space walks 

of the mission along with Mike Lopez-Alegria. Wisoff will wear the space suit bearing the 

vertical red stripes. Wisoff and Lopez-Alegria will help to install and hookup the third 

Pressurized Mating Adapter (PMA 3) to the Unity connecting node and will conduct a test 

of jetpacks worn by spacewalkers as they assemble International Space Station 

hardware. 

Wisoff will also serve as the intravehicular crew member (IV) during the first and third space walks of the flight, 

responsible for the proper choreography of those excursions by Leroy Chiao and Bill McArthur.

STS-57, STS-68 and STS-81. 

Ascent Seating: Mid Deck - Port

Entry Seating: Mid Deck - Port

EV3

Mission Specialist 4: Michael E. Lopez-Alegria

Mike Lopez-Alegria (Cmdr., USN), 42, is making his second flight into space on STS-92, 

having first flown on STS-73. 

Lopez-Alegria most recently served as the Director of Operations for NASA at the 

Gagarin Cosmonaut Training Center in Star City, Russia, supervising astronaut training 

for flights to the International Space Station. On STS-92, Lopez-Alegria will serve as 

Mission Specialist 4 (MS 4), and Extravehicular crew member 4 (EV 4). He will wear the 

space suit bearing the diagonal red stripes.

Lopez-Alegria will be paired with Jeff Wisoff for the second and fourth space walks of the flight, to install and 

hookup the third Pressurized Mating Adapter (PMA 3) to the Unity connecting node and to conduct a test of 

jetpacks worn by spacewalkers as they assemble International Space Station hardware. 

Lopez-Alegria will also be the backup robot arm operator for the first and third space walks of the flight, joining 

Koichi Wakata on the flight deck to maneuver other spacewalkers around the International Space Station.

He will also serve as the "loadmaster" for this mission, responsible for the transfer of logistical supplies from 

Discovery to the ISS.

STS-73.

Ascent Seating: Mid Deck - Center

Entry Seating: Flight Deck - Starboard Aft

EV4

10
Mission Specialist 5: Koichi Wakata

Koichi Wakata, 37, represents the Japanese Space Agency (NASDA) on STS-92, his 

second flight into space. Wakata previously flew on STS-72, using the Shuttle’s robot 

arm to retrieve a Japanese science satellite which had been launched on a Japanese HII rocket. 

Wakata is designated as Mission Specialist 5 (MS 5) during Discovery’s flight, 

responsible for all robot arm operations. Wakata will grapple the Z1 truss structure and 

mate it to the U.S. Unity connecting node and will use the arm to install the third 

Pressurized Mating Adapter (PMA 3) to Unity as an additional docking port for 

International Space Station hardware. 

Wakata will spend four days using the arm to transport four spacewalkers around the ISS as they connect 

umbilicals and electrical cables between the Z1 truss, the new mating adapter and the ISS modules.

STS-72. 

Ascent Seating: Flight Deck - Starboard Aft

Entry Seating: Mid Deck - Center

RMS

Updated: 09/25/2000

11
STS-92

Flight Day Summary 

DATE TIME (EST) DAY MET EVENT

10/05/00 9:38:25 PM 1 000/00:00 Launch

10/06/00 1:22:25 AM 1 000/03:44 NC1 Burn

10/06/00 2:56:25 PM 2 000/17:18 NC2 Burn

10/06/00 6:26:25 PM 2 000/20:48 NPC Burn

10/06/00 11:37:25 PM 2 001/01:59 NC3 Burn

10/07/00 11:19:25 AM 3 001/13:41 NC4 Burn

10/07/00 12:45:25 PM 3 001/15:07 Ti Burn

10/07/00 4:31:25 PM 3 001/18:53 Docking

10/07/00 6:57:25 PM 4 001/21:19 ISS Ingress

10/08/00 11:55:25 AM 4 002/14:17 Z1 Grapple

10/09/00 12:43:25 PM 5 003/15:05 EVA 1 Start

10/10/00 12:54:25 PM 6 004/15:16 EVA 2 Start

10/11/00 12:52:25 PM 7 005/15:14 EVA 3 Start

10/12/00 1:13:25 PM 8 006/15:35 EVA 4 Start

10/14/00 12:33:25 PM 10 008/14:55 Undock

10/16/00 3:35:25 PM 12 010/17:57 Deorbit Burn

10/16/00 4:42:25 PM 12 010/19:04 KSC Landing

Updated: 09/30/2000

12
STS-92

Rendezvous 

Orbiter Operations

Overview

Rendezvous and Docking

The orbiter will rendezvous with the International Space Station (ISS) on Flight Day (FD) 

3 at a time based on ISS-orbiter phasing. 

The orbiter will perform a minus R-bar approach to the ISS to protect communications 

with Russian ground sites and will dock with Pressurized Mating Adapter (PMA) 2 on 

the forward common berthing mechanism (CBM) of Unity. If beta angles exceed 45 

degrees, the ISS may need to perform a 90-degree roll to ensure adequate power 

production. If this occurs, the orbiter will have to perform a 90-degree yaw during the 

approach. 

The primary prerendezvous activities include checking out the orbiter's robotic arm, the 

extravehicular maneuvering units (EMUs), the Ku-band antenna, the orbiter docking 

system (ODS), and the ground command system. It may be necessary to power down 

the station's systems to minimize battery discharge during solar array feathering. The 

power-down will be performed at approximately 2 hours before the terminal phase 

initiation, or Ti, burn.

Next the orbiter rendezvous radar will be activated, but it will not be used until after the 

normal course correction burn. The crew will then power the ODS and activate the 

docking lights at approximately 1 hour before the Ti burn. 

The Ti burn will be performed when the ISS is ready for docking. If ISS power or thermal 

concerns arise, a Ti delay burn will be performed instead of the Ti burn. 

Zvezda and Zarya solar arrays will be feathered (or positioned) and locked when the 

orbiter is 150 feet from the station. Feathering is initiated by a command to reposition 

and hold each beta (and alpha) joint at a predetermined angle, which will limit the 

induced loads. 

The ISS will automatically switch to free drift at capture, and the orbiter will go to free 

drift to avoid imposing excessive loads on the orbiter docking system (ODS).

Light-emitting diodes (LEDs) on PMA-2 will blink when the ISS is in free drift. The crew 

will be able to see the red indicators through the overhead window on the orbiter's aft 

flight deck and verify that the ISS is in the free-drift mode. 

The ODS will begin the automatic rigidization and retraction process, and its capture 

hooks will be closed. Then the ODS will be deactivated, Zvezda and Zarya solar arrays 

will resume sun tracking, and the orbiter-ISS stack will maneuver to the mated attitude.

13
Docked Operations

After the orbiter docks with the ISS, the crew will enter PMA-2, take an air sample from 

Unity, and activate the Unity cabin fan to scrub the atmosphere. After a 2-hour scrub, 

the crew will take another air sample. On FD4, the crew will unberth the Z1 segment of 

the station's integrated truss structure from the orbiter payload bay and mate it with 

Unity's zenith CBM port. After the CBM operations have been completed, the crew will 

enter Unity and Zarya. Over the next four days, crew members will conduct four 

spacewalks to connect umbilicals and configure hardware for future operations. The 

crew will re-enter the ISS on FD9 to transfer EMU hardware and complete any 

remaining entry or transfer tasks. The orbiter will undock from the ISS on FD10.

During mated operations, the orbiter will provide attitude control. 

Departure

Before the orbiter departs on FD10, ground personnel will update the ISS vector and 

mass data. This data includes attitude departure maneuver data, attitude hold data, 

postdeparture mass properties, and postdeparture attitude maneuver data. A navigation 

platform alignment will be performed. Zvezda and Zarya solar arrays will be 

commanded to stop sun tracking and will be feathered to the edge-on position.

The orbiter crew will configure orbiter hardware for departure. This involves enabling 

rendezvous navigation and performing an orbiter-to-target state vector transfer. Next, 

the crew will configure the ODS. The orbiter docking lights will be activated, and orbiter 

power will be applied to the ODS. The orbiter digital autopilot will be configured for 

proximity operations and commanded to free drift. Orbiter interface unit (OIU) 

operations will be terminated, and OIU interfaces will be powered off. 

The crew will undock the orbiter from PMA-2 and perform a series of separation burns 

with the primary RCS jets to move the orbiter away from the ISS.

Zvezda's flight control system will be activated by a separation signal from the 

androgynous peripheral attachment system (APAS). This signal initiates a 250-second 

delay before Zvezda's attitude control system is activated. Separation must occur within 

Russian ground coverage because the APAS indicator is mechanically zero-fault 

tolerant and electrically single-fault tolerant. Russian ground personnel will activate the 

attitude control system if the APAS indicator fails. Zvezda will maneuver to its 

postdeparture attitude and return to the XPOP flight mode. Zvezda and Zarya solar 

arrays will resume sun tracking. Finally, Russian segment equipment needed only 

during mated and departure operations (docking system, lights, etc.) will be deactivated.

Updated: 09/30/2000

14
STS-92

EVAs 

Z1 Assembly, Activation, and Checkout

Overview

Mission Operations Summary

STS-92 will carry a crew of seven on an 11-day mission that includes four spacewalks or 

EVAs on four consecutive days.

On Flight Day 1 (FD1), the DC-to-DC converter unit (DDCU) heaters and the wireless 

instrumentation system (WIS) radio frequency (RF) kit will be activated as initial spacewalk 

preparations begin.

On FD2, the crew will activate and check out orbiter and orbiter-based International Space 

Station (ISS) hardware, including the extravehicular mobility unit (EMU), the shuttle's 

robotic arm, or remote manipulator system (SRMS) and the orbiter interface unit (OIU), and 

assembly power conversion unit (APCU) hardware. After the SRMS checkout, the operator 

will perform a payload bay survey to inspect the integrity of the cargo.

The orbiter will rendezvous and dock with the ISS on FD3. After the docking, the crew will 

enter Pressurized Mating Adaptor (PMA) 2 to collect Unity air samples and initiate Unity air 

scrubbing before entering the ISS the next day.

On FD4, the Z1 element will be unberthed from the orbiter and mated with the zenith port of 

Unity. The first opportunity to enter the ISS will occur after Z1 has been installed. If the 

temperature in Unity is acceptable, the crew will enter Zarya,transfer any necessary 

hardware, install the Z1-to-Unity grounding straps, and remove the CPAs and latches in the 

Z1 vestibule.

The four spacewalks begin on FD5 and end on FD8 to link the Z1 Truss and PMA-3 to the 

ISS. The crew will return to the ISS on FD9 to perform tasks that were not completed 

during the first opportunity.

The orbiter is scheduled to undock from the ISS on FD10 and perform a flyaround, but that 

day has also been reserved as an unscheduled EVA day. 

Two EVA teams, each composed of two crew members, ensure backup support on each 

EVA day. One team will perform spacewalks on FD5 and FD7, and the other team on FD6 

and FD8. FD9 and FD10 are reserved in case an unscheduled EVA is required to properly 

configure any ISS elements or to perform a contingency undocking. The EVA teams are 

cross-trained for all assembly activities to accommodate unexpected situations that require 

shuffling of the EVA teams and crew members.

15
Z1 Installation

Before Z1 is mated with Unity, the zenith active common berthing mechanism (ACBM) is 

activated and checked out. This procedure is initiated by applying power to the controllers 

that control the ACBM. Once power has been applied, the 20 individual controller channels 

begin a power-on self-test (POST). The POST results are relayed after a master motor 

controller has been designated during the initialization.

The master controller controls all ACBM functions and reports statuses that it gathers from 

the slave controllers. The early portable computer system (EPCS) reports the status of the 

power-up and remote power controller module (RPCM) switches, and the crew reviews and 

compares the reported status to expected values.

The crew receives the current overall CBM command status, subsystem identification 

(latch/bolt controller), time, status of the motor current, motor speed, bolt load, and shaft 

position. This operation also sets all powered bolt and latch controller positions. Because 

the system has been initialized, these positions should have a reading of zero. If controller 

channel faults are present, fault detection, isolation, and recovery (FDIR) activities are 

initiated. If no faults are detected, the crew continues with the ACBM checkout.

The crew members test the bolts by turning them two turns out and three turns in before 

berthing to ensure that they are operable. The capture latch actuators fully deploy the four 

capture latches to ensure that they are operable before element berthing begins.

The capture latches are then moved from the deploy position to the capture position as part 

of the preberthing latch operation testing. The capture latches must then be redeployed 

before Z1 is positioned in Unity's zenith ACBM capture envelope. The crew uses the same 

procedure to perform a capture latch deploy test.

After the ACBM has been checked out successfully and the capture latches are extended, 

the robotic arm is maneuvered to the grapple position. Z1 is grappled, and the crew 

unlatches the payload retention latch assemblies (PRLAs) and the active keel assembly 

(AKA) holding Z1. The orbiter is taken to free drift, and the arm unberths and maneuvers Z1 

to a low hover position. After some orbiter attitude cleanup, the orbiter is taken to free drift 

again, and Z1 is maneuvered to the premate position. After more attitude cleanup, the 

orbiter is taken back to free drift, and Z1 is maneuvered to the capture position. The orbiter 

must remain in free drift until initial bolt loading (ABOLT) occurs.

When the bolts are ready, they are taken to the final preload bolt (FBOLT) values by 

tightening the bolts, four at a time. An adequate pressure seal can be achieved if 15 of the 

16 bolts or 14 of 16 (nonadjacent) bolts are at the FBOLT load. After FBOLT has been 

achieved, the capture latches are commanded from the capture position to the close 

position. Power is then removed from Unity's zenith ACBM. 

16
PMA-3 Installation

Before PMA-3 is mated with Unity, the nadir ACBM is activated and checked out. This is 

initiated by applying power to the controllers that control the Node 1 nadir ACBM. After 

power has been applied, the 20 individual controller channels begin a POST. The test 

results are relayed after a master motor controller has been designated during the 

initialization.

The master controller controls all ACBM functions and reports statuses that it gathers from 

the slave controllers. The EPCS reports the status of the power-up and RPCM switches, 

and the crew reviews and compares the status to expected values. The crew receives the 

current overall CBM command status, subsystem identification (latch/bolt controller), time, 

status of the motor current, motor speed, bolt load, and shaft position. This operation also 

sets all powered bolt and latch controller positions.

Before berthing begins, the crew members test the bolts to ensure that they are operable, 

and the capture latch actuators fully deploy the four capture latches to ensure that they are 

operable. The capture latches are moved from the deploy position to the capture position 

as part of the preberthing latch operation testing. The capture latches must then be 

redeployed before Z1 is positioned in Unity's zenith ACBM capture envelope. The crew 

uses the same procedure to perform a capture latch deploy test. After the activation and 

17
checkout of the ACBM, the crew must visually verify that there is no debris or 

contamination that would prevent mating with the PCBM or cause degradation of the CBM 

seals.

18
After the ACBM has been checked out and the capture latches are extended, the EVA crew 

manually releases PMA-3 from the Spacelab pallet (SLP). The arm grapples PMA-3, and 

PMA-3 is released from the SLP. Due to the extremely tight clearances (6 inches) between 

the PMA-3 and the SLP, an EVA crew member monitors the unberthing.

When PMA-3 is clear of structure, it is maneuvered to the capture position. The orbiter 

must remain in free drift until ABOLT occurs.

At this point, the EVA crew connects two sets of umbilicals to PMA-3 and Node 1. An 

operational hold is required to prevent CBM seal scrubbing because of an excessive 

temperature delta between the ACBM and PCBM.

At the beginning of the next flight day, the powered bolts are taken to IBOLT values by 

tightening the bolts, four at a time. In order to achieve an acceptable seal between the 

PCBM and the ACBM, 15 of the 16 powered berthing bolts or 14 of 16 (nonadjacent) bolts 

must be engaged. Once the bolts have reached IBOLT status, they are taken to the FBOLT 

values by furhter tightening the bolts. After the final preload status has been achieved, the 

capture latches are commanded from the CAPTURE position to the CLOSE position.

19
DDCU Heat Pipe (HP) Aliveness Test

During EVA3, Mission Control Center, Houston (MCC-H) will perform a DDCU-HP 

aliveness test to verify that the DDCU-HP blind-mate connections mated. There is no 

visible indication of proper connector mating, so this test mitigates the effect of a DDCU 

failure on follow-on flights. The orbiter crew can perform the procedure, but MCC-H will 

have the primary responsibility.

EVA Operations

Relocation of S-Band Antenna Subassembly (SASA)

Leroy Chiao (EV1) and Bill McArthur (EV2) remove eight bolts on various parts of the 

SASA using the pistol grip tool (PGT). Using handrails on the SASA, McArthur pulls the unit 

away from Z1, ensuring it does not contact any structure. (The SASA can survive for only 1 

hour if it is not attached to a structure.) The shuttle’s robot arm guides McArthur through the 

unstow path as Chiao monitors progress and provides other assistance to McArthur. Chiao 

rotates SASA 180 degrees so that the high-gain antenna (HGA) is pointing in the nadir 

direction and then aligns the mast fasteners with a stowage bracket. Chiao then secures 

the mast to the stowage bracket. McArthur releases a bracket from the SASA and stows it 

in the Z1 stowage bin. Chiao connects cables W34-P4 and W07S-P3 to the SASA. The 

receptacle covers for J3 and J4 are then stowed on Z1 dummy receptacles. The ground 

applies heater power soon after the connectors have been mated.

This task must occur before the DDCU is installed because the SASA is launched in the 

DDCU installation locations. 

20
Connection of Z1-to-Unity Umbilicals

The immediate purpose of this task is to connect the umbilicals that provide keep-alive 

power to critical on-orbit replaceable units (ORUs) on Z1. Several Unity umbilicals provide 

APCU power to the CBMs, and these are not disconnected until CBM operations are 

completed later in the flight. With the exception of Russian-American conversion unit 

(RACU) commanding, all power-down and activation commands are sent via the portable 

computer system (PCS) on board the orbiter.

Installation of Space-to-Ground Antenna (SGANT)

Before the SGANT dish can be removed from its launch location, remote power controller 

mode Z14B-B must be verified to be operational. Once the antenna is removed from the 

structure, the dish cools quickly, and the antenna cannot be reinstalled in the launch 

location. After the orbiter docks with the ISS, the SGANT will violate lower thermal limits 

within 20 hours, according to preflight thermal analysis.

21
To install the SGANT, Chiao positions himself in an APFR on the Z1 zenith bulkhead, and 

McArthur moves on the robot arm near the SGANT. McArthur unfastens two launch 

restraints on the antenna while Chiao holds it. Chiao slides the SGANT to the hard stop on 

Z1 and raises it using the upper deployment handles. The SRMS positions McArthur near 

the boom attachment interface while Chiao removes the APFR and installs and enters the 

APFR near the boom attachment interface.

Using the upper deployment handles and handles on the antenna group interface tube 

(AGIT), McArthur positions the AGIT near the boom. Chiao grasps the AGIT handles and 

aligns the AGIT coarse guide pins with the boom interface holes until the AGIT fine-guide 

pins are aligned and the SGANT dish is secured to the boom with four bolts. McArthur then 

connects the J1 and J2 connectors. The intravehicular (IV) crew member verifies that the 

RPCs feeding connector J1 are open and that closure is inhibited before mating.

Deployment of SGANT Boom

Before deploying the boom, the crew verifies that the SGANT and transmit/receive control 

(TRC) heaters are active. Preflight thermal analysis shows that the SGANT will violate 

lower thermal limits within 30 minutes after the boom is deployed.

The spacewalker on the arm removes a pip pin from the launch restraint bolt and then 

unfastens the bolt. The SGANT with boom is unstowed by both spacewalkers. McArthur 

attaches the antenna to Z1 by fastening two captive EVA bolts on the boom to the Z1 

zenith bulkhead using the PGT.

22
Installation of DDCU-HP

Two DDCU-HPs are being flown on STS-92. DDCU Z1-3B is on the starboard sidewall of 

the payload bay, and DDCU Z1-4B is on the port sidewall. The installation procedures for 

the two units are similar.

After the DDCU is released, the heat pipe radiator causes heat to leak from the DDCU. The 

DDCU will reach its survival thermal limit in one hour.

McArthur aligns the DDCU-HP on Z1, using the alignment marks on the unit with marks on 

the truss, and guides the box onto the aft tie-down nuts until the box is seated. When the 

DDCU-HP front screw is fastened to a hard stop, the front status indicator changes from 

UNLOCK to LOCK and the rear status indicator goes from LOCK to UNLOCK. After the 

rear screw is fastened to a hard stop, the rear status indicator goes from UNLOCK to 

LOCK. The ORU is installed.

The DDCU will continue to cool down until heater power is applied, which should occur 

within 1 hour of the installation on the Z1 structure.

23
Release of PMA-3 From SLP

Once the Unity nadir ACBM has been activated and checked out and the ACBM capture 

latches are all extended, EVA crew members Jeff Wisoff and Michael Lopez-Alegria can 

remove PMA-3 from the Spacelab pallet (SLP).

With the robotic arm attached to PMA-3, the two spacewalkers loosen the berthing bolts 

while deploying the capture latches. This series of maneuvers prevents asymmetrical 

loading on the PMA-3 from the MBM.

PMA-3 must be unberthed from the SLP in daylight because cameras must be used during 

the procedure. Because of the tight clearances, camera views are necessary to assist the 

SRMS crew member.

One spacewalker must open and close the capture latches to ensure that they are 

operational.

With the latches closed, the spacewalkers first loosen all bolts one-half turn to relieve 

pressure on the PCBM seal. This is done so that unloading occurs symmetrically; however, 

it does not have to be performed by the two crew members simultaneously. Because of the 

loading conditions, the crew members fully loosen and release all but the four bolts next to 

the MBM capture latches. The EVA astronauts fully unbolt the remaining berthing bolts 

simultaneously while working opposite from each other. PMA-3 is now free of all berthing 

bolts on the MBM ring.

The remaining restraints are the four capture latches. An EVA crew member takes up a 

position at the capture latch drive assembly and drives the latches at a low torque setting 

until they are one turn from being fully deployed. Then the latches are fully opened at a 

higher torque setting. After PMA-3 has been moved from the envelope of the MBM, an EVA 

astronaut drives the capture latches back to their closed position for entry, first within one 

turn of being fully deployed at a higher torque and then to the final position at a lower 

torque.

Monitoring PMA-3 Unberthing

Because of the 6-inch clearance between the PMA-3 docking light assembly and the SLP, 

an EVA crew member must monitor the unberthing. The orbiter-ISS stack must be in free 

drift; once the PMA-3 is clear of the SLP, the orbiter's vernier reaction control system 

(VRCS) is used for attitude control.

Connection of PMA-3 Umbilicals

Once PMA-3 has been successfully berthed and latched to Node 1, an EVA crew member 

can support the primary umbilical release. Lopez-Alegria releases launch restraints and 

clamps so that he can release the zero-gravity connectors and slide the thermal covers off 

the zero-g connectors. Lopez-Alegria then releases one PMA-3 primary umbilical launch 

24
fitting and disconnects the primary umbilicals from the dummy panel. The PMA-3 primary 

umbilical bunch is free of the PMA when the stanchion fitting is released. (A minimum of 20 

pounds is required to release the stanchion after the EVA bolt is free.)

Lopez-Alegria moves to the node after releasing EVA clamp C4 and is ready to help guide 

the umbilical around the grapple fixture. The remaining clamps are released, and the 

primary umbilical is controlled to ensure that the lower part of the umbilical does not float 

into the path of the Wisoff or the SRMS. Wisoff carries the PMA-3 primary umbilical on the 

SRMS around the starboard side (orbiter port) of PMA-3 to the Node 1 end cone handrail, 

where it is temporarily secured. Wisoff holds the PMA-3 primary umbilical stanchion and 

controls the umbilical while moving to the Node 1 end cone on the SRMS. Lopez-Alegria 

must help guide the umbilical around the grapple fixture and ensure that it does not 

become entangled. The RPCs for the PMA shell heaters are opened to provide adequate 

safety inhibits before connectors are mated.

The primary PMA-3 umbilicals will be connected to Unity's forward nadir end cone 

connector panel. After a heater is used overnight to equalize the temperature of the two 

elements, the mating of PMA-3 with Node 1 will be completed at the start of FD7. After the 

two elements have been mated, the redundant cable bundle is connected at the beginning 

of EVA3 on FD7.

25
History/Background

Installation of EVA Tool Stowage Devices (ETSDs)

The robot arm transports the EVA astronaut to the SLP, and he removes the launch 

restraints for ETSD 1. The arm then transports the astronaut and ETSD to the ISS port side 

(orbiter starboard side) of the Z1 aft face, where the ETSD corner pads are aligned with the 

Z1 brackets. Engagement is indicated when the Z1 corner guide soft-dock tabs latch over 

the ETSD corner brackets. Once the ETSD is soft docked, eight fasteners (four at each 

bracket) are engaged with a hard stop. The process is then repeated for the ETSD 2.

Figure 11 shows the ETSD installation.

Final Z1 and PMA-3 Umbilicals

Temporary EPS (APCU) jumpers A (W1002) and B (W1001), which were installed on Flight 

2A, are removed from locations on the Unity forward zenith connector panel so the four Z1 

umbilicals can be installed. Once APCU power is verified off, an EVA astronaut removes 

two APCU jumpers from the zenith side of Unity’s forward end cone connector panel. 

Figure 12 shows this task, and Table 8 lists the function of each umbilical that is to be 

connected.

26
After the four Z1 umbilicals are mated, the DDCU aliveness test can be performed on the 

ground. The PMA-3 and PMA-2 APCU node connectors can be swapped after the 

aliveness test is performed.

At this point, APCU power is no longer available to the ISS. This puts the ISS in a 

configuration to accept 140-volt APCU power on Mission 4A in support of P6 activation.

RTAS Launch Locks Removal

Lopez-Alegria uses the PGT to drive the primary bolt counterclockwise approximately 24 

turns until it reaches a hard stop. Lopez-Alegria then pulls up on the pip pin to disengage 

the launch restraint from the bulkhead and disposes of it in a trash bag. This process is 

repeated for the three remaining bolts. This task must be completed after the Ku-band 

boom is deployed because of clearance problems.

Z1 Flight-Releasable Grapple Fixture (FRGF) Relocation

An EVA crew member disengages eight fasteners on the FRGF using the PGT and 

removes the FRGF from Z1. The SRMS transports the astronaut to the FRGF stowage 

location on Z1. The crew member inserts the FRGF through a hole in the zenith bulkhead 

and draws it up into the stowage bracket until it reaches a hard stop. Using the PGT, the 

astronaut engages two fasteners with a hard stop. Due to clearance problems, this task 

also must be completed after the Ku-band boom is deployed.

Deployment of Z1 Tray

This task supports the power-up of the U.S. Lab and PMA-2 relocation on Mission 5A. It 

must be performed after all Z1-to-Node 1 connectors are mated because access to the 

Node 1 zenith connector panel will be severely restricted once the tray is deployed. Wisoff, 

on the SRMS, holds the tray and applies 40 lbf against it in the stowed position while 

Lopez-Alegria, free floating, removes two pip pins and two restraint pins on the left and 

right of the tray. Lopez-Alegria then lifts the tray launch restraint bracket 3 inches. 

27
Grasping the upper portion (zenith end) of the tray, Wisoff is transported through the 

unstow path by the robot arm. Wisoff adjusts the grasp position, as necessary, during the 

operation. Wisoff must apply up to 40 lbf from the initial unstow until the tray reaches the 

neutral position halfway through the deployment.

Final unstow also requires Wisoff to apply a force of up to 40 lbf until the tray reaches a 

hard stop. While Wisoff holds the tray in the deployed position, Lopez-Alegria engages pip 

pins at the starboard and port hinges to lock the tray. Wisoff folds the clevis down onto Z1 

and stows the pip pins and restraint pins in the clevis. Lopez-Alegria free floats to the ISS 

starboard side of the clevis and attaches an adjustable tether to a loop on Z1. LopezAlegria feeds the free end of the tether through a loop on the clevis, hooks it on Z1, and 

tightens the tether. The tether stabilizes the position of the two space vision system (SVS) 

targets on the clevis, making the Z1 SVS target array ready for the P6-to-Z1 mating on 

Flight 4A.

Z1 Keel Pin Relocation

An EVA crew member unfastens two bolts using the Pistol grip tool with a 7/16-inch 

extension until the fastener plate tick marks are aligned with the end of the guide pins. A 

simultaneous push force of 5 pounds must be applied during torquing. The fasteners are 

stowed on the keel by attaching an adjustable tether to each fastener's lanyard and 

diagonally opposite handholds, snugging the tethers. The keel pin is detached, and the 

SRMS relocates the EVA astronaut and keel to the storage location on the starboard side 

of the zenith bulkhead. The crew member aligns the keel with the stowage lugs until a hard 

stop is reached. Using the Pistol grip tool, the astronaut inserts two bolts to attach the keel 

pin.

Cycle Z1 MBM Latch

This task must be performed after the Z1 tray is deployed because of structural clearance 

problems. The MBM capture latches are cycled, by fully deploying and then fully closing 

the latches, to verify their operability for future operations on Mission 5A/STS-98.

The crew retrieves the power tool that is used to actuate the capture latches on the Z1 

MBM, removes an APFR from stowage, and moves to the worksite on the side of the Z1 

MBM. The crew installs and enters the APFR and places the power tool over the 

microconical interface on the capture latch drive mechanism. The power tool drives the 

capture latches open to the deployed position. This position is the widest point at which the 

latches rotate away from the interface plane of the MBM.

While the power tool is turning the latch drive assembly, the EVA crew monitors the 

position indicator on the drive shaft. Two parallel columns of two holes each are 60 

degrees apart on the shaft. These holes show the capture latches' position as the crew 

drives the mechanism. Markings fill the holes as the drive shaft moves the capture latches 

into position. When the markings fill the holes closest to the MBM, the latches are correctly 

deployed for capture, and the crew can stop the power tool. The crew removes the tool 

from the microconical interface and leaves the APFR. The operation is complete, and all 

supporting tools and equipment are returned to stowage.

28
Rocketdyne Truss Attachment System (RTAS) Capture Latch Cycle

The RTAS capture latch (the "claw") will be used to capture the P6 truss on Mission 4A. 

During Mission 3A, an EVA crew member cycles the latch to verify its functionality and 

leave it in a configuration to support P6 capture. This task is done in parallel with the keel 

pin relocation. The free-floating EVA astronaut moves to the worksite and opens the latch 

using the PGT. Once this is complete, the crew member cycles the latch closed and then 

open.

EVA Timeline for Z1 Assembly, Activation, and Checkout

Time Event

3/15:20 EVA 1 Egress

3/15:00 EVA 1 Setup

3/16:06 EVA 1 Z1 String 1 Umb

3/17:40 EVA 1 SASA Relocate

3/19:11 EVA 1 Z1 String 2 Umb

3/18:10 EVA 1 SGANT Dish Install

3/19:09 EVA 1 SGANT Boom Deploy

3/19:30 EVA 1 Port ETSD Install

3/20:30 EVA 1 Cleanup

3/21:30 EVA 1 Ingress

4/15:10 EVA 2 Egress

4/15:15 EVA 2 Setup

4/15:55 EVA 2 PMA 3 Release from SLP

4/17:10 EVA 2 RTAS LL Release

4/17:10 EVA 2 CIDS Relocate

4/18:05 EVA 2 PMA 3 Install

4/18:40 EVA 2 RTAS LL Release

4/18:40 EVA 2 IAPFR Relocate

4/19:35 EVA 2 Connect PMA 3 Umb

4/20:25 EVA 2 Cleanup

4/21:40 EVA 2 Ingress

5/15:10 EVA 3 Egress

5/15:15 EVA 3 Setup

5/16:15 EVA 3 DDCU-HP 4B Install

5/17:05 EVA 3 DDCU-HP 3B Install

29
Time Event

5/17:51 EVA 3 Final Z1 Umbs

5/18:20 EVA 3 Starboard ETSD Install

5/19:05 EVA 3 Final PMA 3 UMB

5/19:35 EVA 3 Z1 Keel Relocate

5/19:35 EVA 3 OHTS Bag Relocate

5/20:20 EVA 3 Cleanup

5/21:40 EVA 3 Ingress

6/15:30 EVA 4 Egress

6/15:35 EVA 4 Setup

6/16:35 EVA 4 RTAS Latch Cycle

6/17:05 EVA 4 Z1 Tray Deploy

6/17:35 EVA 4 MBM Cycle

6/18:07 EVA 4 4A APFR Relocated

6/18:35 EVA 4 Cleanup

6/19:35 EVA 4 SAFER Demo

6/21:05 EVA 4 Incap Crew Demo

6/21:45 EVA 4 Cleanup

6/22:00 EVA 4 Ingress

Updated: 09/30/2000

30
STS-92

Payloads 

Pressurized Mating Adapter 3

Payload Bay

Prime:

Backup:

Overview

Pressurized Mating Adaptor (PMA) 3 will provide a place for an orbiter to dock with the 

U.S. segment of the International Space Station (ISS). PMA-3 includes mechanical 

interfaces, spacewalk hardware and thermal control equipment, electrical power 

subsystem (EPS) and command and data handling (C&DH) passthroughs. PMA-3 will not 

be used for docking until Mission 4A/STS-97.

Command and Data Handling

PMA-3 provides a hard-line 1553 data bus connection between the orbiter and Unity via 

X-connectors on the PMA-3 androgynous peripheral attachment system (APAS), which 

interfaces with the orbiter docking system (ODS). Umbilical connections between Unity 

and PMA-3 complete this hard-line path. This path allows the orbiter interface units 

(OIUs) and orbiter portable computer system machines to talk on the ISS orbiter buses. 

Extravehicular Activity (EVA) Support

The PMA-3 segment is equipped with the following spacewalk aids: A portable foot 

restraint (PFR) top-mounted worksite interface (WIF) fixture, two flight-releasable grapple 

fixtures (FRGFs), camera and laser targets, a number of Space Vision System (SVS) 

targets, handholds, and handrails.

The PFR WIF fixtures are used to attach the PFR workstation stanchions.

The FRGF provides the standard mechanical interface between the shuttle's robotic arm 

or remote manipulator system (SRMS) and payloads. It is compatible with all large ISS 

manipulator systems. The FRGF can be released during an EVA by rotating two release 

rods that allow the fixture's grapple shaft to be removed. A spare shaft can be installed on 

orbit, enabling the interface to be restored to a capture configuration for retrieving 

payloads.

Camera and laser targets consist of a camera target on the APAS hatch, a hemispherical 

laser target on PMA-3, and planar laser reflectors on the side of the APAS.

Handholds and handrails help EVA crew members move about. They have been placed 

in preplanned paths in and around worksites.

31
Motion Control System

PMA-3 has two sets of four red light-emitting diodes (LEDs) that tell the orbiter crew with 

the status of the ISS attitude control system. The crew can see the LEDs through the 

overhead window on the aft flight deck. Each set of four LEDs is controlled by a separate 

Node 1 MDM for one-fault tolerance during arrival or departure. Free drift is indicated 

when the two sets of LEDs alternately flash on and off at a 5-hertz rate. Every other state 

of the MCS is indicated by a steady on.

Structures and Mechanisms

The PMA-3 is a truncated conical shell with a 24-inch axial offset in the diameters 

between the end rings. It is a ring-stiffened shell structure machined from 2219 aluminum 

alloy roll ring forgings welded together. PMA-3 mechanical interfaces include a passive 

common berthing mechanism and a Russian APAS.

Thermal Control System

PMA-3 has ten 60-watt passive thermal control system heaters, temperature 

instrumentation, and multilayer insulation (MLI).

The PMA-3 shell's temperature is maintained above the minimum level by electrical 

resistance heater circuits, which are controlled by RPCM N1-RS2-B. Each heater has a 

resistive thermal device, which provides temperature data to the Node 1 MDMs. The 

ground or crew can control the shell heater states by altering the setpoints of the heaters.

Radiative heat loss and excessive radiative heating from the space environment are 

minimized by MLI blankets between a micrometeorite/orbital debris shield and the PMA's 

primary structure.

Updated: 09/30/2000

32
STS-92

Payloads 

Z1 Integrated Truss Segment

Payload Bay

Prime:

Backup:

Overview

The Z1 is the base structure for the U.S. solar array. It includes two plasma contactors, two 

DC-to-DC converter units (DDCUs), four control moment gyro (CMG) assemblies, part of one 

string of the S-band communications system, one Ku-band communications system, primary 

and secondary power distribution, thermal control system hardware, mechanical interfaces, 

and EVA/extravehicular robotics (EVR) hardware.

Command and Tracking Subsystem

S-Band Communications System. The S-band communications system consists of two 

redundant strings, each of which comprises three on-orbit replaceable units (ORUs) and two 

antennas, the baseband signal processor (BSP), the Tracking and Data Relay Satellite 

System (TDRSS) transponder, and the antenna RF group, which includes a low-gain 

antenna and a high-gain directional antenna.

The radio frequency group (RFG) has two main functions. It amplifies and filters the radio 

signal and receives and radiates the signal, providing the interface with free space. It also 

controls antenna switching and pointing.

The RFG unit has two antennas. The high-gain antenna (HGA) can support the high-rate 

data link, but it requires TDRSS pointing updates that are generated by the guidance, 

navigation, and control (GN&C) pointing function (which will not be available until the GN&C 

multiplexer/demultiplexer [MDM] is activated on Mission 5A). The low-gain antenna (LGA) is 

fixed in position and needs no pointing data. 

The RFG is launched on the Z1 starboard bulkhead. Since this is the DDCU heat pipe (HP) 

installation location, the RFG must be relocated temporarily during EVA1 to the stowage 

bracket. 

The BSP is the heart of the S-band system. It provides data and voice processing for 

downlink and uplink. A voice line connects directly to the BSP to provide voice 

communications. 

The TDRSS transponder, which will arrive on Mission 4A, receives downlink information from 

the BSP and modulates the RF carrier for transmission to the ground. The transponder 

receives the uplink information from the RF group, demodulates the signal, and routes it to 

the BSP for information separation and distribution.

The downlink output of the BSP is a constant-rate data stream of either 192 kbps or 12 kbps, 

as specified by a configuration command.

33
To support the uplink, the BSP inputs a digital bit stream from the standard TDRSS 

transponder at either 72 or 6 kbps depending on whether the BSP is configured for the high 

or low data rate. The data is processed by decryption, CCSDS header validation, and 

demultiplexing functions. If uplink audio channel data is present, it is expanded to 192 kbps 

per channel for interface to the AUAIU and is routed to the appropriate output port. The 

uplink core data channel packets are passed to the 1553 interface for transfer to the 

command and control processor.

Ku-Band Communications System. The Ku-band system is the primary return link for 

International Space Station (ISS) video and payload data transmitted in digital format to the 

ground. The space-to-ground antenna (SGANT) will be relocated and attached to the singlebeam boom of the current Ku-band antenna on EVA1. 

The Ku-band provides a 50-Mbps fixed-rate downlink with up to four video signals or up to 

43-Mbps high-rate data with 7 Mbps overhead. A communication outage recorder records 

payload data in the zone of exclusion and during structural blockage outages. There are 12 

logical channels (4 video and 8 payload) in the Ku-band downlink. 

Like the S-band system, the Ku-band system does not inspect the data passing through it. 

The ISS transmits at a constant rate of 50 Mbps. Of this, 43 Mbps is available to the user. 

The video channels can be configured to downlink full-motion video or stop-action 

video, which consists of skipping video frames. Normally, the service is configured on 

the ground and commanded through the S-band uplink to the on-board C&DH, which 

routes commands via the 1553 local bus to the Ku-band.

The on-board Ku-band equipment includes a single string of four avionics units-the video 

baseband signal processor (VBSP), high-rate frame multiplexer (HRFM), high-rate modem 

(HRM), and the transmit/receive controller (TRC)-plus an erectable, steerable antenna. Video 

signals must be processed by the VBSP before they are routed to the HRFM for interleaving 

with high-rate data (HRD). HRD comes from the payload patch directly to the HRFM for 

processing and interleaving with the video signal. The HRFM builds the downlink transfer 

frames and routes the data stream to the HRM for modulation on an S-band interface (I/F) 

frequency for routing to the TRC for translation to the Ku-band frequency. The data stream is 

power amplified and routed to the steerable antenna for transmission through the TDRSS to 

the ground. An uplink carrier is received and processed by the TRC to maintain antenna 

autotracking only.

Video channels are directed to the VBSP for processing. The VBSP has four audio and four 

video input channels and four video plus audio output channels. Connectors are also 

provided for power and 1553 bus and I/F data out. The four video plus audio output channels 

are directed to the HRFM. The VBSP performs audio and video analog-to-digital (A/D) format 

conversions, packet formatting, built-in test (BIT), and ORU control functions. The VBSP 

converts incoming analog signals to digital, formats them for the selected mode of operation, 

and forwards them to the HRFM for interleaving with high-rate payload data.

34
Electrical Power Subsystem

The Z1 truss assembly has several power distribution components. These include two 

initialization diode assemblies (IDAs), two secondary power distribution assemblies (SPDAs), 

two plasma contactor units (PCUs), two DDCUs, and two patch panels.

The IDAs provide diode-protected power from the shuttle assembly power conversion unit 

(APCU) to P6 for module initialization on Mission 4A. They also provide a connection path 

from P6 to the lab DDCUs from Mission 5A until Mission 12A, when the P4 module will be 

delivered.

The SPDAs consist of remote power controller modules (RPCMs), power and data 

connections (central utility rail), and a cold plate. They control, protect, and isolate secondary 

distribution lines. The SPDAs accept secondary power from the DDCUs and distribute this 

power to RPDAs or downstream loads. The SPDA cold plates transfer heat from the RPCMs 

via radiant fins. 

The central utility rail provides power connections from the DDCU power feed to the RPCMs 

and data connections via redundant 1553B data buses. Because the DDCUs will not be 

active during Mission 3A, SPDAs Z1-3B and Z1-4B will receive power directly from RussianAmerican conversion units (RACUs) 5 and 6, respectively.

RPCMs are electronic switches that control, protect, and isolate secondary distribution lines. 

There are six types of RPCMs, which differ in their rated output current, number of switches 

per module, and trip functions. 

The PCU emits electrons through a self-generated plasma and is self-regulating. The PCUs 

control the voltage between the space plasma and the ISS structure. PCUs mounted on the 

Z1 truss maintain the structure potential of the ISS within 40 volts of the plasma potential. A 

central element of each PCU is the hollow cathode assembly (HCA), which emits up to 10 

amps of electron current to the ambient plasma. The PCU actively emits when the ISS is in 

sunlight. Without the PCUs, the ISS could reach structure potentials of approximately -150 

volts. 

Two DDCUs, which are launched attached to cargo handling interface adapters (CHIAs), will 

be delivered during STS-92. The DDCUs will be removed from the CHIAs and attached to 

the starboard side of Z1 during EVA1. The units convert power from primary (115 to 173 volts 

DC) to secondary (123 to 126 volts DC). They will not be activated until Mission 4A, when P6 

power comes on line.

Two patch panels on the port side of Z1 allow the Z1 input power source to be changed. On 

the outside of the panels are a fixed output and three interchangeable input connectors. To 

change the panel configuration, an umbilical is simply demated and mated to a different 

connector on the patch panel. The launch configuration of the patch panels provides Russian 

segment (RS) power to Z1 components for keep-alive purposes. The panels will be 

reconfigured during Mission 4A to prepare for the transition to U.S. power via P6.

35
Extravehicular Activity Subsystems

The Z1 truss segment is equipped with several spacewalk aids: Two EVA tool stowage 

devices (ETSDs), 22 worksite interface (WIF) sockets, 1 flight-releasable grapple fixture 

(FRGF), 11 trusses, 2 tray launch restraints, numerous standard handholds and handrails, 

and several custom handles.

Motion Control Subsystem

The motion control subsystem (MCS) hardware launched as part of the Z1 element includes 

the CMGs and the CMG assemblies. This hardware will not be activated until Mission 5A, 

when the GN&C MDM will be activated with the U.S. Lab.

The CMG assembly consists of four CMGs and a micrometeorite/orbital debris shield. The 

four CMGs, which will control the attitude of the ISS, have a spherical momentum storage 

capability of 14,000 ft-lb/sec, the scalar sum of the individual CMG wheel moments. The 

momentum stored in the CMG system at any given time equals the vector sum of the 

individual CMG momentum vectors. 

To maintain the ISS in the desired attitude, the CMG system must cancel, or absorb, the 

momentum generated by the disturbance torques acting on the station. If the average 

disturbance torque is nonzero, the resulting CMG output torque is also nonzero, and 

momentum builds up in the CMG system. When the CMG system saturates, it is unable to 

generate the torque required to cancel the disturbance torque, which results in the loss of 

attitude control. 

The CMG system saturates when momentum vectors have become parallel and only 

momentum vectors change. When this happens, control torques perpendicular to this parallel 

line are possible, and controllability about the parallel line is lost. 

Russian segment thrusters are used to desaturate the CMGs.

An ISS CMG consists of a large flat wheel that rotates at a constant speed (6,600 rpm) and 

develops an angular momentum of 3,500 ft-lb/sec about its spin axis. This rotating wheel is 

mounted in a two-degree-of-freedom gimbal system that can point the spin axis (momentum 

vector) of the wheel in any direction.

At least two CMGs are needed to provide attitude control. The CMG generates an output 

reaction torque that is applied to the ISS by inertially changing the direction of its wheel 

momentum. The CMG's output torque has two components, one proportional to the rate of 

change of the CMG gimbals and a second proportional to the inertial body rate of the ISS as 

sensed at the CMG base. Because the momentum along the direction of the spin axis is 

fixed, the output torque is constrained to lie in the plane of the wheel. That is why one CMG 

cannot provide the three-axis torque needed to control the attitude of the ISS.

Each CMG has a thermostatically controlled survival heater to keep it within thermal limits 

before the CMGs are activated on Mission 5A. The heaters are rated at 120 watts and have 

an operating temperature range of -42 to -35°F.

36
Structures and Mechanisms

Mechanical interfaces are provided between Z1 and the P6 long spacer on the Z1 zenith 

face, between Z1 and Node 1 on the Z1 nadir face, between Z1 and PMA-2 on the Z1 

forward face, and across a cable tray from the Z1 forward face to the U.S. Lab.

A manual berthing mechanism (MBM) on the forward face of Z1 allows temporary stowage of 

PMA-2. The device is similar to the one used on the Spacelab pallet to stow PMA-3 on this 

flight (but without the perimeter bolts). 

A hinged cable tray on the forward face of Z1 provides a direct interface for power, data, and 

coolant lines to the U.S. Lab. Mission 3A EVA crew members will deploy the tray to expose 

the MBM and will release 32 bolts on the 4 coolant lines. 

The Z1 truss structure is designed to maximize component packaging and support load paths 

during launch and on orbit. A solid plate beneath the CMG face provides additional protection 

from micrometeorites and orbital debris for the CMGs and other Z1 components. The 

structural framework beam and shell elements are aluminum 2219-T851. The trunnion and 

keel pin beam elements are INCO 718, and the Ku-band antenna boom beam elements are 

steel.

Thermal Control Subsystem

The Z1 truss assembly includes the following elements of the early external active thermal 

control system (EEATCS): 4 accumulators, ammonia, and 12 quick disconnects and 

associated plumbing. The accumulators charge the EEATCS with ammonia on orbit, 

accommodate thermal expansion in the fluid, and maintain the system's operating pressure. 

The quick disconnects facilitate the connection of ammonia transfer lines between P6 and Z1 

and between Z1 and the U.S. Lab.

Updated: 09/30/2000

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Incapacitated EVA Crewmember Translation

DTO 675

Overview

The object of this DTO is to verify techniques to allow a spacewalking crewmember to 

return a fellow EVA crewmember who is incapacitated to the shuttle airlock. The DTO 

will evaluate several techniques to determine how efficient each is.

This DTO will only be performed if all station assembly tasks are completed and 

consumables are available.

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Safer Flight Demonstration

DTO 689

Overview

The purpose of the DTO is to demonstrate that the Simplified Aid for EVA Rescue 

(SAFER) performs as expected, and to build confidence in the design through an endto-end, on-orbit functional checkout.

There are two test objectives that check out rotation/translation controllers and perform 

a point-to-point translation.

This DTO will be performed only if all station assembly tasks are complete and 

consumables are available.

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Single-String Global Positioning System

DTO 700-14

Overview

DTO 700-14 will demonstrate the performance and operation of the Global Positioning 

System (GPS) during ascent, orbital and entry phases of the flight. It uses a modified 

military GPS receiver/processor and the existing orbiter GPS antennas and records 

data on a payload and general support computer (PGSC) hard drive.

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Solid-State Star Tracker (SSST) Size Limitations

DTO 847

Overview

The objective of this DTO is to characterize the performance of the Solid State Star 

Tracker (SSST) with a large and bright target. The assumption is that the SSST (-Y star 

tracker) will be able to track an object large enough and bright enough to eliminate the 

potential problems with the Image Dissecting Tracker (-Z star tracker) during day tracker 

passes for station flights.

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Structural Dynamics Model Verification

DTO 257

Overview

This DTO will fire a sequence of primary reaction control system (RCS) jets and 

observe the effects of the firings on the structural dynamics of the mated shuttle and 

International Space Station complex. The test is designed to confirm that primary 

reaction control system jet firings are acceptable for use if vernier reaction control 

system jets are not available.

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Shuttle Reference and Data

Shuttle Abort Modes

Selection of an ascent abort mode may become necessary if there is a failure that 

affects vehicle performance, such as the failure of a space shuttle main engine or an 

orbital maneuvering system. Other failures requiring early termination of a flight, such as 

a cabin leak, might also require the selection of an abort mode.

There are two basic types of ascent abort modes for space shuttle missions: intact 

aborts and contingency aborts. Intact aborts are designed to provide a safe return of the 

orbiter to a planned landing site. Contingency aborts are designed to permit flight crew 

survival following more severe failures when an intact abort is not possible. A 

contingency abort would generally result in a ditch operation.

INTACT ABORTS

There are four types of intact aborts: abort to orbit (ATO), abort once around (AOA), 

transoceanic abort landing (TAL) and return to launch site (RTLS).

ABORT TO ORBIT (ATO)

The ATO mode is designed to allow the vehicle to achieve a temporary orbit that is 

lower than the nominal orbit. This mode requires less performance and allows time to 

evaluate problems and then choose either an early deorbit maneuver or an orbital 

maneuvering system thrusting maneuver to raise the orbit and continue the mission.

ABORT ONCE AROUND (AOA)

The AOA is designed to allow the vehicle to fly once around the Earth and make a 

normal entry and landing. This mode generally involves two orbital maneuvering system 

thrusting sequences, with the second sequence being a deorbit maneuver. The entry 

sequence would be similar to a normal entry.

TRANSOCEANIC ABORT LANDING (TAL)

The TAL mode is designed to permit an intact landing on the other side of the Atlantic 

Ocean. This mode results in a ballistic trajectory, which does not require an orbital 

maneuvering system maneuver.

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RETURN TO LAUNCH SITE (RTLS)

The RTLS mode involves flying downrange to dissipate propellant and then turning 

around under power to return directly to a landing at or near the launch site.

ABORT DECISIONS

There is a definite order of preference for the various abort modes. The type of failure 

and the time of the failure determine which type of abort is selected. In cases where 

performance loss is the only factor, the preferred modes would be ATO, AOA, TAL and 

RTLS, in that order. The mode chosen is the highest one that can be completed with the 

remaining vehicle performance. 

In the case of some support system failures, such as cabin leaks or vehicle cooling 

problems, the preferred mode might be the one that will end the mission most quickly. In 

these cases, TAL or RTLS might be preferable to AOA or ATO. A contingency abort is 

never chosen if another abort option exists.

The Mission Control Center-Houston is prime for calling these aborts because it has a 

more precise knowledge of the orbiter's position than the crew can obtain from onboard 

systems. Before main engine cutoff, Mission Control makes periodic calls to the crew to 

tell them which abort mode is (or is not) available. If ground communications are lost, 

the flight crew has onboard methods, such as cue cards, dedicated displays and display 

information, to determine the current abort region.

Which abort mode is selected depends on the cause and timing of the failure causing 

the abort and which mode is safest or improves mission success. If the problem is a 

space shuttle main engine failure, the flight crew and Mission Control Center select the 

best option available at the time a space shuttle main engine fails.

If the problem is a system failure that jeopardizes the vehicle, the fastest abort mode 

that results in the earliest vehicle landing is chosen. RTLS and TAL are the quickest 

options (35 minutes), whereas an AOA requires approximately 90 minutes. Which of 

these is selected depends on the time of the failure with three good space shuttle main 

engines.

The flight crew selects the abort mode by positioning an abort mode switch and 

depressing an abort push button.

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RETURN TO LAUNCH SITE OVERVIEW

The RTLS abort mode is designed to allow the return of the orbiter, crew, and payload 

to the launch site, Kennedy Space Center, approximately 25 minutes after lift-off. 

The RTLS profile is designed to accommodate the loss of thrust from one space shuttle 

main engine between lift-off and approximately four minutes 20 seconds, at which time 

not enough main propulsion system propellant remains to return to the launch site.

An RTLS can be considered to consist of three stages-a powered stage, during which 

the space shuttle main engines are still thrusting; an ET separation phase; and the glide 

phase, during which the orbiter glides to a landing at the Kennedy Space Center. The 

powered RTLS phase begins with the crew selection of the RTLS abort, which is done 

after solid rocket booster separation. The crew selects the abort mode by positioning 

the abort rotary switch to RTLS and depressing the abort push button. The time at 

which the RTLS is selected depends on the reason for the abort. For example, a threeengine RTLS is selected at the last moment, approximately three minutes 34 seconds 

into the mission; whereas an RTLS chosen due to an engine out at lift-off is selected at 

the earliest time, approximately two minutes 20 seconds into the mission (after solid 

rocket booster separation).

After RTLS is selected, the vehicle continues downrange to dissipate excess main 

propulsion system propellant. The goal is to leave only enough main propulsion system 

propellant to be able to turn the vehicle around, fly back towards the Kennedy Space 

Center and achieve the proper main engine cutoff conditions so the vehicle can glide to 

the Kennedy Space Center after external tank separation. During the downrange phase, 

a pitch-around maneuver is initiated (the time depends in part on the time of a space

shuttle main engine failure) to orient the orbiter/external tank configuration to a heads 

up attitude, pointing toward the launch site. At this time, the vehicle is still moving away 

from the launch site, but the space shuttle main engines are now thrusting to null the 

downrange velocity. In addition, excess orbital maneuvering system and reaction control 

system propellants are dumped by continuous orbital maneuvering system and reaction 

control system engine thrustings to improve the orbiter weight and center of gravity for 

the glide phase and landing.

The vehicle will reach the desired main engine cutoff point with less than 2 percent 

excess propellant remaining in the external tank. At main engine cutoff minus 20 

seconds, a pitch-down maneuver (called powered pitch-down) takes the mated vehicle 

to the required external tank separation attitude and pitch rate. After main engine cutoff 

has been commanded, the external tank separation sequence begins, including a 

reaction control system translation that ensures that the orbiter does not recontact the 

external tank and that the orbiter has achieved the necessary pitch attitude to begin the 

glide phase of the RTLS.

After the reaction control system translation maneuver has been completed, the glide 

phase of the RTLS begins. From then on, the RTLS is handled similarly to a normal entry.

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TRANSATLANTIC LANDING ABORT OVERVIEW

The TAL abort mode was developed to improve the options available when a space 

shuttle main engine fails after the last RTLS opportunity but before the first time that an 

AOA can be accomplished with only two space shuttle main engines or when a major 

orbiter system failure, for example, a large cabin pressure leak or cooling system failure, 

occurs after the last RTLS opportunity, making it imperative to land as quickly as 

possible.

In a TAL abort, the vehicle continues on a ballistic trajectory across the Atlantic Ocean 

to land at a predetermined runway. Landing occurs approximately 45 minutes after 

launch. The landing site is selected near the nominal ascent ground track of the orbiter 

in order to make the most efficient use of space shuttle main engine propellant. The 

landing site also must have the necessary runway length, weather conditions and U.S. 

State Department approval. Currently, the three landing sites that have been identified 

for a due east launch are Moron,, Spain; Dakar, Senegal; and Ben Guerur, Morocco (on 

the west coast of Africa).

To select the TAL abort mode, the crew must place the abort rotary switch in the 

TAL/AOA position and depress the abort push button before main engine cutoff. 

(Depressing it after main engine cutoff selects the AOA abort mode.) The TAL abort 

mode begins sending commands to steer the vehicle toward the plane of the landing 

site. It also rolls the vehicle heads up before main engine cutoff and sends commands 

to begin an orbital maneuvering system propellant dump (by burning the propellants 

through the orbital maneuvering system engines and the reaction control system 

engines). This dump is necessary to increase vehicle performance (by decreasing 

weight), to place the center of gravity in the proper place for vehicle control, and to 

decrease the vehicle's landing weight.

TAL is handled like a nominal entry.

ABORT TO ORBIT OVERVIEW

An ATO is an abort mode used to boost the orbiter to a safe orbital altitude when 

performance has been lost and it is impossible to reach the planned orbital altitude. If a 

space shuttle main engine fails in a region that results in a main engine cutoff under 

speed, the Mission Control Center will determine that an abort mode is necessary and 

will inform the crew. The orbital maneuvering system engines would be used to place 

the orbiter in a circular orbit.

ABORT ONCE AROUND OVERVIEW

The AOA abort mode is used in cases in which vehicle performance has been lost to 

such an extent that either it is impossible to achieve a viable orbit or not enough orbital 

maneuvering system propellant is available to accomplish the orbital maneuvering 

system thrusting maneuver to place the orbiter on orbit and the deorbit thrusting 

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STS-92

maneuver. In addition, an AOA is used in cases in which a major systems problem 

(cabin leak, loss of cooling) makes it necessary to land quickly. In the AOA abort mode, 

one orbital maneuvering system thrusting sequence is made to adjust the post-main 

engine cutoff orbit so a second orbital maneuvering system thrusting sequence will 

result in the vehicle deorbiting and landing at the AOA landing site (White Sands, N.M.; 

Edwards Air Force Base; or the Kennedy Space Center.. Thus, an AOA results in the 

orbiter circling the Earth once and landing approximately 90 minutes after lift-off.

After the deorbit thrusting sequence has been executed, the flight crew flies to a landing 

at the planned site much as it would for a nominal entry.

CONTINGENCY ABORT OVERVIEW

Contingency aborts are caused by loss of more than one main engine or failures in 

other systems. Loss of one main engine while another is stuck at a low thrust setting 

may also necessitate a contingency abort. Such an abort would maintain orbiter integrity 

for in-flight crew escape if a landing cannot be achieved at a suitable landing field.

Contingency aborts due to system failures other than those involving the main engines 

would normally result in an intact recovery of vehicle and crew. Loss of more than one 

main engine may, depending on engine failure times, result in a safe runway landing. 

However, in most three-engine-out cases during ascent, the orbiter would have to be 

ditched. The in-flight crew escape system would be used before ditching the orbiter.

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Space Shuttle External Tank

The external tank contains the liquid hydrogen fuel and liquid oxygen oxidizer and 

supplies them under pressure to the three space shuttle main engines in the orbiter 

during lift-off and ascent. When the SSMEs are shut down, the ET is jettisoned, enters 

the Earth's atmosphere, breaks up, and impacts in a remote ocean area. It is not 

recovered. 

The largest and heaviest (when loaded) element of the space shuttle, the ET has three

major components: the forward liquid oxygen tank, an unpressurized intertank that 

contains most of the electrical components, and the aft liquid hydrogen tank. The ET is 

153.8 feet long and has a diameter of 27.6 feet. 

The ET is attached to the orbiter at one forward attachment point and two aft points. In 

the aft attachment area, there are also umbilicals that carry fluids, gases, electrical 

signals and electrical power between the tank and the orbiter. Electrical signals and 

controls between the orbiter and the two solid rocket boosters also are routed through 

those umbilicals. 

Liquid Oxygen Tank

The liquid oxygen tank is an aluminum monocoque structure composed of a fusionwelded assembly of preformed, chem-milled gores, panels, machined fittings and ring 

chords. It operates in a pressure range of 20 to 22 psig. The tank contains anti-slosh 

and anti-vortex provisions to minimize liquid residuals and damp fluid motion. The tank 

feeds into a 17-inch-diameter feed line that conveys the liquid oxygen through the 

intertank, then outside the ET to the aft right-hand ET/orbiter disconnect umbilical. The 

17-inch-diameter feed line permits liquid oxygen to flow at approximately 2,787 pounds 

per second with the SSMEs operating at 104 percent or permits a maximum flow of 

17,592 gallons per minute. The liquid oxygen tank's double-wedge nose cone reduces 

drag and heating, contains the vehicle's ascent air data system (for nine tanks only) and 

serves as a lightning rod. The liquid oxygen tank's volume is 19,563 cubic feet. It is 331 

inches in diameter, 592 inches long and weighs 12,000 pounds empty. 

Intertank

The intertank is a steel/aluminum semimonocoque cylindrical structure with flanges on 

each end for joining the liquid oxygen and liquid hydrogen tanks. The intertank houses 

ET instrumentation components and provides an umbilical plate that interfaces with the 

ground facility arm for purge gas supply, hazardous gas detection and hydrogen gas 

boiloff during ground operations. It consists of mechanically joined skin, stringers and 

machined panels of aluminum alloy. The intertank is vented during flight. The intertank 

contains the forward SRB/ET attach thrust beam and fittings that distribute the SRB 

loads to the liquid oxygen and liquid hydrogen tanks. The intertank is 270 inches long, 

331 inches in diameter and weighs 12,100 pounds. 

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Liquid Hydrogen Tank

The liquid hydrogen tank is an aluminum semimonocoque structure of fusion-welded 

barrel sections, five major ring frames, and forward and aft ellipsoidal domes. Its 

operating pressure range is 32 to 34 psia. The tank contains an anti-vortex baffle and 

siphon outlet to transmit the liquid hydrogen from the tank through a 17-inch line to the 

left aft umbilical. The liquid hydrogen feed line flow rate is 465 pounds per second with 

the SSMEs at 104 percent or a maximum flow of 47,365 gallons per minute. At the 

forward end of the liquid hydrogen tank is the ET/orbiter forward attachment pod strut, 

and at its aft end are the two ET/orbiter aft attachment ball fittings as well as the aft 

SRB/ET stabilizing strut attachments. The liquid hydrogen tank is 331 inches in 

diameter, 1,160 inches long, and has a volume of 53,518 cubic feet and a dry weight of 

29,000 pounds. 

ET Thermal Protection System

The ET thermal protection system consists of sprayed-on foam insulation and 

premolded ablator materials. The system also includes the use of phenolic thermal 

insulators to preclude air liquefaction. Thermal isolators are required for liquid hydrogen 

tank attachments to preclude the liquefaction of air-exposed metallic attachments and 

to reduce heat flow into the liquid hydrogen. 

ET Hardware

Each propellant tank has a vent and relief valve at its forward end. This dual-function 

valve can be opened by ground support equipment for the vent function during 

prelaunch and can open during flight when the ullage (empty space) pressure of the 

liquid hydrogen tank reaches 38 psig or the ullage pressure of the liquid oxygen tank 

reaches 25 psig. 

The liquid oxygen tank contains a separate, pyrotechnically operated, propulsive tumble 

vent valve at its forward end. At separation, the liquid oxygen tumble vent valve is 

opened, providing impulse to assist in the separation maneuver and more positive 

control of the entry aerodynamics of the ET. 

There are eight propellant-depletion sensors, four each for fuel and oxidizer. The fueldepletion sensors are located in the bottom of the fuel tank. The oxidizer sensors are 

mounted in the orbiter liquid oxygen feed line manifold downstream of the feed line 

disconnect. During SSME thrusting, the orbiter general-purpose computers constantly 

compute the instantaneous mass of the vehicle due to the usage of the propellants. 

Normally, main engine cutoff is based on a predetermined velocity; however, if any two 

of the fuel or oxidizer sensors sense a dry condition, the engines will be shut down. 

The locations of the liquid oxygen sensors allow the maximum amount of oxidizer to be 

consumed in the engines, while allowing sufficient time to shut down the engines before 

the oxidizer pumps cavitate (run dry). In addition, 1,100 pounds of liquid hydrogen are 

loaded over and above that required by the 6:1 oxidizer/fuel engine mixture ratio. This 

assures that MECO from the depletion sensors is fuel-rich; oxidizer-rich engine 

shutdowns can cause burning and severe erosion of engine components. 

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STS-92

Four pressure transducers located at the top of the liquid oxygen and liquid hydrogen 

tanks monitor the ullage pressures. 

Each of the two aft external tank umbilical plates mate with a corresponding plate on 

the orbiter. The plates help maintain alignment among the umbilicals. Physical strength 

at the umbilical plates is provided by bolting corresponding umbilical plates together. 

When the orbiter GPCs command external tank separation, the bolts are severed by 

pyrotechnic devices. 

The ET has five propellant umbilical valves that interface with orbiter umbilicals: two for 

the liquid oxygen tank and three for the liquid hydrogen tank. One of the liquid oxygen 

tank umbilical valves is for liquid oxygen, the other for gaseous oxygen. The liquid 

hydrogen tank umbilical has two valves for liquid and one for gas. The intermediatediameter liquid hydrogen umbilical is a recirculation umbilical used only during the liquid 

hydrogen chill-down sequence during prelaunch. 

The ET also has two electrical umbilicals that carry electrical power from the orbiter to 

the tank and the two SRBs and provide information from the SRBs and ET to the 

orbiter. 

A swing-arm-mounted cap to the fixed service structure covers the oxygen tank vent on 

top of the ET during the countdown and is retracted about two minutes before lift-off. 

The cap siphons off oxygen vapor that threatens to form large ice on the ET, thus 

protecting the orbiter's thermal protection system during launch. 

ET Range Safety System

A range safety system provides for dispersing tank propellants if necessary. It includes 

a battery power source, a receiver/decoder, antennas and ordnance. 

Various parameters are monitored and displayed on the flight deck display and control 

panel and are transmitted to the ground. 

The contractor for the external tank is Martin Marietta Aero space, New Orleans, La. 

The tank is manufactured at Michoud, La. Motorola, Inc., Scottsdale, Ariz., is the 

contractor for range safety receivers. 

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Space Shuttle Rendezvous Maneuvers

COMMON SHUTTLE RENDEZVOUS MANEUVERS

OMS-1 (Orbit insertion) - Rarely used ascent abort burn 

OMS-2 (Orbit insertion) - Typically used to circularize the initial orbit following ascent, 

completing orbital insertion. For ground-up rendezvous flights, also considered a 

rendezvous phasing burn

NC (Rendezvous phasing) - Performed to hit a range relative to the target at a future 

time 

NH (Rendezvous height adjust) - Performed to hit a delta-height relative to the target 

at a future time 

NPC (Rendezvous plane change) - Performed to remove planar errors relative to the 

target at a future time

NCC (Rendezvous corrective combination) - First on-board targeted burn in the 

rendezvous sequence. Using star tracker data, it is performed to remove phasing and 

height errors relative to the target at Ti 

Ti (Rendezvous terminal intercept) - Second on-board targeted burn in the 

rendezvous sequence. Using primarily rendezvous radar data, it places the Orbiter on 

a trajectory to intercept the target in one orbit 

MC-1, MC-2, MC-3, MC-4 (Rendezvous midcourse burns) - These on-board 

targeted burns use star tracker and rendezvous radar data to correct the post-Ti 

trajectory in preparation for the final, manual proximity operations phase

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Space Shuttle Solid Rocket Boosters

The two SRBs provide the main thrust to lift the space shuttle off the pad and up to an 

altitude of about 150,000 feet, or 24 nautical miles (28 statute miles). In addition, the 

two SRBs carry the entire weight of the external tank and orbiter and transmit the weight 

load through their structure to the mobile launcher platform. 

Each booster has a thrust (sea level) of approximately 3,300,000 pounds at launch. 

They are ignited after the three space shuttle main engines'' thrust level is verified. The 

two SRBs provide 71.4 percent of the thrust at lift- off and during first-stage ascent. 

Seventy- five seconds after SRB separation, SRB apogee occurs at an altitude of 

approximately 220,000 feet, or 35 nautical miles (41 statute miles). SRB impact occurs 

in the ocean approximately 122 nautical miles (141 statute miles) downrange.

The SRBs are the largest solid- propellant motors ever flown and the first designed for 

reuse. Each is 149.16 feet long and 12.17 feet in diameter.

Each SRB weighs approximately 1,300,000 pounds at launch. The propellant for each 

solid rocket motor weighs approximately 1,100,000 pounds. The inert weight of each 

SRB is approximately 192,000 pounds.

Primary elements of each booster are the motor (including case, propellant, igniter and 

nozzle), structure, separation systems, operational flight instrumentation, recovery 

avionics, pyrotechnics, deceleration system, thrust vector control system and range 

safety destruct system.

Each booster is attached to the external tank at the SRB''s aft frame by two lateral sway 

braces and a diagonal attachment. The forward end of each SRB is attached to the 

external tank at the forward end of the SRB''s forward skirt. On the launch pad, each 

booster also is attached to the mobile launcher platform at the aft skirt by four bolts and 

nuts that are severed by small explosives at lift-off.

During the downtime following the Challenger accident, detailed structural analyses 

were performed on critical structural elements of the SRB. Analyses were primarily 

focused in areas where anomalies had been noted during postflight inspection of 

recovered hardware.

One of the areas was the attach ring where the SRBs are connected to the external 

tank. Areas of distress were noted in some of the fasteners where the ring attaches to 

the SRB motor case. This situation was attributed to the high loads encountered during 

water impact. To correct the situation and ensure higher strength margins during 

ascent, the attach ring was redesigned to encircle the motor case completely (360 

degrees). Previously, the attach ring formed a C and encircled the motor case 270 

degrees.

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Additionally, special structural tests were performed on the aft skirt. During this test 

program, an anomaly occurred in a critical weld between the hold-down post and skin of 

the skirt. A redesign was implemented to add reinforcement brackets and fittings in the 

aft ring of the skirt.

These two modifications added approximately 450 pounds to the weight of each SRB.

The propellant mixture in each SRB motor consists of an ammonium perchlorate 

(oxidizer, 69.6 percent by weight), aluminum (fuel, 16 percent), iron oxide (a catalyst, 

0.4 percent), a polymer (a binder that holds the mixture together, 12.04 percent), and an 

epoxy curing agent (1.96 percent). The propellant is an 11-point star- shaped 

perforation in the forward motor segment and a double- truncated- cone perforation in 

each of the aft segments and aft closure. This configuration provides high thrust at 

ignition and then reduces the thrust by approximately a third 50 seconds after lift-off to 

prevent overstressing the vehicle during maximum dynamic pressure.

The SRBs are used as matched pairs and each is made up of four solid rocket motor 

segments. The pairs are matched by loading each of the four motor segments in pairs 

from the same batches of propellant ingredients to minimize any thrust imbalance. The 

segmented-casing design assures maximum flexibility in fabrication and ease of 

transportation and handling. Each segment is shipped to the launch site on a heavyduty rail car with a specially built cover.

The nozzle expansion ratio of each booster beginning with the STS-8 mission is 7-to-79. 

The nozzle is gimbaled for thrust vector (direction) control. Each SRB has its own 

redundant auxiliary power units and hydraulic pumps. The all-axis gimbaling capability 

is 8 degrees. Each nozzle has a carbon cloth liner that erodes and chars during firing. 

The nozzle is a convergent- divergent, movable design in which an aft pivot- point 

flexible bearing is the gimbal mechanism.

The cone- shaped aft skirt reacts the aft loads between the SRB and the mobile 

launcher platform. The four aft separation motors are mounted on the skirt. The aft 

section contains avionics, a thrust vector control system that consists of two auxiliary 

power units and hydraulic pumps, hydraulic systems and a nozzle extension jettison 

system.

The forward section of each booster contains avionics, a sequencer, forward separation 

motors, a nose cone separation system, drogue and main parachutes, a recovery 

beacon, a recovery light, a parachute camera on selected flights and a range safety 

system.

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Each SRB has two integrated electronic assemblies, one forward and one aft. After 

burnout, the forward assembly initiates the release of the nose cap and frustum and 

turns on the recovery aids. The aft assembly, mounted in the external tank/SRB attach 

ring, connects with the forward assembly and the orbiter avionics systems for SRB 

ignition commands and nozzle thrust vector control. Each integrated electronic 

assembly has a multiplexer/ demultiplexer, which sends or receives more than one 

message, signal or unit of information on a single communication channel.

Eight booster separation motors (four in the nose frustum and four in the aft skirt) of 

each SRB thrust for 1.02 seconds at SRB separation from the external tank. Each solid 

rocket separation motor is 31.1 inches long and 12.8 inches in diameter.

Location aids are provided for each SRB, frustum/ drogue chutes and main parachutes. 

These include a transmitter, antenna, strobe/ converter, battery and salt water switch 

electronics. The location aids are designed for a minimum operating life of 72 hours and 

when refurbished are considered usable up to 20 times. The flashing light is an 

exception. It has an operating life of 280 hours. The battery is used only once.

The SRB nose caps and nozzle extensions are not recovered.

The recovery crew retrieves the SRBs, frustum/ drogue chutes, and main parachutes. 

The nozzles are plugged, the solid rocket motors are dewatered, and the SRBs are 

towed back to the launch site. Each booster is removed from the water, and its 

components are disassembled and washed with fresh and deionized water to limit salt 

water corrosion. The motor segments, igniter and nozzle are shipped back to Thiokol for 

refurbishment.

Each SRB incorporates a range safety system that includes a battery power source, 

receiver/ decoder, antennas and ordnance.

HOLD-DOWN POSTS

Each solid rocket booster has four hold- down posts that fit into corresponding support 

posts on the mobile launcher platform. Hold- down bolts hold the SRB and launcher 

platform posts together. Each bolt has a nut at each end, but only the top nut is 

frangible. The top nut contains two NASA standard detonators, which are ignited at solid 

rocket motor ignition commands.

When the two NSDs are ignited at each hold- down, the hold- down bolt travels 

downward because of the release of tension in the bolt (pretensioned before launch), 

NSD gas pressure and gravity. The bolt is stopped by the stud deceleration stand, 

which contains sand. The SRB bolt is 28 inches long and is 3.5 inches in diameter. The 

frangible nut is captured in a blast container.

54
STS-92

The solid rocket motor ignition commands are issued by the orbiter''s computers 

through the master events controllers to the hold- down pyrotechnic initiator controllers 

on the mobile launcher platform. They provide the ignition to the hold- down NSDs. The 

launch processing system monitors the SRB hold- down PICs for low voltage during the 

last 16 seconds before launch. PIC low voltage will initiate a launch hold.

SRB IGNITION

SRB ignition can occur only when a manual lock pin from each SRB safe and arm 

device has been removed. The ground crew removes the pin during prelaunch 

activities. At T minus five minutes, the SRB safe and arm device is rotated to the arm 

position. The solid rocket motor ignition commands are issued when the three SSMEs 

are at or above 90-percent rated thrust, no SSME fail and/or SRB ignition PIC low 

voltage is indicated and there are no holds from the LPS.

The solid rocket motor ignition commands are sent by the orbiter computers through the 

MECs to the safe and arm device NSDs in each SRB. A PIC single-channel capacitor 

discharge device controls the firing of each pyrotechnic device. Three signals must be 

present simultaneously for the PIC to generate the pyro firing output. These signalsarm, fire 1 and fire 2-originate in the orbiter general- purpose computers and are 

transmitted to the MECs. The MECs reformat them to 28-volt dc signals for the PICs. 

The arm signal charges the PIC capacitor to 40 volts dc (minimum of 20 volts dc).

The fire 2 commands cause the redundant NSDs to fire through a thin barrier seal down 

a flame tunnel. This ignites a pyro booster charge, which is retained in the safe and arm 

device behind a perforated plate. The booster charge ignites the propellant in the igniter 

initiator; and combustion products of this propellant ignite the solid rocket motor initiator, 

which fires down the length of the solid rocket motor igniting the solid rocket motor 

propellant.

The GPC launch sequence also controls certain critical main propulsion system valves 

and monitors the engine- ready indications from the SSMEs. The MPS start commands 

are issued by the onboard computers at T minus 6.6 seconds (staggered start- engine 

three, engine two, engine one- all approximately within 0.25 of a second), and the 

sequence monitors the thrust buildup of each engine. All three SSMEs must reach the 

required 90-percent thrust within three seconds; otherwise, an orderly shutdown is 

commanded and safing functions are initiated.

Normal thrust buildup to the required 90-percent thrust level will result in the SSMEs 

being commanded to the lift- off position at T minus three seconds as well as the fire 1 

command being issued to arm the SRBs. At T minus three seconds, the vehicle base 

bending load modes are allowed to initialize (movement of approximately 25.5 inches 

measured at the tip of the external tank, with movement towards the external tank).

55
STS-92

At T minus zero, the two SRBs are ignited, under command of the four onboard 

computers; separation of the four explosive bolts on each SRB is initiated (each bolt is 

28 inches long and 3.5 inches in diameter); the two T-0 umbilicals (one on each side of 

the spacecraft) are retracted; the onboard master timing unit, event timer and mission 

event timers are started; the three SSMEs are at 100 percent; and the ground launch 

sequence is terminated.

The solid rocket motor thrust profile is tailored to reduce thrust during the maximum 

dynamic pressure region.

ELECTRICAL POWER DISTRIBUTION

Electrical power distribution in each SRB consists of orbiter- supplied main dc bus 

power to each SRB via SRB buses A, B and C. Orbiter main dc buses A, B and C 

supply main dc bus power to corre sponding SRB buses A, B and C. In addition, orbiter 

main dc bus C supplies backup power to SRB buses A and B, and orbiter bus B 

supplies backup power to SRB bus C. This electrical power distribution arrangement 

allows all SRB buses to remain powered in the event one orbiter main bus fails.

The nominal dc voltage is 28 volts dc, with an upper limit of 32 volts dc and a lower limit 

of 24 volts dc.

HYDRAULIC POWER UNITS

There are two self- contained, independent HPUs on each SRB. Each HPU consists of 

an auxiliary power unit, fuel supply module, hydraulic pump, hydraulic reservoir and 

hydraulic fluid manifold assembly. The APUs are fueled by hydrazine and generate 

mechanical shaft power to a hydraulic pump that produces hydraulic pressure for the 

SRB hydraulic system. The two separate HPUs and two hydraulic systems are located 

on the aft end of each SRB between the SRB nozzle and aft skirt. The HPU 

components are mounted on the aft skirt between the rock and tilt actuators. The two 

systems operate from T minus 28 seconds until SRB separation from the orbiter and 

external tank. The two independent hydraulic systems are connected to the rock and tilt 

servoactuators.

The APU controller electronics are located in the SRB aft integrated electronic 

assemblies on the aft external tank attach rings.

The APUs and their fuel systems are isolated from each other. Each fuel supply module 

(tank) contains 22 pounds of hydrazine. The fuel tank is pressurized with gaseous 

nitrogen at 400 psi, which provides the force to expel (positive expulsion) the fuel from 

the tank to the fuel distribution line, maintaining a positive fuel supply to the APU 

throughout its operation.

56
STS-92

The fuel isolation valve is opened at APU startup to allow fuel to flow to the APU fuel 

pump and control valves and then to the gas generator. The gas generator''s catalytic 

action decomposes the fuel and creates a hot gas. It feeds the hot gas exhaust product 

to the APU two- stage gas turbine. Fuel flows primarily through the startup bypass line 

until the APU speed is such that the fuel pump outlet pressure is greater than the 

bypass line''s. Then all the fuel is supplied to the fuel pump.

The APU turbine assembly provides mechanical power to the APU gearbox. The 

gearbox drives the APU fuel pump, hydraulic pump and lube oil pump. The APU lube oil 

pump lubricates the gearbox. The turbine exhaust of each APU flows over the exterior 

of the gas generator, cooling it, and is then directed overboard through an exhaust duct.

When the APU speed reaches 100 percent, the APU primary control valve closes, and 

the APU speed is controlled by the APU controller electronics. If the primary control 

valve logic fails to the open state, the secondary control valve assumes control of the 

APU at 112-percent speed.

Each HPU on an SRB is connected to both servoactuators on that SRB. One HPU 

serves as the primary hydraulic source for the servoactuator, and the other HPU serves

as the secondary hydraulics for the servoactuator. Each sevoactuator has a switching 

valve that allows the secondary hydraulics to power the actuator if the primary hydraulic 

pressure drops below 2,050 psi. A switch contact on the switching valve will close when 

the valve is in the secondary position. When the valve is closed, a signal is sent to the 

APU controller that inhibits the 100-percent APU speed control logic and enables the 

112-percent APU speed control logic. The 100-percent APU speed enables one 

APU/HPU to supply sufficient operating hydraulic pressure to both servoactuators of 

that SRB.

The APU 100-percent speed corresponds to 72,000 rpm, 110-percent to 79,200 rpm, 

and 112-percent to 80,640 rpm.

The hydraulic pump speed is 3,600 rpm and supplies hydraulic pressure of 3,050, plus 

or minus 50, psi. A high- pressure relief valve provides overpressure protection to the 

hydraulic system and relieves at 3,750 psi.

The APUs/HPUs and hydraulic systems are reusable for 20 missions.

THRUST VECTOR CONTROL

Each SRB has two hydraulic gimbal servoactuators: one for rock and one for tilt. The 

servoactuators provide the force and control to gimbal the nozzle for thrust vector 

control.

57
STS-92

The space shuttle ascent thrust vector control portion of the flight control system directs 

the thrust of the three shuttle main engines and the two SRB nozzles to control shuttle 

attitude and trajectory during lift- off and ascent. Commands from the guidance system 

are transmitted to the ATVC drivers, which transmit signals proportional to the 

commands to each servoactuator of the main engines and SRBs. Four independent 

flight control system channels and four ATVC channels control six main engine and four 

SRB ATVC drivers, with each driver controlling one hydraulic port on each main and 

SRB servoactuator.

Each SRB servoactuator consists of four independent, two- stage servovalves that 

receive signals from the drivers. Each servovalve controls one power spool in each 

actuator, which positions an actuator ram and the nozzle to control the direction of 

thrust.

The four servovalves in each actuator provide a force- summed majority voting 

arrangement to position the power spool. With four identical commands to the four 

servovalves, the actuator force-sum action prevents a single erroneous command from 

affecting power ram motion. If the erroneous command persists for more than a 

predetermined time, differential pressure sensing activates a selector valve to isolate 

and remove the defective servovalve hydraulic pressure, permitting the remaining 

channels and servovalves to control the actuator ram spool.

Failure monitors are provided for each channel to indicate which channel has been 

bypassed. An isolation valve on each channel provides the capability of resetting a 

failed or bypassed channel.

Each actuator ram is equipped with transducers for position feedback to the thrust 

vector control system. Within each servoactuator ram is a splashdown load relief 

assembly to cushion the nozzle at water splashdown and prevent damage to the nozzle 

flexible bearing.

SRB RATE GYRO ASSEMBLIES

Each SRB contains two RGAs, with each RGA containing one pitch and one yaw gyro. 

These provide an output proportional to angular rates about the pitch and yaw axes to 

the orbiter computers and guidance, navigation and control system during first- stage 

ascent flight in conjunction with the orbiter roll rate gyros until SRB separation. At SRB 

separation, a switchover is made from the SRB RGAs to the orbiter RGAs.

The SRB RGA rates pass through the orbiter flight aft multiplexers/ demultiplexers to 

the orbiter GPCs. The RGA rates are then mid-value- selected in redundancy 

management to provide SRB pitch and yaw rates to the user software. The RGAs are 

designed for 20 missions.

58
STS-92

SRB SEPARATION

SRB separation is initiated when the three solid rocket motor chamber pressure 

transducers are processed in the redundancy management middle value select and the 

head- end chamber pressure of both SRBs is less than or equal to 50 psi. A backup cue 

is the time elapsed from booster ignition.

The separation sequence is initiated, commanding the thrust vector control actuators to 

the null position and putting the main propulsion system into a second-stage 

configuration (0.8 second from sequence initialization), which ensures the thrust of each 

SRB is less than 100,000 pounds. Orbiter yaw attitude is held for four seconds, and 

SRB thrust drops to less than 60,000 pounds.

The SRBs separate from the external tank within 30 milliseconds of the ordnance firing 

command.

The forward attachment point consists of a ball (SRB) and socket (ET) held together by 

one bolt. The bolt contains one NSD pressure cartridge at each end. The forward 

attachment point also carries the range safety system cross-strap wiring connecting 

each SRB RSS and the ET RSS with each other.

The aft attachment points consist of three separate struts: upper, diagonal and lower. 

Each strut contains one bolt with an NSD pressure cartridge at each end. The upper 

strut also carries the umbilical interface between its SRB and the external tank and on 

to the orbiter.

There are four booster separation motors on each end of each SRB. The BSMs 

separate the SRBs from the external tank. The solid rocket motors in each cluster of 

four are ignited by firing redundant NSD pressure cartridges into redundant confined 

detonating fuse manifolds.

The separation commands issued from the orbiter by the SRB separation sequence 

initiate the redundant NSD pressure cartridge in each bolt and ignite the BSMs to effect 

a clean separation.

59
STS-92

Space Shuttle Super Light Weight Tank (SLWT)

The super lightweight external tank (SLWT) made its first Shuttle flight June 2, on 

mission STS-91. The SLWT is 7,500 pounds lighter than the standard external tank. 

The lighter weight tank will allow the shuttle to deliver International Space Station 

elements (such as the service module) into the proper orbit. 

The SLWT is the same size as the previous design. But the liquid hydrogen tank and 

the liquid oxygen tank are made of aluminum lithium, a lighter, stronger material than 

the metal alloy used for the Shuttle's current tank. The tank's structural design has 

also been improved, making it 30% stronger and 5% less dense. 

The SLWT, like the standard tank, is manufactured at Michoud Assembly, near New 

Orleans, Louisiana, by Lockheed Martin. 

The 154-foot-long external tank is the largest single component of the space shuttle. It 

stands taller than a 15-story building and has a diameter of about 27 feet. The external 

tank holds over 530,000 gallons of liquid hydrogen and liquid oxygen in two separate 

tanks. The hydrogen (fuel) and liquid oxygen (oxidizer) are used as propellants for the 

Shuttle's three main engines. 

60
ACRONYMS AND ABBREVIATIONS

A/D Analog to Digital

AC Assembly Contingency

ACBM Active Common Berthing Mechanism

ACBSP Assembly Contingency Baseband Signal Processor

ACRFG Assembly Contingency RF Group

ACS Attitude Control System

AD Active Device

AFD Aft Flight Deck

AGIT Antenna Group Interface Tube

AKA Active Keel Assembly

AOH Assembly Operations Handbook

APAS Androgynous Peripheral Attachment System

APCU Assembly Power Conversion Unit

APFR Articulating Portable Foot Restraint

AR Atmosphere Revitalization

ASCR Assured Safe Crew Return

AUAIU Audio Interface Unit

B/U Backup

BIT Built-In Test

BITE Built-In Test Equipment

BIU Bus Interface Unit

BMR Body-Mounted Radiator

BSP Baseband Signal Processor

C&C Command and Control

C&DH Command and Data Handling

C&T Communications and Tracking

C&W Caution and Warning

CBM Common Berthing Mechanism

CC Central Computer

CCSDS Consultative Committee for Space Data System

CCTV Closed Circuit Television

CCW Counterclockwise

CDRA Carbon Dioxide Removal Assembly

CHIA Cargo Handling Interface Adapter

CL Close

CMG Control Moment Gyro

COTS Commercial-off-the-Shelf

CP Coldplate

CTP Command/Telemetry Processor

DAP Digital Auto Pilot

DCCU DC to DC Connector Unit

DDCU Direct Current-to-Direct Current/Converter Unit

DMS Data Management Subsystem

DSCU Docking System Control Unit

61
EATC External Active Thermal Control

ECLSS Environmental Control and Life Support System

ECOMM Early Communications System

EDAC Error Detection and Correction

EEATCS Early External Active Thermal Control System

EEPROM Electrically Erasable Programmable Read Only Memory

EMU Extravehicular Maneuvering Unit

ENA Enable

EPCS Early Portable Computer System

EPS Electrical Power Subsystem

ERSU External Remote Sensing Unit

ESA European Space Agency

ETSD EVA Tool Stowage Device

EVA Extravehicular Activity

EVAS Extravehicular Activity Support

EVR Extravehicular Robotics

FBOLT Final Preload Bolt

FD Flight Day

FDF Flight Data File

FDIR Fault Detection, Isolation, and Recovery

FDS Fire Detection and Suppression

FGB Functional Cargo Block

FRGF Flight Releasable Grapple Fixture

FRM CTR Frame Counter

GFE Government-Furnished Equipment

GLONASS Global Navigational Satellite System

GN&C Guidance, Navigation, and Control

GPC General Purpose Computer

HCA Hollow Cathode Assembly

HEI Hernandez Engineering, Inc.

HGA High-Gain Antenna

HRD High-Rate Data

HRFM High-Rate Frame Multiplexer

HRM High-Rate Modem

HX Heat Exchanger

I/F Interface

IATC Internal Active Thermal Control

IBOLT Initial Preload Bolt

ICD Interface Control Document

IDA Initialization Diode Assembly

IMV Intermodule Ventilation

INH Inhibit

INP Input

ISS International Space Station

ITS Integrated Truss Segment

IVA Intravehicular Activity

IWIS Internal Wireless Instrumentation System

62
JOP Joint Operations Panel

JSC Lyndon B. Johnson Space Center

L/L Liquid to Liquid

LAN Land Area Network

LGA Low-Gain Antenna

LSM Life Support Module

LT Low Temperature

LTA Launch to Activation

LVLH Local Vertical/Local Horizontal

M/OD Micrometeorite/Orbital Debris

MBM Manual Berthing Mechanism

Modified Berthing Mechanism

MCC Mission Control Center

MCC-H Mission Control Center - Houston

MCC-M Mission Control Center - Moscow

MCDS Multi-CRT Display System

MCS Motion Control System

MDM Multiplexer/Demultiplexer

MIL-STD Military Standard

MLI Multilayer Insulation

MOD Mission Operations Directorate

MPEV Manual Pressure Equalization Valve

MPM Manipulator Positioning Mechanism

MRK Moisture Removal Kit

MRL Manipulator Retention Latch

MT Moderate Temperature

NASA National Aeronautics and Space Administration

NCC Nominal Course Correction

NCS Node Control Software

NCU Network Control Unit

NPRV Negative Pressure Relief Valve

NIU Network Interface Unit

OCA Orbiter Communications Adapter

OCC Onboard Computer System

ODF Operations Data File

ODS Orbiter Docking System

ODTS

OIU Orbiter Interface Unit

OMS Onboard Measuring Subsystem

OP Open

ORU On-Orbit Replaceable Unit

OS On-Orbit Segment

OUT Output

63
PCBM Passive Common Berthing Mechanism

PCM Pulse Code Modulation

PCMMU Pulse Code Modulation Master Unit

PCN Page Change Notice

PCS Portable Computer System

PCU Plasma Contactor Unit

PD Physical Device

PDU Power Distribution Unit

PFR Portable Foot Restraint

PFRWS Portable Foot Restraint Workstation Stanchion

PGSC Payload and General Support Computer

PGT Pistol Grip Tool

PMA Pressurized Mating Adapter

PRI Primary

psid Pounds Per Square Inch - Differential

PRLA Payload Retention Latch Assembly

PTCS Passive Thermal Control System

RACU Russian American Conversion Unit

RCS Remote 

RF Radio Frequency

RFG Radio Frequency Group

RMS Remote Manipulator System

ROS Russian On-Orbit Segment

RPC Remote Power Connector

RPCM Remote Power Controller Mode

RPDA Remote Power Distribution Assembly

RR Rendezvous Radar

RS Russian Segment

RSA Russian Space Agency

RSU Remote Sensing Unit

RTAS Rocketdyne Truss Attachment System

RTD Resistive Temperature Device

RTS Radiotelemetry System

S&M Structures and Mechanisms

SAG S-Band Antenna Group

SASA S-Band Antenna Subassembly

SEC Secondary

SECAM

SFOC Space Flight Operations Contract

SGANT Space-to-Ground Antenna

SGTRC Space-to-Ground Transmit/Receive Controller

SHF S-Band High Frequency

SLP Spacelab-Pallet

SM Service Module

SMCC Shuttle Mission Control CenterR

SODF Station Operations Data File

SPDA Secondary Power Distribution Assembly

SRMS Shuttle Remote Manipulator System

SSP Space Shuttle Program

64
Standard Switch Panel

SSRMS Space Station Remote Manipulator System

STBD Starboard

SWIS Shuttle-Based Wireless Instrumentation System

TC Terminal Computer

TCS Thermal Control Subsystem

TDRS Tracking and Data Relay Satellite

TDRSS Tracking and Data Relay Satellite System

TEA Torque Equilibrium Attitude

THC Temperature and Humidity Control

Ti Terminal Phase Initiation

TRC Transmit/Receive Control

TTCS Telephone-Telegraph Communication Subsystem

UHF Ultra High Frequency

US

USA United Space Alliance

USL United States Laboratory

USOS United States On-Orbit Segment

USW Ultra Short Wave

VBSP Video Baseband Signal Processor

VRCS Vector/Vernier Reaction Control System

VTS Video-Teleconferencing System

WIF Worksite Interface

WIS Wireless Instrumentation System

WM Waste Management

WRM Water Recovery and Management

WS Worksite

WSGT White Sands Ground Terminal

XCVR Transmitter/Receiver

65
STS-92

Mission Benefits

Research on the International Space Station

The International Space Station represents a quantum leap in our capability to conduct 

research on orbit. It will serve as a laboratory for exploring basic questions in a variety 

of disciplines, and as a testbed and springboard for exploration. Research on the ISS 

will include commercial, science, and engineering research in the following areas:

Advanced Human Support Technology: Researchers develop technologies, 

systems, and procedures to enable safe and efficient human exploration and 

development of space.

Long Term Benefits: Reduce the cost of space travel while enhancing safety; develop 

small, low power monitoring and sensing technologies with applications in 

environmental monitoring in space and on Earth; and develop advanced waste 

processing and agricultural technologies with applications in space and on Earth.

Biomedical Research and Countermeasures: Researchers seek to understand and 

control the effects of the space environment on space travelers (e.g. muscle atrophy, 

bone loss, fluid shifts, . . .).

Long Term Benefits: Enhance the safety of space travel; develop methods to keep 

humans healthy in low-gravity environments; and advance new fields of research in the 

treatment of diseases,

Fundamental Biology: Scientists study gravity's influence on the evolution, 

development, growth, and internal processes of plants and animals. Their results 

expand fundamental knowledge that will benefit medical, agricultural, and other 

industries.

Long Term Benefits: Advance understanding of cell, tissue, and animal behavior;use of 

plants as sources of food and oxygen for exploration; improved plants for agricultural 

and forestry,

Biotechnology: Microgravity allows researchers to grow three-dimensional tissues that 

have characteristics more similar to tissues in the body than has ever been previously 

available and to produce superior protein crystals for drug development. 

Long Term Benefits: Culture realistic tissue for use in research (cancerous tumors, 

organ pieces); and provide information to design a new class of drugs to target specific 

proteins and cure specific diseases, 

66
Fluid Physics: The behavior of fluids is profoundly influenced by gravity. Researchers 

use gravity as an experimental variable to explain and model fluid behavior in systems 

on Earth and in space.

Long Term Benefits: Improved spacecraft systems designs for safety and efficiency; 

better understanding of soil behavior in earthquake conditions; and improved 

mathematical models for designing fluid handling systems for powerplants, refineries 

and innumerable other industrial applications 

Materials Science: Researchers use low gravity to advance our understanding of the 

relationships among the structure, processing and properties of materials. In low gravity, 

differences in weight of liquids used to form materials do not interfere with the ability to 

mix these materials opening the door to a whole new world of composite materials.

Long Term Benefits: Advance understanding of processes for manufacturing 

semiconductors, metals, ceramics, polymers, and other materials; and determine 

fundamental physical properties of molten metal, semiconductors, and other materials 

with precision impossible on Earth.

Combustion Science: The removal of gravity allows scientists to simplify the study of 

complex combustion (burning) processes. Since combustion is used to produce 85 

percent of Earth's energy, even small improvements in efficiency and reduction of soot 

production (a major source of pollution on earth) will have large environmental and 

economic benefits. 

Long Term Benefits: Enhance efficiency of combustion processes; enhance fire 

detection and safety on Earth and in Space; and improve control of combustion 

emissions and pollutants 

Fundamental Physics: Scientists use the low gravity and low temperature 

environment to slow down reactions allowing them to test fundamental theories of 

physics with degrees of accuracy that far exceed the capacity of Earthbound science.

Long Term Benefits: Challenge and expand theories of how matter organizes as it 

changes state (important in understanding superconductivity); test fundamental theories 

in physics with precision beyond the capacity of Earth-bound science; and potential for 

improved magnetic materials 

Earth Science and Space Science: Space Station will be a unique platform with 

multiple exterior attach points from which to observe the Earth and the Universe.

Long Term Benefits: Space Scientists will use the location above the atmosphere to 

collect and search for cosmic rays, cosmic dust, antimatter and "dark" matter. Earth 

Scientists can obtain global profiles of aerosols, ozone, water vapor, and oxides in order 

to determine their role in climatological processes and take advantage of the longevity 

of ISS to observe global changes over many years.

67
STS-92

Mission Benefits

The Space Shuttle, Still Young at 100

Safer, More Capable and More Reliable Than Ever Before

1. The shuttle has amassed an amazing array of accomplishments in the past 20 years:

It has launched 3 million pounds of cargo and almost 600 (596 counting the STS-92 

crew) passengers and pilots.

The shuttle fleet has cumulatively spent almost three years in flight (2 years, 336 days, 

14 hours, 47 minutes, 59 seconds to be exact, not including the planned duration of 

STS-92).

The fleet has amassed more than 15 years of passenger-hours in space.

Over 850 payloads have flown, including hundreds of individual experiments.

The Shuttle has deployed more than 60 payloads and retrieved more than two dozen.

The shuttle has traveled more than 345 million miles (not including STS-92) and 

completed more than 13,500 orbits of Earth (exactly 13,573, not including STS-92).

2. The shuttle has enabled us to make unprecedented discoveries about ourselves, our 

planet and our universe.

The shuttle has supported two space stations, made three maintenance flights to the 

Hubble Space Telescope and launched planetary missions to study Jupiter, Venus and 

the Sun. All of NASA's 

68
STS-92

Media Assistance

NASA Television Transmission

NASA Television is available through the GE2 satellite system which is located on 

Transponder 9C, at 85 degrees west longitude, frequency 3880.0 MHz, audio 6.8 MHz. 

The schedule for television transmissions from the orbiter and for mission briefings will 

be available during the mission at Kennedy Space Center, FL; Marshall Space Flight 

Center, Huntsville, AL; Dryden Flight Research Center, Edwards, CA; Johnson Space 

Center, Houston, TX; and NASA Headquarters, Washington, DC. The television 

schedule will be updated to reflect changes dictated by mission operations.

Status Reports

Status reports on countdown and mission progress, on-orbit activities and landing 

operations will be produced by the appropriate NASA newscenter

Briefings

A mission press briefing schedule will be issued before launch. During the mission, 

status briefings by a flight director or mission operations representative and when 

appropriate, representatives from the payload team, will occur at least once each day. 

The updated NASA television schedule will indicate when mission briefings are planned. 

Internet Information

Information is available through several sources on the Internet. The primary source for 

mission information is the NASA Human Space Flight Web, part of the World Wide 

Web. This site contains information on the crew and its mission and will be updated 

regularly with status reports, photos and video clips throughout the flight. The NASA 

Shuttle Web's address is:

http://spaceflight.nasa.gov

General information on NASA and its programs is available through the NASA Home 

Page and the NASA Public Affairs Home Page:

http://www.nasa.gov

or

http://www.nasa.gov/newsinfo/index.html

Shuttle Pre-Launch Status Reports

http://www-pao.ksc.nasa.gov/kscpao/status/stsstat/current.htm

69
STS-92

Information on other current NASA activities is available through the Today@NASA 

page:

http://www.nasa.gov/today/index.html

The NASA TV schedule is available from the NTV Home Page:

http://spaceflight.nasa.gov/realdata/nasatv/schedule.html

Resources for educators can be found at the following address:

http://education.nasa.gov

Access by Compuserve

Users with Compuserve accounts can access NASA press releases by typing "GO 

NASA" (no quotes) and making a selection from the categories offered.

70
STS-92

Media Contacts

NASA PAO Contacts

Eileen Hawley 281 483-5111

NASA, Johnson Space Center

Houston, TX

Astronauts/Mission Operations

eileen.hawley1@jsc.nasa.gov

Kari Kelley Allen 281 336-4844

The Boeing Company

Houston, TX

International Space Station 

kari.k.allen@boeing.com

Bruce Buckingham 321 867-2468

NASA, Kennedy Space Center

Kennedy Space Center, FL

Launch Operations

bruce.buckingham-1@kmail.ksc.nasa.gov

Jack King 321 861-4358

United Space Alliance

Kennedy Space Center, FL

Shuttle Processing

KingJW@usafoo.unitedspacealliance.com

Beth Hill 714 372-4736

The Boeing Company

Huntington Beach, CA

Space Shuttle

sara.b.hill@boeing.com

Dan Beck 818 586-4572

Boeing Rocketdyne

Canoga Park, CA

Space Shuttle Main Engines

daniel.c.beck@boeing.com

71
STS-92

Dwayne Brown 202 358-1726

NASA Headquarters

Washington, DC

Space Shuttle/Space Station 

Policy

dwayne.brown@hq.nasa.gov

Kyle Herring 281 483-5111

NASA, Johnson Space Center

Houston, TX

Space Station Operations

kyle.j.herring1@jsc.nasa.gov

72
Table prepared by Richard W. Orloff, 02/2001 

SHUTTLE FLIGHTS AS OF OCTOBER 2000 

99 TOTAL FLIGHTS OF THE SHUTTLE SYSTEM -- 74 SINCE RETURN TO FLIGHT 

STS-103 

12/19/99 - 12/27/99 

STS-93 

07/23/99 - 07/27/99 

STS-96 

05/27/99 - 06/06/99 

STS-90 

04/17/98 - 05/03/98 

STS-95 

10/29/98 - 11/07/98 

STS-87 

11/19/97 - 12/05/97 

STS-91 

06/02/09 - 06/12/98 

STS-94 

07/01/97 - 07/17/97 

STS-85 

08/07/97 - 08/19/97 

STS-83 

04/04/97 - 04/08/97 

STS-82 

02/11//97 - 02/21/97 

STS-106 

09/08/00 - 09/20/00 

STS-80 

11/19/96 - 12/07/96 

STS-70 

07/13/95 - 07/22/95 

STS-101 

05/19/00 - 05/29/00 

STS-78 

06/20/96 - 07/07/96 

STS-63 

02/03/95 - 02/11/95 

STS-86 

09/25/97 - 10/06/97 

STS-75 

02/22/96 - 03/09/96 

STS-64 

09/09/94 - 09/20/94 

STS-84 

05/15/97 - 05/24/97 

STS-73 

10/20/95 - 11/05/95 

STS-60 

02/03/94 - 02/11/94 

STS-81 

01/12/97 - 01/22/97 

STS-65 

07/08/94 - 07/23/94 

STS-51 

09/12/93 - 09/22/93 

STS-79 

09/16/96 - 09/26/96 

STS-62 

03/04/94 - 03/18/94 

STS-56 

04/08/83 - 04/17/93 

STS-76 

03/22/96 - 03/31/96 

STS-58 

10/18/93 - 11/01/93 

STS-53 

12/02/92 - 12/09/92 

STS-74 

11/12/95 - 11/20/95 

STS-55 

04/26/93 - 05/06/93 

STS-42 

01/22/92 - 01/30/92 

STS-71 

06/27/95 - 07/07/95 

STS-99 

02/11/00 - 02/22/00 

STS-52 

10/22/92 - 11/01/92 

STS-48 

09/12/91 - 09/18/91 

STS-66 

11/03/94 - 11/14/94 

STS-88 

12/04/98 - 12/15/98 

STS-50 

06/25/92 - 07/09/92 

STS-39 

04/28/91 - 05/06/91 

STS-46 

07/31/92 - 08/08/92 

STS-89 

01/22/98 - 01/31/98 

STS-40 

06/05/91 - 06/14/91 

STS-41 

10/06/90 - 10/10/90 

STS-45 

03/24/92 - 04/02/92 

STS-77 

05/19/96 - 05/29/96 

STS-35 

12/02/90 - 12/10/90 

STS-51L 

01/28/86 

STS-31 

04/24/90 - 04/29/90 

STS-44 

11/24/91 - 12/01/91 

STS-72 

01/11/96 - 11/20/96 

STS-32 

01/09/90 - 01/20/90 

STS-61A 

10/30/85 - 11/06/85 

STS-33 

11/22/89 - 11/27/89 

STS-43 

08/02/91 - 08/11/91 

STS-69 

09/07/95 - 09/18/95 

STS-28 

08/08/89 - 08/13/89 

STS-51F 

07/29/85 - 08/06/85 

STS-29 

03/13/89 - 03/18/89 

STS-37 

04/05/91 - 04/11/91 

STS-67 

03/02/95 - 03/18/95 

STS-61C 

01/12/86 - 01/18/86 

STS-51B 

04/29/85 - 05/06/85 

STS-26 

09/29/88 - 10/03/88 

STS-38 

11/15/90 - 11/20/90 

STS-68 

09/30/94 - 10/11/94 

STS-9 

11/28/83 - 12/08/83 

STS-41G 

10/05/84 - 10/13/84 

STS-51-I 

08/27/85 - 09/03/85 

STS-36 

02/28/90 - 03/04/90 

STS-59 

04/09/94 – 04/20/94 

STS-5 

11/11/82 - 11/16/82 

STS-41C 

04/06/84 - 04/13/84 

STS-51G 

06/17/85 - 06/24/85 

STS-34 

10/18/89 - 10/23/89 

STS-61 

12/02/93 - 12/13/93 

STS-4 

06/27/82 - 07/04/82 

STS-41B 

02/03/84 - 02/11/84 

STS-51D 

04/12/85 - 04/19/85 

STS-30 

05/04/89 - 05/08/89 

STS-57 

06/21/93 - 07/01/93 

STS-3 

03/22/82 - 03/30/82 

STS-8 

08/30/83 - 09/05/83 

STS-51C 

01/24/85 - 01/27/85 

STS-27 

12/02/88 - 12/06/88 

STS-54 

01/13/93 - 01/19/93 

STS-2 

11/12/81 - 11/14/81 

STS-7 

06/18/83 - 06/24/83 

STS-51A 

11/08/84 - 11/16/84 

STS-61B 

11/26/85 - 12/03/85 

STS-47 

09/12/92 - 09/20/92 

STS-1 

04/12/81 - 04/14/81 

STS-6 

04/04/83 - 04/09/83 

STS-41D 

08/30/84 - 09/05/84 

STS-51J 

10/03/85 - 10/07/85 

STS-49 

05/07/92 - 05/16/92 

 

OV-102 

Columbia 

(26 flights) 

OV-099 

Challenger 

(10 flights) 

OV-103 

Discovery 

(27 flights) 

OV-104 

Atlantis 

(22 flights) 

OV-105 

Endeavour 

(14 flights)