STS-121 Press Kit 
August 2006 CONTENTS i

CONTENTS 

Section Page

STS-115 MISSION OVERVIEW: SPACE STATION ASSEMBLY RESUMES................................ 1

STS-115 TIMELINE OVERVIEW ............................................................................................... 10

MISSION PRIORITIES............................................................................................................. 12

LAUNCH AND LANDING ........................................................................................................... 14

LAUNCH............................................................................................................................................... 14

ABORT-TO-ORBIT (ATO)...................................................................................................................... 14

TRANSATLANTIC ABORT LANDING (TAL) ............................................................................................. 1 4

RETURN-TO-LAUNCH-SITE (RTLS)....................................................................................................... 14

ABORT ONCE AROUND (AOA)............................................................................................................... 14

LANDING ............................................................................................................................................. 14

MISSION PROFILE................................................................................................................... 15

STS-115 ATLANTIS CREW ...................................................................................................... 16

MISSION PERSONNEL ............................................................................................................. 21

RENDEZVOUS AND DOCKING .................................................................................................. 22

UNDOCKING, SEPARATION AND DEPARTURE ....................................................................................... 25

SPACEWALKS ......................................................................................................................... 26

PAYLOAD OVERVIEW .............................................................................................................. 32

EXPERIMENTS ......................................................................................................................... 51

SHUTTLE REFERENCE DATA .................................................................................................... 54

ACRONYMS AND ABBREVIATIONS ......................................................................................... 67

MEDIA ASSISTANCE ............................................................................................................... 82

PUBLIC AFFAIRS CONTACTS .................................................................................................. 83
August 2006 MISSION OVERVIEW 1

STS-115 MISSION OVERVIEW: 

SPACE STATION ASSEMBLY RESUMES 

Configuration of the International Space Station at the completion of STS‐115

For the first time since late 2002, assembly of

the International Space Station (ISS) will

resume on the STS‐115 mission.

Six astronauts will be launched on the Space

Shuttle Atlantis on the 116th shuttle mission and

the 27th flight of Atlantis to deliver and install

the 17‐and‐a‐half ton P3 / P4 truss segment to

the port side of the integrated truss system of

the orbital outpost.  The truss, part of the

station’s girder‐like backbone, includes a new

set of photovoltaic solar arrays.  When unfurled

to their full length of 240 feet, the arrays will

provide additional power for the station in

preparation for the delivery of international

science modules over the next two years.  Each

of the 82 active array blankets that are grouped

into 31‐and‐a‐half “bays” contain 16,400

individual silicon photovoltaic cells to convert

sunlight into electricity.

The truss also contains a device called the Solar

Alpha Rotary Joint (SARJ), which will rotate

360 degrees either clockwise or

counterclockwise to position the P4 and P6

solar arrays to track the sun for electrical power

generation.
August 2006 MISSION OVERVIEW 2

Space Shuttle Atlantis rolled over from the Orbiter Processing Facility to the  

Vehicle Assembly Building in late July in preparation for STS‐115

Atlantis is scheduled to liftoff from Launch Pad

39‐B at NASA’s Kennedy Space Center, Florida,

no earlier than late August.  As was the case for

the just completed STS‐121 mission, Atlantis’

launch is governed by daylight lighting

requirements at the Florida spaceport and the

proper lighting in orbit for photography of the

external fuel tank once it is jettisoned eight and

a half minutes after launch.

Each day’s launch opportunity will be about

five minutes long, timed to when the rotation of

the Earth places the Kennedy Space Center in

the plane of the station’s orbit.

STS‐115 is expected to last 11 days with the

three scheduled spacewalks planned for flight

days 4, 5 and 7.  If a focused inspection of

Atlantisʹ heat shield is requested once the

shuttle is at the station, the flight could be

extended by one or two days, changing the

flight days on which the three spacewalks will

be conducted.

The mission is commanded by Navy Capt.  

Brent Jett, 47, a veteran of three previous

spaceflights, including the STS‐97 mission in

2000 to the ISS that delivered the first U.S. solar

arrays on the P6 truss structure.

Jett is joined by Navy Capt. Chris Ferguson, 45,

a first‐time flyer who will serve as the missionʹs

pilot. Joe Tanner, 56, will lead two teams of

STS‐115 Commander Brent Jett
August 2006 MISSION OVERVIEW 3

spacewalkers in his fourth flight into space.  

Tanner was a crewmate of Jettʹs on the STS‐97

mission.  Coast Guard Cmdr. Dan Burbank, 45,

makes his second flight to the ISS on the

STS‐115 mission and will serve as flight

engineer during launch and landing.  Navy

Cmdr. Heide Stefanyshyn‐Piper (Hyeʹ‐dee

Steph‐uh‐nihʹ‐shun Pyeʹ‐pur), 43, makes her

first spaceflight on Atlantisʹ mission, joining

Tanner for two of the three planned

spacewalks.  Canadian Space Agency astronaut

Steve MacLean (Muh‐klaineʹ), 51, rounds out

the crew and will join Burbank for the second

spacewalk of the flight in his second mission

into space.  He flew a scientific research mission

aboard Columbia 14 years ago and will become

the second Canadian astronaut to walk in

space, following in the footsteps of his

colleague, Chris Hadfield.

Tanner has conducted five spacewalks in his

career both in service of the Hubble Space

Telescope and at the station.  Burbank, Piper

and MacLean will be conducting their first

spacewalks.

The primary purpose of the mission is to install

the huge P3/P4 truss to the existing P1 truss on

the port side of the station.  Complex work with

both the shuttle and the station’s robotic arms

will be required to not only install the new truss

segment, but also inspect Atlantis’ wings and

thermal protection system, duplicating

inspections performed on STS‐114 and STS‐121

missions.  This heat shield inspection

conducted over several days will ensure the

shuttle is safe to return to Earth.

MacLean, who helped develop the Canadian‐

built robotic arms for both shuttle and station,

will become the first Canadian to operate both

systems during the STS‐115 mission.

A view of Orbiter Discoveryʹs underside during the STS‐121 Rendezvous Pitch Maneuver
August 2006 MISSION OVERVIEW 4

Expedition 13 crew ‐ Commander Pavel Vinogradov (center), flight engineer and NASA science

officer Jeff Williams (right) and European Space Agency (ESA) astronaut Thomas Reiter (left)

When Atlantis arrives at the station two days

after its launch, the six‐person shuttle crew will

be greeted by the three‐man Expedition 13

crew.  Russian Commander Pavel Vinogradov,

53, and Flight Engineer and NASA Science

Officer Jeff Williams, 48, have been aboard the

complex since April 1 following their launch on

the Soyuz TMA‐8 spacecraft from the Baikonur

Cosmodrome in Kazakhstan.  They are

scheduled to return to Earth in late September

following the arrival of the next station crew,

Expedition 14, which will launch in another

Soyuz spacecraft right after the STS‐115

mission.  The third member of the Expedition

13 crew is European Space Agency astronaut

Thomas Reiter, 48, who arrived at the outpost

on the STS‐121 mission.  He will remain on

board with the Expedition 14 crew through the

end of the year.

The STS‐115 flight marks the second time as

many as four out of the five partner agencies

will have been represented on the station

together.
August 2006 MISSION OVERVIEW 5

After launching, the crew will spend its first

hours in space setting up equipment and

downlinking initial imagery of Atlantis’

external fuel tank following its jettison.

The next day, Ferguson and Burbank will join

MacLean at the aft flight deck of Atlantis to

grapple and unberth the 50‐foot‐long Orbiter

Boom Sensor System (OBSS), a crane extension

for the shuttle’s robotic arm.  The extension

uses two lasers and a high‐resolution television

camera to examine the leading edges of

Atlantis’ wings and its nose cap for any sign of

damage that may have occurred during launch.  

Additional inspections of the wings and other

orbiter surfaces will take place the day prior to

undocking from the station and in the hours

following undocking.  Imagery analysts at the

Johnson Space Center in Houston will pore

over the data and report their findings to

members of the Mission Management Team.

While the boom inspection is under way,

Tanner and Piper will prepare the spacesuits

and tools that they, Burbank and MacLean will

use during their spacewalks.  Jett and Ferguson

will conduct periodic firings of Atlantis’ jets in

the standard plan to place Atlantis on a precise

course to rendezvous with the station the

following day.  Once the initial boom

inspection is complete, the boom will be put

back onto the starboard sill, or edge, of the

shuttle cargo bay for use later in the mission.
August 2006 MISSION OVERVIEW 6

On the third day of the mission, Jett will take

over manual control of Atlantis at a point about

1,000 feet directly below the station.  He then

will fly the shuttle to just 600 feet under the

complex and begin a slow back‐flip rotation of

the orbiter.  This will enable Vinogradov and

Williams to capture high‐resolution digital

imagery of Atlantis for engineers on the ground

to analyze the condition of the shuttle’s heat

shield.

With the backflip completed, Jett will slowly

maneuver Atlantis to a docking with the station

at the forward end of the Destiny Laboratory,

setting the stage for a week of joint operations

between the two crews.

Even before the hatches are opened to allow the

nine crew members to greet one another,

Ferguson and Burbank will use the shuttle’s

robotic arm to grapple the massive P3/P4 truss.  

Once hatches are opened, MacLean will join

Williams in Destiny at the robotic work station

to maneuver the station’s Canadarm2 robotic

arm for a handoff of the truss from Ferguson

and Burbank.  The truss will remain grappled

to Canadarm2 overnight.

Tanner and Piper will then begin spacewalk

preparations as they use a new technique called

an extravehicular activity (EVA) “campout”

protocol which requires them to prebreathe

pure oxygen in masks to begin to cleanse

nitrogen from their bloodstreams.  That will

prevent a condition known as decompression

sickness, commonly called the “bends.”

Tanner and Piper will sleep in Quest overnight

as its pressure is lowered to 10.2 pounds per

square inch to further rid their bloodstreams of

nitrogen and shorten their final spacewalk

preparations.  Burbank and MacLean will

follow the same regime prior to the second

spacewalk of the mission, and Tanner and Piper

will repeat the procedure prior to the third

spacewalk.

As Tanner and Piper prepare to leave the Quest

airlock for their first spacewalk, MacLean and

Williams will slowly and carefully position the

P3/P4 truss at the edge of the P1 truss.  They

will align it properly through the use of

television cameras and install it to P1 so that the

two segments are captured and hard mated to

one another through a series of four structural

bolts.  Once the bolts are secure and the truss

segments are mated, the Canadarm2 will

release its grasp of the truss’ grapple fixture.

After emerging from Quest, Tanner and Piper

will move to the newly installed truss segment

where they will first connect power cables in

one of two trays housing the truss’ electronics.  

The next operation calls for Piper to release

launch restraints for the Solar Array Blanket

Boxes (SABB) that house the folded arrays.  

Tanner and Piper will then release similar

restraints for the Beta Gimbal Assemblies

(BGA) that serve as the structural link between

the truss’ integrated electronics and the massive

Solar Array Wings (SAWs).  The BGAs will

rotate to the correct position for the actual

deployment of the arrays on flight day 6.
August 2006 MISSION OVERVIEW 7

Before the spacewalkers can complete the

unstowing of the array blanket boxes, they

must move a keel pin out of the way.  The pin

held the P4 truss segment in place at the bottom

of the payload bay for launch.

Next, Tanner and Piper will begin preparations

for operation of the Solar Alpha Rotary Joint

and completing the connection of electrical

cables between the new P3 truss and the P1

truss.

As the spacewalk concludes, Mission Control in

Houston will command the activation of the P4

truss for a thorough checkout of its systems and

the arrays as they remain folded in their

canisters.

The second spacewalk by Burbank and

MacLean on flight day 5 will be devoted to final

preparations for the activation of the Solar

Alpha Rotary Joint.  Six launch restraints, four

thermal covers and sixteen launch locks will be

removed and released to set the stage for its

activation.

Before the crew is awakened to begin the sixth

day of the mission, flight controllers will initiate

the deployment of the new arrays on the P4

truss.  Only one half of the 31.5 ʺbaysʺ 

encapsulating the arrays will be immediately

unfurled so that flight controllers can confirm a

new deploy procedure developed after the P6

array deployment on the STS‐97 mission in

2000.

At that time, the solar array panels experienced

a phenomenon called ʺstictionʺ that caused a

tension bar designed to keep the arrays taut to

oscillate back and forth.  The effect was similar

to a sail flapping while held on the top and

bottom.  It was determined after the flight that a

silicon coating designed to protect the copper

wiring in the solar cells against the degrading

effects of atomic oxygen particles had fused

over time as the arrays were in their blanket

boxes.  In addition, the ʺstictionʺ could be

lessened by deploying the arrays after they

have been warmed in the sun first and by using

a high‐ tension deployment mode. A second

solar array wing was deployed on STS‐97 using

the high‐tension mode without incident.  

Cameras and visual clues have also been added

to make it easier for the crew to better monitor

deployment.

For this flight, the arrays will be deployed at a

high tension to avoid large motion by the

tension bar itself.  The temperature of the arrays

also will be controlled to avoid a repeat of the

“stiction” issue.

Once the crew is awake, they will take over the

array deployment process, commanding the
August 2006 MISSION OVERVIEW 8

wings to unfold to an interim position of 49 per

cent of full deployment, then to their full length

of 240 feet.  The incremental deployment

sequence is performed when the sun is at the

correct angle relative to the station for the

proper thermal conditioning of the array

blankets.

After the arrays are fully deployed, the

Canadarm2 will “walk‐off” from the Mobile

Transporter back to the Destiny Lab to enable

the rail car to move to a work site at the P3 truss

for a checkout of its functionality the next day.  

Work site No. 8 on the truss is where the

station’s robotic arm will be positioned during

the next flight, STS‐116, for the installation of

the P5 truss component.

The following day, flight day 7, will find

Tanner and Piper conducting the third and final

spacewalk of the mission to release restraints

for the deployment of the heat‐dissipating

radiator on the P4 truss.  They also will deploy

braces to hold the Solar Alpha Rotary Joint in

place, remove a large keel pin that restrained

the P3 truss during launch in Atlantis’ cargo

bay, and install bolt clips to properly position

the Beta Gimbal Assemblies for the newly

deployed P4 arrays.

In addition, Tanner and Piper expect to remove

a materials science experiment from the hull of

the Quest airlock to return to Earth, remove and

replace components of the S1 (starboard) truss’

S‐band communications system and install a

new external wireless communications antenna

on forward end of Destiny to enable improved

television transmissions from cameras mounted

on the helmets of future spacewalkers.  They

will also install a thermal shroud on the

station’s KU‐band antenna to protect it from the

extreme heat and cold of low‐Earth orbit.

Once the health check of the rail’s work site No.

8 is completed, the rail car will return to its

normal parking place on work site No. 4 at the

S0 truss in the middle of the truss system.  The

station’s arm will move from its grapple fixture

on Destiny to the Mobile Transporter at work

site No. 4 the next day to support additional

inspections of Atlantis’ thermal insulation.

The crew enjoy a half day of off‐duty time on

flight day 8 and conduct the final transfer of

items to the station.

On flight day 9, the crew will complete its

transfer work, say farewell to the Expedition 13

crew and close hatches between the two

vehicles.  Undocking from the station is now

expected late in the shuttle crewʹs work day

with a possible flyaround of the complex

dependent on crew time and propellant

capability.
August 2006 MISSION OVERVIEW 9

Expedition 13 crew ‐ Commander Pavel

Vinogradov (left), flight engineer and NASA

science officer Jeff Williams (right), and

European Space Agency (ESA) astronaut

Thomas Reiter (center)

Atlantis will separate from the station in a

trajectory that will enable the crew to document

the outpost with its newly installed truss that

will make it asymmetrical as it orbits the Earth.  

The starboard truss will be built out in a similar

fashion when the S3 / S4 truss segments are

delivered to the ISS on the STS‐117 mission next

year.

Jett and Ferguson will monitor Atlantis’

departure from the station as the shuttle moves

to a distance of 40 nautical miles.  At that point,

an engine firing will place Atlantis in a station

keeping orbit in the unlikely event it would

have to return to the complex due to damage to

its heat shield.

On flight day 10, the shuttle crew will use the

shipʹs robotic arm to grapple the boom sensor

system once again to conduct a complete ʺlateʺ 

inspection of Atlantisʹ wings and nose cap in a

procedure almost identical to the activity on

flight day 2.

Ferguson, Burbank and MacLean will take

turns monitoring the maneuvers of the boom

sensor crane to view the reinforced carbon‐

carbon paneling on the leading edges of the

shuttleʹs wings and nose cap.  That will provide

the imagery needed to permit mission

managers to give Atlantis and its crew the final

clearance to return home.

On shuttle mission STS‐121, part of this

inspection was performed while the shuttle was

docked to the station.  Using a lesson learned

from STS‐121, the inspection plan has been

changed to provide a more efficient way for

Atlantisʹ astronauts to conduct the final

examination of the shuttleʹs heat shield.  During

Discoveryʹs flight on STS‐121, lighting and

clearance issues caused time to be added to the

inspection of the port wing.  The revised

timeline for STS‐115 will eliminate such issues.

On flight day 11, the astronauts will stow their

gear, test Atlantis’ flight control surfaces and

steering jets and review their entry and landing

procedures.

Atlantis is scheduled to land on the morning of

flight day 12, setting the stage for subsequent

shuttle missions to continue the expansion of

the International Space Station.
August 2006 TIMELINE OVERVIEW 10

STS-115 TIMELINE OVERVIEW 

FLIGHT DAY 1: 

• Launch

• Payload Bay Door Opening

• Ku‐Band Antenna Deployment

• Shuttle Robot Arm Power Up

• External Tank Handheld Video, Umbilical

Well Imagery and Wing Leading Edge

Sensor Data Downlink

FLIGHT DAY 2: 

• Shuttle Robot Arm Checkout

• Shuttle Robot Arm Grapple of Orbiter

Boom Sensor System (OBSS)

• Inspection of Shuttle Thermal Protection

System and Wing Leading Edge reinforced

carbon‐carbon (RCC)

• Spacesuit Checkout

• Orbiter Docking System Outer Ring

Extension

• Airlock Preparations

• Rendezvous Tool Checkout

FLIGHT DAY 3: 

• Rendezvous Operations

• Terminal Initiation Engine Firing

• Rendezvous Pitch Maneuver and ISS Digital

Photography of Atlantis

• Docking to the International Space Station

• Hatch Opening and Welcoming by

Expedition 13 Crew

• Shuttle robot arm grapple of P3/P4 Truss

and handoff to Station robot arm

• Tanner/Piper sleep in Quest Airlock for

spacewalk pre‐breathe campout protocol

FLIGHT DAY 4: 

• Station robot arm installs P3/P4 Truss

installation on P1 Truss attachment

• Tanner/Piper EVA No. 1 to connect power

cables, release Solar Array Blanket Box and

Beta Gimbal Assembly Restraints and to

prepare the Solar Alpha Rotary Joint for

operations

• Activation of the P4 array assembly

• Burbank/MacLean sleep in Quest Airlock

for spacewalk pre‐breathe campout protocol

FLIGHT DAY 5: 

• Burbank/MacLean EVA No. 2 to complete

preparations for the activation of the Solar

Alpha Rotary Joint for operations

• Solar Alpha Rotary Joint activation and

checkout during crew sleep

• Station robot arm is maneuvered to P4 array

deploy viewing position
August 2006 TIMELINE OVERVIEW 11

FLIGHT DAY 6: 

• First bay of P4 Solar Array 4A and 2A masts

are deployed before crew wake up

• Remaining bays of the P4 Solar Array 4A

and 2A masts are deployed (31.5 bays per

array)

• Station robot arm conducts “double walk

off” from the Mobile Base System to the

Destiny Lab to prepare for EVA No. 3 and

Truss work site No. 8 checkout in

preparation for use during the STS‐116

mission

• Tanner/Piper sleep in Quest Airlock for

spacewalk pre‐breathe campout protocol

FLIGHT DAY 7: 

• Tanner/Piper EVA No. 3 to release the

photovoltaic radiator restraints for its

deployment, deploy the Solar Alpha Rotary

Joint braces, remove the P3 keel pin to clear

the truss path for the Mobile Transporter,

retrieve the MISSE 5 experiment, install an

external wireless TV transmission antenna

and conduct other get‐ahead work for

future spacewalks

• The Mobile Transporter is moved from

work site 7 on the truss to work site 8 to

checkout that location in advance of

STS‐116, then moves back to its normal

“parking space” on work site 4

FLIGHT DAY 8: 

• Crew off‐duty period for a half a day

• Joint Crew News Conference

• Transfer activities and water supply

transfers from shuttle to station

FLIGHT DAY 9: 

• Final transfers

• Rendezvous tool checkout in preparation

for undocking

• Final farewells and hatch closure

• Undocking

FLIGHT DAY 10: 

• Shuttle robotic arm grapple of OBSS

• OBSS late inspection of Atlantis’ starboard

wing and nose cap

• OBSS final berthing by shuttle robot arm

• Final shuttle robot arm powerdown

FLIGHT DAY 11: 

• Flight Control System Checkout

• Reaction Control System Hot‐Fire Test

• Cabin Stowage

• Deorbit Timeline Review

• Ku‐Band Antenna Stowage

FLIGHT DAY 12: 

• Deorbit Preparations

• Payload Bay Door Closing

• Deorbit Burn

• KSC Landing
August 2006 MISSION PRIORITIES 12

MISSION PRIORITIES 

Priorities for the STS‐115 space shuttle mission

include (in order):

• Perform flight day 2 inspections

• Perform rendezvous pitch maneuver during

rendezvous and docking for orbiter thermal

protection system inspection using

International Space Station imagery

• Install integrated truss structure (ITS) P3/P4

to ITS P1

• Activate ITS P3/P4 systems to receive

survival power from P6

• Transfer mandatory quantities of water

• Transfer critical items per transfer priority

list

• Configure P4 for power generation and

distribution and deploy P4 solar array

wings

• Deploy Solar Alpha Rotary Joint (SARJ)

drive lock assemblies

• Install four of four Alpha Joint Interface

Structure (AJIS) struts

• Remove launch locks and restraints,

activate and check out SARJ

• Install all SARJ brace beams

• Deploy P4 photovoltaic radiator

• Complete spacewalk tasks necessary to

enable the mobile transporter (MT) to be

moved to work site No. 8

• Remove circuit interrupt devices 6 and 8

• Remove and replace S‐band string 1 S‐band

Antenna Support Assembly (SASA)

• Conduct orbiter late inspection for

micrometeoroid orbital debris

• Relocate MT to worksite No. 8 and check

out Mobile Servicing System

• Transfer remaining items per transfer

priority list

• Remove and replace starboard S‐band base

band processor and transponder

• Install extravehicular activity (EVA)

temporary rail stop, stow P3 MT stop and

EVA shuttle tether stop

• Install P6 shoulder bolt retainer

• Retrieve Materials on the International

Space Station Experiment (MISSE) 5

• Transfer minimum of 25 pounds of oxygen

to stationʹs high‐pressure gas tanks

• Perform daily middeck activities to support

payload operations:

− MICROBE

− Bioavailability and Performance Effects

of Promethazine During Spaceflight

− Sleep Wake Actigraphy and Light

Exposure During Spaceflight

− Latent Virus:  Monitoring Reactivation

and Shedding in Astronauts
August 2006 MISSION PRIORITIES 13

• Perform U.S./Russian daily space station

payload status checks as required

• Perform internal thermal control system

sampling

• Perform defibrillator checkout

• Install antenna group interface tube heat

shield

• Set the soft capture latches for the P5 install

• Conduct modified Rocketdyne truss

attachment system visual inspection and

foreign object debris check

• Engage P6 beta gimbal assembly hinge lock

• Reboost space station with orbiter, using

available propellant, to maintain altitude

and rendezvous requirements

• Perform imagery survey of the space

stationʹs exterior during orbiter flyaround

(if propellant available)

• Perform ram burn observation and Maui

analysis of upper‐atmospheric injections

payloads of opportunity

• Perform Station Detailed Test Objective

(SDTO) 15003‐U:  Microgravity

Environment Definition for the SARJ

checkout

• Perform SDTO 12004‐U:  Shuttle Booster

Fan Bypass

• Perform SDTO 13005‐U:  ISS Structural Life

Validation and Extension

• Perform program‐approved EVA get‐ahead

tasks:

− Release P4 integrated equipment

assembly micrometeoroid orbital debris

shield bolt torques

− Position auxiliary portable foot

restraints for STS‐116’s EVA 1

− Install external wireless instrumentation

system antennas and cable on U.S. lab

− Install S1 crew and equipment

translation aid light

− Install non‐propulsive vent on lab

− Mate P1/P3 fluid umbilical
August 2006 LAUNCH & LANDING 14

LAUNCH AND LANDING 

LAUNCH 

As with all previous space shuttle launches,

Atlantis on STS‐115 will have several modes

available that could be used to abort the ascent

if needed due to engine failures or other

systems problems.  Shuttle launch abort

philosophy aims toward safe recovery of the

flight crew and intact recovery of the orbiter

and its payload.  Abort modes include:

ABORT-TO-ORBIT (ATO) 

Partial loss of main engine thrust late enough to

permit reaching a minimal 105 by 85 nautical

mile orbit with orbital maneuvering system

engines.

TRANSATLANTIC ABORT LANDING 

(TAL) 

Loss of one or more main engines midway

through powered flight would force a landing

at either Zaragoza, Spain; Moron, Spain; or

Istres, France.  For launch to proceed, weather

conditions must be acceptable at one of these

TAL sites.

RETURN-TO-LAUNCH-SITE (RTLS) 

Early shutdown of one or more engines, and

without enough energy to reach Zaragoza,

would result in a pitch around and thrust back

toward KSC until within gliding distance of the

Shuttle Landing Facility.  For launch to

proceed, weather conditions must be forecast to

be acceptable for a possible RTLS landing at

KSC about 20 minutes after liftoff.

ABORT ONCE AROUND (AOA) 

An AOA is selected if the vehicle cannot

achieve a viable orbit or will not have enough

propellant to perform a deorbit burn, but has

enough energy to circle the Earth once and land

about 90 minutes after liftoff.

LANDING 

The primary landing site for Atlantis on

STS‐115 is the Kennedy Space Centerʹs Shuttle

Landing Facility.  Alternate landing sites that

could be used if needed due to weather

conditions or systems failures are at Edwards

Air Force Base, California, and White Sands

Space Harbor, New Mexico.
MISSION PROFILE 

August 2006 MISSION PROFILE 15

CREW 

Commander: Brent Jett

Pilot: Chris Ferguson

Mission Specialist 1: Joe Tanner

Mission Specialist 2: Dan Burbank

Mission Specialist 3: Heidemarie

Stefanyshyn‐Piper

Mission Specialist 4: Steve MacLean

LAUNCH 

Orbiter: Atlantis (OV‐104)

Launch Site: Kennedy Space Center

Launch Pad 39B

Launch Date: No Earlier Than August

27, 2006

Launch Time: 4:30 p.m. EDT (Preferred

In‐Plane launch time for

8/27)

Launch Window: 5 Minutes

Altitude: 122 Nautical Miles (140

Miles) Orbital Insertion;

190 NM (218 Miles)

Rendezvous

Inclination: 51.6 Degrees

Duration: 10 Days 20 Hours 16

Minutes

VEHICLE DATA 

Shuttle Liftoff Weight: 4,525,808

pounds

Orbiter/Payload Liftoff Weight: 269,840

pounds

Orbiter/Payload Landing Weight: 199,679

pounds

Software Version: OI‐30

Space Shuttle Main Engines: 

SSME 1: 2044

SSME 2: 2048

SSME 3: 2047

External Tank: ET‐118

SRB Set: BI‐127

RSRM Set: 94

SHUTTLE ABORTS 

Abort Landing Sites 

RTLS: Kennedy Space Center Shuttle

Landing Facility

TAL: Primary – Zaragoza, Spain.  

Alternates Moron and Istres, France

AOA: Primary – Kennedy Space Center

Shuttle Landing Facility; Alternate

White Sands Space Harbor

Landing 

Landing Date: No Earlier Than Sept. 7,

2006

Landing Time: 12:02 p.m. EDT

Primary landing Site: Kennedy Space Center

Shuttle Landing Facility

PAYLOADS 

P3/P4 Truss
August 2006 CREW 16

STS-115 ATLANTIS CREW 

The STS‐115 mission will resume the assembly

of the International Space Station with the

delivery and installation of stationʹs second port

truss segment (P3/P4) onto the P1 truss and the

second set of solar arrays and batteries.  

Following the installation of the segments

utilizing both the shuttle and the station robotic

arms, a series of three spacewalks will complete

the final connections and prepare for the

deployment of the station’s second set of solar

arrays.

To reflect the primary mission of the flight, the

patch depicts a solar panel as the main element.  

As Atlantis launches toward the space station,

its trail depicts the symbol of the Astronaut

Office.  The starburst, representing the power of

the sun, rises over the Earth and shines on the

solar panel.  The shuttle flight number 115 is

shown at the bottom of the patch, along with

the space station assembly designation 12A (the

12th American assembly mission).  The blue

Earth in the background reminds us of the

importance of space exploration and research to

all of Earth’s inhabitants.

Short biographical sketches of the crew follow

with detailed background available at:  

http://www.jsc.nasa.gov/Bios/.
August 2006 CREW 17 A veteran of three space missions, Brent Jett

will command the crew of STS‐115, the shuttleʹs

19th mission to the International Space Station.  

He served as the pilot of STS‐72 in January 1996

and then again on STS‐81 in 1997.  He flew as

the commander on STS‐97 in 2000.  A captain in

the U.S. Navy, he has traveled more than 12

million miles over the three missions and

logged 699 hours, 15 minutes and 57 seconds in

space.  He has overall responsibility for the on‐

orbit execution of the mission, orbiter systems

operations and flight operations including

landing the orbiter.  In addition, Jett will fly the

shuttle ina procedure called the rendezvous

pitch maneuver while Atlantis is 600 feet below

the station before docking to enable the station

crew to photograph the orbiterʹs heat shield.  

He will then dock Atlantis to the station.  Jett

will also be heavily involved in spacewalk

preparations and transferring cargo to the

station during the docked phase of the mission. Brent Jett
August 2006 CREW 18

Chris Ferguson

A captain in the U.S. Navy, Chris Ferguson will

make his first journey into space as the pilot of

the STS‐115 shuttle mission.  He will be

responsible for orbiter systems operations and

assisting Jett in the rendezvous and docking to

the International Space Station.  Along with

Dan Burbank, he will use the shuttle’s robotic

arm and the Orbiter Boom Sensor System to

inspect Atlantisʹ heat shield.  He will also assist

Burbank with shuttle arm operations during the

unberthing of the P3/P4 truss from the shuttle

cargo bay and handing it off to the station

robotic arm as well as transfer cargo to and

from the shuttle.  Ferguson will undock

Joe Tanner

Atlantis from the station at the end of the

mission.

With five spacewalks to his credit, STS‐115

Mission Specialist 1 (MS1) Joe Tanner will make

his fourth venture into space.  Tanner will

perform two of the three extravehicular

activities (EVAs), or spacewalks, as EV1.  He

will conduct the first and the third planned

spacewalks with Heide Stefanyshyn‐Piper on

flight days 4 and 7.  He will share

intravehicular duties with Stefanyshyn‐Piper

during the second spacewalk on flight day 5.  

He will be seated on the middeck for launch

and landing.
August 2006 CREW 19

Dan Burbank

Mission Specialist 2 (MS2) Dan Burbank, a

commander in the U.S. Coast Guard, will be

making his second spaceflight.  He previously

flew on Atlantisʹ STS‐106 mission in 2000.  

Burbank, as EV3, will conduct the second

spacewalk of the mission with his colleague

Steven MacLean.  He will serve as the

intravehicular activity crew member helping to

suit up and choreograph spacewalkers Tanner

and Heidemarie Stefanyshyn‐Piper for the first

and third spacewalks.  He will be the prime

operator of the shuttle robotic arm, working

with Ferguson to unberth the P3/P4 truss

segment from the shuttleʹs cargo bay and hand

it off to the station robotic arm, as well as

conduct robotic inspections of the shuttleʹs

thermal protection system.  During the

rendezvous, docking and undocking, he will

manage the rendezvous navigation tools used

to guide the shuttleʹs trajectory relative to the

station.  He will be seated on the flight deck for

launch and landing, serving in the role of flight

engineer.

Heidemarie Stefanyshyn‐Piper

Mission Specialist 3 (MS3) Heide Stefanyshyn‐

Piper will be making her first flight into space.  

A captain in the U.S. Navy, she will serve as

EV2 and conduct the first and third spacewalks

of the mission with Tanner.  She will be the

prime operator of the space station robotic arm

during the second spacewalk, during the

deployment of the solar arrays and during the

transfer of the station arm from the Mobile Base

System to the Destiny laboratory.  She will also

assist Steve MacLean during the unberthing,

handoff and berthing of the Orbiter Boom

Sensor System.  She will serve as the overall

lead for transferring supplies from the shuttleʹs

cargo module to the station.  During the

rendezvous, docking and undocking, she will

manage the handheld laser and the Orbiter

Docking System.  She will oversee payload bay

door closing operations.  She will be seated on

the middeck for launch and the flight deck for

landing.
August 2006 CREW 20

Mission Specialist 4 (MS4) Steve MacLean of the

Canadian Space Agency will make his second

spaceflight on STS‐115.  He first flew on STS‐52

in 1992.  MacLean will execute the second

spacewalk of the mission with Burbank.  He

will be the prime operator of the space station

robotic arm for the mission, overseeing its

operation during the P3/P4 truss segment

installation, during the Orbiter Boom Sensor

System handoffs to the shuttle arm and during

EVA 1 and EVA3.  In addition, MacLean will

assist Burbank and Ferguson during the robotic

inspections of the shuttle’s thermal protection

system.  MacLean will share intravehicular

activity duties with Burbank helping to

choreograph EVA1 and EVA 3 He will also

oversee shuttle payload bay door opening once

Atlantis reaches orbit.  MacLean will be seated

on the flight deck for launch and the middeck

for landing. Steve MacLean
August 2006 MISSION PERSONNEL 21

MISSION PERSONNEL 

CONSOLE POSITIONS FOR STS-115 

Flt. Director CAPCOM PAO

Ascent Steve Stich Tony Antonelli

Ken Ham (Wx)

Kyle Herring

Orbit 1 (Lead) Paul Dye Megan McArthur Kyle Herring (Lead)

Orbit 2 Cathy Koerner Terry Virts Kylie Clem

Planning Bryan Lunney Hans Schlegel Nicole Cloutier

Entry Steve Stich Tony Antonelli

Ken Ham (Wx)

Kelly Humphries

Shuttle Team 4 Mike Sarafin N/A N/A

ISS Orbit 1 Kelly Beck Kevin Ford N/A

ISS Orbit 2 (Lead) John McCullough Pam Melroy N/A

ISS Orbit 3 Kwatsi Alibaruho Zach Jones N/A

Station Team 4 Annette Hasbrook N/A N/A

Mission Control,

Korolev, Russia

Mark Ferring N/A N/A

    

JSC PAO Representative at KSC for Launch – Kelly Humphries

KSC Launch Commentator – George Diller

KSC Launch Director – Mike Leinbach

NASA Launch Test Director – Jeff Spaulding
August 2006 RENDEZVOUS & DOCKING 22

RENDEZVOUS AND DOCKING 

Atlantis’ approach to the International Space

Station during the STS‐115 rendezvous and

docking process will include a tricky maneuver

first demonstrated on STS‐114 and repeated

during the STS‐121 mission.  The orbiter will be

commanded to conduct a back flip pirouette,

enabling station crew members to take digital

images of the shuttle’s heat shield.

With Atlantis’ Commander Brent Jett at the

controls, the shuttle will perform the

360‐degree pitch‐around maneuver with the

orbiter about 600 feet below the station.  The

flip will take about nine minutes to complete,

offering Expedition 13 Commander Pavel

Vinogradov and Flight Engineer Jeff Williams

time to capture required tile surface imagery of

Atlantis.

The photos will then be downlinked through

the station’s Ku‐band communications system

for analysis by systems engineers and mission

managers.

The photography will be performed out of

windows 6 and 8 in the Zvezda Service Module

with Kodak DCS 760 digital cameras and

400mm and 800mm lenses.  The imagery

during the so‐called Rendezvous Pitch

Maneuver (RPM) is among several inspection

procedures instituted after the Columbia

accident.  The procedures are designed to detect

and determine the extent of any damage the

orbiter’s protective tiles and reinforced carbon‐

carbon surfaces might have sustained.

The sequence of events that brings Atlantis to

its docking with the station begins with the

precisely timed launch of the shuttle, placing

the orbiter on the correct trajectory and course

bottomside_800mm.cnv

NOTE

1. indicates

critical focus

point.

2. Sequence is

~16 shots;

repeat

sequence, as

time allows.

bottomside_800mm.cnv

NOTE

1. indicates

critical focus

point.

2. Sequence is

~16 shots;

repeat

sequence, as

time allows.

for its two‐day chase to arrive at the station.  

During the first two days of the mission,

periodic engine firings will gradually bring

Atlantis to a point about nine‐and‐a‐half statute

miles behind the station, the starting point for a

final approach.

About two‐and‐a‐half hours before the

scheduled docking time on flight day 3,

Atlantis will reach that point, about 50,000 feet

behind the ISS.  There, Atlantis’ jets will be

fired in what is called the Terminal Initiation

(TI) burn to begin the final phase of the

rendezvous.  Atlantis will close the final miles

to the station during the next orbit.
August 2006 RENDEZVOUS & DOCKING 23

As Atlantis moves closer to the station, the

shuttle’s rendezvous radar system and

trajectory control sensor (TCS) will begin

tracking the complex and providing range and

closing rate information to the crew.  During

the final approach, Atlantis will execute several

small mid‐course corrections at regular

intervals with its steering jets.  That will place

Atlantis at a point about 1000 feet directly

below the station.  Jett then will take over the

manual flying of the shuttle up the radial

vector, or R‐bar.  The R‐bar is the the imaginary

line drawn between the station and the Earth.

He will slow Atlantis’ approach and fly to a

point about 600 feet directly below the station,

and if required, wait for the proper lighting

conditions.  The rendezvous is designed to

optimize lighting for inspection imagery as well

as crew visibility for critical rendezvous events.

On verbal cue from Pilot Chris Ferguson to

alert the station crew, Jett will command

Atlantis to begin a nose‐forward, three‐quarter

of a degree per second rotational back flip.  At

RPM start, the ISS crew will begin of series of

photographs for inspection.  The sequence of

photography mapping provides optimization

of the lighting conditions.

Rendezvous Approach Profile

Space Shuttle Rendezvous Maneuvers

OMS-1 (Orbit insertion) - Rarely used ascent 

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.
August 2006 RENDEZVOUS & DOCKING 24

Both the 400 and 800 mm digital camera lenses

will be used to capture imagery of the required

surfaces of the orbiter.  The 400 mm lens

provides up to three inch resolution and the 800

mm lens can provide up to one inch resolution

as well as detect gap filler protrusions of

greater than .25 inch.  The imagery includes the

upper surfaces of the shuttle as well as Atlantis’

belly, capturing pictures of the nose landing

gear door seals, the main landing gear door

seals and the elevon cove with one inch

analytical resolution.  Since the STS‐114

mission, additional zones were added for the

800 mm lens to focus on the gap fillers on

Atlantis’ belly when the orbiter is at 145 and

230 degree angles during the flip.  There should

be enough time for two sets of pictures.

When Atlantis completes its rotation, it will

return to an orientation with its payload bay

facing the station.

Jett then will move Atlantis to a position about

400 feet in front of the station along the V‐Bar,

or the velocity vector, the direction of travel for

both spacecraft.  Ferguson will provide Jett

with navigation information as he slowly

inches the shuttle toward the docking port at

the forward end of the station’s Destiny

Laboratory.

Ferguson will join mission specialists Joe

Tanner, Dan Burbank and Heidemarie‐

Stefanyshyn‐Piper in playing key roles in the

rendezvous.  They will operate laptop

computers processing the navigational data, the

laser range systems and Atlantis’ docking

mechanism.

Using a view from a camera mounted in the

center of Atlantis’ docking mechanism as a key

alignment aid, Jett will precisely match up the

docking ports of the two spacecraft.  He will fly

to a point where the docking mechanisms are

30 feet apart and pause to check the alignment.

For Atlantis’ docking, Jett will maintain the

shuttleʹs speed relative to the station at about

one‐tenth of a foot per second (while both

Atlantis and the station are traveling at about

17,500 mph), and keep the docking mechanisms

aligned to within a tolerance of three inches.  

When Atlantis makes contact with the station,

preliminary latches will automatically attach

the two spacecraft.  Immediately after Atlantis

docks, the shuttleʹs steering jets will be

deactivated to reduce the forces acting at the

docking interface.  Shock absorber‐like springs

in the docking mechanism will dampen any

relative motion between the shuttle and the

station.
August 2006 RENDEZVOUS & DOCKING 25

Once that motion between the spacecraft has

been stopped, Tanner will secure the docking

mechanism, sending commands for Atlantis’

docking ring to retract and to close a final set of

latches between the two vehicles.

UNDOCKING, SEPARATION AND 

DEPARTURE 

With additional inspections of Atlantis’ heat

shield scheduled on flight day 9, the day before

undocking, and flight day 10 immediately post

undocking, the orbiter will undock with the

shuttle robotic arm and Orbiter Boom Sensor

System (OBBS) deployed in their mated

configuration.  The OBSS will be stowed after

the inspections are completed.

Once Atlantis is ready to undock, Tanner will

send a command to release the docking

mechanism.  At initial separation of the

spacecraft, springs in the docking mechanism

will push the shuttle away from the station.  

Atlantis’ steering jets will be shut off to avoid

any inadvertent firings during the initial

separation.

Once Atlantis is about two feet from the station,

with the docking devices clear of one another,

Ferguson will turn the steering jets back on and

fire them to very slowly move away.  From the

aft flight deck, Ferguson will manually control

Atlantis within a tight corridor as the orbiter

separates from the station, essentially the

reverse of the task performed by Jett just before

Atlantis docked.

Atlantis will continue moving away to a

distance of about 450 feet, where Ferguson will

initiate a maneuver to fly the shuttle directly

above the station.  Once Atlantis reaches that

point, Ferguson will fire Atlantis’ jets to depart

the vicinity of the station for the final time.  

Atlantis will separate to a distance of about 40

nautical miles from the station.  The shuttle will

remain there to protect for a return to the

complex in the unlikely event the late

inspection reveals any damage to the heat

shield.
August 2006 SPACEWALKS 26

SPACEWALKS 

The primary focus for STS‐115’s spacewalks are

to install the Port 3/Port 4 (P3/P4) truss segment

to the port side of the integrated truss system

and prepare it for on‐orbit activities.  Three

spacewalks are planned and will be performed

on flight days four, five and seven.  EVAs 1 and

2 will be back‐to‐back spacewalks.

If a focused inspection of the space shuttle is

required, EVA 1 and 2 will be performed on

flight days four and six and EVA 3 will be on

flight day 8.

Mission specialists Joe Tanner and Heide

Stefanyshyn‐Piper will perform EVAs 1 and 3.  

Mission specialists Dan Burbank and Steve

MacLean, a Canadian Space Agency astronaut,

will perform EVA 2.  Tanner has conducted five

previous spacewalks.  These will be the first

spacewalks for Stefanyshyn‐Piper, Burbank and

MacLean.

Mission Specialist Joe Tanner

Mission Specialist Heide Stefanyshyn‐Piper

On EVAs 1 and 3, Tanner will wear the

spacesuit marked with solid red stripes, while

Stefanyshyn‐Piper will wear an all‐white

spacesuit.  On EVA 2 Burbank will wear the

spacesuit marked with broken horizontal red

stripes, while MacLean will wear a spacesuit

marked with broken diagonal red stripes.  Each

spacewalk is estimated to last six and a half

hours.
August 2006 SPACEWALKS 27

Mission Specialist Dan Burbank

All the spacewalks will be conducted from the

station’s joint airlock, the U.S. Quest airlock.  

Before each EVA, the astronauts will prepare

using a new prebreathe protocol first tested

during the handover of Expedition 12 in April.  

The astronauts will spend the night as the

procedure is called, in the airlock.  “Camping

out,” helps reduce the amount of time typically

required for the prebreathe exercise and, in

some cases, the complexity of the next

morning’s EVA preparations.  Consequently, it

aids the crew to get outside earlier to perform

the day’s EVA tasks.

The protocol calls for the crew members to

isolate themselves in the airlock and lower the

air pressure to 10.2 pounds per square inch, or

psi—the station is kept at 14.7 psi, or near sea‐

level pressure—and then camp out for the

night.  Astronauts aboard the shuttle perform a

similar procedure for shuttle‐based EVAs,

lowering the spacecraft’s air pressure a day or

so before a spacewalk.

The morning of EVA 1, the Integrated Truss

Segment (ITS) P3/P4 will be attached to the Port

1 (P1) segment.  The P3 segment consists of the

Mission Specialist Steve MacLean

P3 truss and the Solar Alpha Rotary Joint

(SARJ), a device which will rotate 360 degrees

clockwise and counterclockwise to position the

P4 and P6 solar arrays to track the sun for

electrical power.  The P4 segment provides the

station with a second set of photovoltaic Solar

Array Wings (SAWs) that will provide

additional power for the station once unfurled

to their full length of 240 feet.

MacLean will become the first Canadian to

operate the Space Station Remote Manipulator

System (SSRMS), or Canadarm2, when the crew

mates the P3/P4 truss to the station.  He will be

the second Canadian to walk in space.

The station eventually will have 11 integrated

truss segments that stretch 356 feet from end to

end.  They will support four virtually identical

solar array assemblies that provide electrical

power.  They also will support radiators that

will cool the station.

ISS assembly sequence, the trusses will provide

a path for the Special Purpose Dexterous

Manipulator (SPDM) that will slide along rails

on the station’s main truss structure.  Major P3

subsystems include the Segment‐to‐Segment
August 2006 SPACEWALKS 28

Attach System (SSAS), Solar Alpha Rotary Joint

(SARJ), and Unpressurized Cargo Carrier

Attach System (UCCAS).  Major P4 subsystems

include the Photovoltaic Module (PVM),

Photovoltaic Radiator (PVR), and Modified

Rocketdyne Truss Attachment System

(MRTAS).

The STS‐115 spacewalks also include removal

of Circuit Interrupt Devices (CIDs) 6 and 8 to

support power reconfiguration tasks on

Mission 12A.1, and an upgrade of the

Integrated Truss Segment S1 truss S‐band

system Baseband Signal Processor (BSP) and

Transponder.

EVA 1 

The P3/P4 activation operations are complex

and challenging for both the crew and the

ground teams.  The activation involves detailed

sequential task choreography, time and

cooperation between the intra‐vehicular,

extra‐vehicular crew members and multiple

Mission Control systems disciplines.

During the first EVA, the crew will prepare the

truss for activation and prepare the solar arrays

for deployment.

As Tanner and Piper prepare for the spacewalk,

Steve MacLean and Jeff Williams will use the

station’s robotic arm to slowly move the

17‐and‐a‐half ton P3/P4 truss to the port side of

the integrated truss system.  They will align it

using a television camera and then mate it to

P1.

Once Tanner and Piper have left the airlock,

they will move to the newly installed truss.  

Upon a “go” from Mission Control that the

proper electrical inhibits are in place, Tanner

will connect the power cables in the P1‐to‐P3

lower utility tray, where the electrical

connections are housed.  He will also close the

number 7 Circuit Interrupt Devices.  Upon

completion, Tanner will give the ground

control team a “go” to begin the activation of

the P3/P4 truss.

Piper will also be working on the P3/P4 truss.  

Her first task will be to release the aft and

forward Solar Array Blanket Box (SABB) launch

restraints, unbolting the SABBs from the

Integrated Equipment Assembly.  The SABBs

hold the folded solar arrays.  Piper will release

the aft SABB first.
August 2006 SPACEWALKS 29

Piper and Tanner will then release similar

restraints for the Beta Gimbal Assemblies

(BGA).  The BGAs are the structural link

between the truss’ integrated electronics and

the Solar Array Wings.  Piper will release the

forward wing BGA; Tanner will release the aft

wing BGA.

Prior to the forward wing BGA release, Tanner

will rotate a keel pin that held the P4 truss

segment in place for launch.  Tanner and Piper

will then complete unstowing the SABBs.  Next,

the spacewalk crew will begin preparing the

Solar Alpha Rotary Joint (SARJ) for rotation by

installing the drive lock assemblies (DLA) to

their on‐orbit position.  Piper will also deploy

and rigidize the four Alpha Joint Interface

Structure (AJIS) struts.  The AJIS struts must be

rigidized, for purposes of structural loading,

prior to removing any of the launch locks on

the subsequent EVA.

When Mission Control has completed the P3/P4

activation to a point where the next set of

electrical inhibits are in place, Tanner will

connect the electrical cables in the P1‐to‐P3

upper utility tray.  He will also remove the

number 6 and 8 Circuit Interrupt Devices in

preparation for work on the next shuttle flight,

STS‐116.

With the conclusion of the spacewalk, Mission

Control Houston will command the activation

of the P4 truss to check out its systems and the

still folded solar arrays.

The BGAs will rotate to the correct position for

the deployment of the solar arrays on flight day

6.

EVA 2 

On flight day 5, Burbank and MacLean will

continue preparing the SARJ for rotation by

releasing and removing 16 launch locks, six of

10 outer launch restraints.  The launch locks

and launch restraints constrain the SARJ and

handle loads during ascent.  All of the launch

locks must be removed before any of the launch

restraints can be removed.  In addition, the

drive lock assemblies must be deployed to

provide a method of controlling uncontrolled

SARJ rotation.

Both Burbank and MacLean will work

simultaneously to remove the launch locks.  

Each launch lock is located under a separate

insulation cover which is in turn connected to

the SARJ inboard bulkhead by four to six bolts,

and connected to the outboard bulkhead by one

to three spring‐loaded clamp bolts.  The cover

must be removed to access the launch lock.  

After removing the cover, the launch lock is

removed by releasing four bolts.  Once the

launch lock is removed, the cover is replaced

and reattached to the SARJ inboard bulkhead.  

The outboard spring clamp bolts are left open

to allow for SARJ rotation.

After the launch locks are removed, Burbank

and MacLean will begin removing the SARJ

launch restraints.  These restraints hold

together the inner and outer SARJ bulkheads.  

There are two additional launch restraints

holding the forward face nadir and zenith SARJ

stub rails to the inboard SARJ bulkhead as well

as two additional launch restraints on aft truss

beams.

The final task of EVA 2 will be the deployment

of the SARJ brace beams.  These beams are

located on the P3 inboard side of the SARJ.  

These beams also help rigidize the SARJ

interface.
August 2006 SPACEWALKS 30

Once these tasks are completed, the SARJ can

be rotated and Mission Control will command a

checkout that involves as series of tests,

including rotating the SARJ a full 360 degrees,

plus another 180 degrees to position the P4

photovoltaic radiator (PVR) pointing toward

Earth.  This position corresponds to the 0

degree SARJ position.

The solar array and SARJ preparation work

must be completed by the end of EVA 2 so the

SARJ can be checked out allowing the solar

arrays’ deployment on flight day 6.

EVA 3 

On flight day 7, Tanner and Piper will again

don their spacesuits for the third and final

spacewalk.  During the six‐and‐one‐half‐hour

EVA, the crew will install bolt retainers on the

P6 beta gimbal assembly (BGA), prepare the P4

PVR, a photovoltaic or heat‐dissipating

radiator, for deployment, clear the path on P3’s

forward face for the MT’s move to work site 8,

remove and replace the S1 S‐band antenna

support assembly (SASA), remove and replace

the string 1 baseband signal processor (BSP)

and transponder on the S1 truss, install the

antenna group interface tube (AGIT) heat

shield, and retrieve the Materials for ISS

Experiment (MISSE) 5 from P6.

Before moving to P4, Tanner will move to the

P6 zenith bulkhead to install eight bolt retainers

on the P6 beta gimbal assemblies and engage a

remaining P6 4‐bar hinge lock.  The bolt

retainers will prevent the potential backing out

of those bolts.  One of the eight hinge locks did

not engage during the STS‐97 mission.  

Although previous attempts to engage the lock

were unsuccessful, the team believes they have

the correct tool for the non‐standard task.   

Piper will spend some time preparing the work

sites for the SASA task.

The crew then will move to P4 where they will

prepare the P4 radiator for deployment by

removing the cinches and winches on the

radiator.  These must be released before the

crew inside the station can deploy the radiator.

The cinches are wire braided cables with nut

assemblies on the end that serve as launch

restraints for the PVR.  One end of the cable is

permanently attached to the PVR base plate, the

opposite end contains an adjustable nut

assembly that slides into a receptacle on the

outermost PVR panel.

The nut assemblies secure the PVR for launch.  

The EVA crew will use a pistol grip tool (PGT)

on the nuts to release the tension in the cable,

remove the cable nut assembly from its

receptacle on the PVR, and attach the cinch to a

clip on the PVR base plate.

The next task will be to release the winch bar,

which secured the PVR during launch, from the

PVR.  A PIP pin secures the winch bar to the

outermost PVR panel.  Tanner will release one

of the winch PIP pins while Piper releases the

other.  After the winch bar is released from the

PVR, the pip pin will be reinstalled into the
August 2006 SPACEWALKS 31

winch bar.  The radiator will then be ready for

deployment.

Tanner and Piper will also ʺclearʺ the P3

forward face so the mobile transporter (MT) can

move onto the P3 truss.  The keel pin and drag

link are located on P3 face 1.  Tanner and Piper

will remove the drag link and keel pin and stow

them within the P3 structure.  Tanner and Piper

will remove the P3 space vision system (SVS)

target, as well as rotate P1, P3 MT stops and

tether shuttle stops.  In addition, the crew will

install the EVA temporary rail stop (ETRS).  The

ETRS will now perform the functions of the

other MT and tether shuttle stop mechanisms

while still allowing a Crew and Equipment

Translation Aid (CETA) cart to be outboard of

the MT/MBS while the MT/MBS is at the P3

work site.  The crew will also relocate the

articulating portable foot restraints (APFR)

used during their tasks to the P4 truss where

they will not interfere with the MT movement.

After completing their P3 and P4 tasks, Tanner

and Piper will move back to the Z1 truss via the

airlock.  The EVA crew will start working on

the removal and replacement of the SASA on S1

with a spare unit that is located on the Z1 truss.  

The task will require several handoffs between

the crew to move the spare SASA up to the S1

truss.  The crew will also need to temporarily

stow the failed SASA while the spare SASA is

installed.  The crew will reverse their path and

use the same technique to move the failed

SASA back to the Z1 truss.  The failed SASA

will be returned to Earth on a later mission.

Piper will retrieve upgraded S‐band BSP and

transponders from the airlock.  The new Orbital

Replacement Units (ORUs) will be carried to

the work site in a large ORU transfer bag.  Piper

will temporarily stow the bag on the S1 truss

while she changes out the S‐band transponder

and base‐band signal processor (BSP).  She will

return to the airlock with the older version BSP

and transponder.

While Piper is busy on the S1 truss, Tanner will

install a heat shield on the Ku‐band antenna

group interface tube (AGIT).  This will help

prevent over heating in this area during certain

ISS attitudes.  Tanner will then return to the P6

truss to remove MISSE 5 from the P6 truss for

return to Earth.
August 2006 PAYLOAD OVERVIEW 32

PAYLOAD OVERVIEW 

Space Shuttle Atlantis will carry equipment

and supplies in its cargo bay to the

International Space Station.  Additional items

will be carried on the shuttle middeck,

including supplies, food, water and clothing for

the station crew.

STAGE 12A ASSEMBLY HARDWARE 

The primary element to be delivered is the

P3/P4 Integrated Truss Segment.  The P3/P4

truss segment provides an additional

photovoltaic module (PVM) with two power

channels (2A and 4A), the port Solar Alpha

Rotary Joint (SARJ) and two Unpressurized

Cargo Carrier Attachment System sites.  The

P3/P4 segment is 45.3 ft long and weighs 34,885

pounds.

Integrated Truss Segment P3/P4 

The P3 structure has an elongated hexagonal

cross section with two bays arrayed along the

long axis.  The P4 truss segment is integrated to

the P3 segment before flight.  A detailed

description of this prime payload for the

STS‐115 mission can be found on page 39.
August 2006 PAYLOAD OVERVIEW 33

Solar Alpha Rotary Joint 

The Solar Alpha Rotary Joint (SARJ) enables the

new outboard solar arrays to always point to

the sun by rotating like a Ferris wheel.  The port

SARJ on P3/P4 provides tracking for the P4 and

P6 solar arrays once P6 is relocated to P5 on the

STS‐120 mission.  The starboard SARJ for the

S3/S4 arrays, to be delivered on STS‐117, will

provide tracking for the S4 and S6 solar arrays

once S6 is installed on the STS‐119 mission.  The

SARJ can rotate 360 degrees clockwise and

counterclockwise.  The SARJ also provides the

structural interface between the P3 or S3 and P4

or S4 elements.  It includes hardware to route

power and data through the rotating SARJ

interface to the outboard truss segments.

Race Rings 

The SARJ includes an inboard and an outboard

race ring, which provide the structural

connection between the P3 and P4 elements.  

Along the circumference of each race ring are

gear teeth that mesh with the drive lock

assembly (DLA) pinion gear to rotate the SARJ.  

The outboard race ring is used by the DLA for

rotating the SARJ.

Trundle Bearings 

The 12 equally‐spaced trundle bearings hold

the SARJ inboard and outboard race rings

together.  Each trundle bearing is fixed to the

inboard race ring and is clamped onto the

outboard race ring with a roller interface to

allow for SARJ rotation.  There are three rollers

on the trundle bearing that interface with the

outboard race ring; the inner and outer upper

rollers and the center roller.  Each roller consists

of two bearings:  the primary and journal

bearings.  The primary bearing rotates.  If the

primary bearing seizes up, the journal bearings

will begin rotating.  The journal bearing is

designed to operate for about 30 days.  There

are micro‐switches in the trundle bearing that

allow the ground to know if the journal bearing

is rotating.

Drive Lock Assembly 

Two Drive Lock Assemblies (DLAs) are

responsible for rotating and locking the SARJ.  

Each DLA includes the engage‐disengage

mechanism (EDM) motor, drive motor, pinion

gear, lock rack and two follower arms.  The

EDM motor is a stepper motor that pivots the

lock rack and pinion gear about the lock/engage

pivot point to the desired position.  The DLA

positions are locked, engaged and neutral.  For

the locked position, the lock rack is in contact

with the race ring gear teeth to prevent the
August 2006 PAYLOAD OVERVIEW 34

SARJ from rotating.  For the engaged position,

the DLA pinion gear is meshed with the race

ring gear to rotate the SARJ with the drive

motor.  In the neutral position, neither the lock

rack nor pinion gear is engaged to the race ring

gear.  The follower arms, which are of the same

design as the trundle bearings, are used to

secure the DLA to the race ring.

Rotary Joint Motor Controller 

The two rotary joint motor controllers (RJMCs)

are mounted on the structural ribs of the

inboard race ring and control the operation of

the DLA motors via commands from the P3

multiplexers/demultiplexers (MDMs).  Each

RJMC has two heaters and a resistive thermal

device sensor monitored by the P3 MDMs.  The

RJMC supplies and receives voltage signals

from the resolvers on the Utility Transfer

Assembly to determine the position of the SARJ

(resolver A is connected to RJMC 1, resolver B

is connected to RJMC 2).  For SARJ outboard

operations, one RJMC is moved to the outboard

race ring.

SARJ Structural Support 

At launch, the SARJ is secured with 16 launch

locks and 10 launch restraints.  All the launch

locks and the six outboard launch restraints will

be removed on orbit before the SARJ is rotated,

first in a short test, then in its operational

configuration.  On the longerons for two sides

of the SARJ, there are stub rail segments that

bridge over the interface that enables the SARJ

to rotate.  The launch restraints on the inboard

and outboard side hold the stub rail segments

in place.  When the SARJ is operating in the

inboard mode, the outboard launch restraints

are removed and the inboard launch restraints

are left in place so that the stubs rails are

cantilevered over the rotating SARJ interface.  

The SARJ inboard bulkhead is structurally

attached to P3 at the P3 longerons and four

SARJ braces.  The braces are stowed on the

diagonals and not connected to the inboard

bulkhead for launch so that the SARJ loads path

is supported entirely by the launch restraints.

Integrated Truss Segment P4 

The major components of the P4 truss are the

port inboard photovoltaic module and

modified Rocketdyne truss attachment system.

Photovoltaic Module 

The P4 photovoltaic module contains two

power channels, 2A and 4A.  The individual

components of these power channels are

described in more detail below.

Solar Array Wing (SAW):  Each power channel

begins with the solar array wing where

electrical generation takes place.  Each solar

array wing is divided in half by its deployment

mast into two PV blankets (right and left).  Each

blanket contains 16,400 individual silicon PV

cells, eight square centimeters apiece, for the

conversion of sunlight into electricity.  The

photovoltaic cells are grouped into 82 active

panels ‐‐ each consisting of 200 cells connected

in series.  The active panels are wired in sets of

two, providing a total of 41 independent

circuits, or strings, per blanket.  Two flat

collector circuits run along the sides of each
August 2006 PAYLOAD OVERVIEW 35

blanket, providing connectivity between two

active panels.  The circuits also route primary

power from each string to the containment box

wire harness, which forms the blanket power

output.  This setup allows each string to be

treated as an independent power source.  Two

inactive panels are located at the top and

bottom of each blanket to prevent the

shadowing of the arrays and to protect the

active panels while the arrays are stowed.  They

also provide space between them and the solar

array wing, which must be directly in line with

the sun whenever possible for maximum power

generation.  This is accomplished by

maneuvering the station into certain

orientations, as well as by positioning the

arrays using the Beta Gimbal Assemblies

(BGAs) and the SARJ.

Sequential Shunt Unit (SSU):  The primary

power output of each solar array wing is

divided into 82 individual paths, or strings—41

per blanket—which are routed through the

solar array wing mast canister to the SSU

located at the base of the canister.  The main

purpose of the SSU is to regulate the voltage

level of the primary power bus, the input from

the solar array wing.  It also can be used to

isolate the array from the rest of the electrical

power system during contingency or

maintenance operations.  The solar arrays are

controlled via the photovoltaic controller

application (PVCA) software within the

photovoltaic controller unit (PVCU) routing

system.  To regulate the primary voltage level, a

closed‐loop control system is used to achieve a

voltage set point.  The voltage set point is sent

by the PVCA to the three photovoltaic

controller elements (PVCEs) within the SSU.  

Closed‐loop control is achieved mainly through

the SSU Shunt Electronics Module (SEM) and

the input of the three PVCEs.

These internal electronics are powered by two

sources:  the channel primary power bus and a

direct current control power bus (provided by a

battery charge/discharge unit on the adjacent

power channel).  The PVCEs compare the SSU

output voltage (the sum of the 82 strings) with

the PVCA set point and sends a signal to the

SEM, which decides how many strings need to

be routed back to the array or deselected to

achieve the voltage set point.  Shunting is

performed through field effect transistors

(FETs) connected to each string within the SEM.  

The process always allows one transitional

string for exceeding the set point.  The

transitional string is continuously shunted and

unshunted at a specified rate in a process called

dithering that results in fine voltage level

control.

Beta Gimbal Assembly (BGA):  The Beta Gimbal

Assembly is the mechanism used to orient the

face of each solar array wing toward the sun to

maximize the power generation of the solar

cells.  The BGAs provide rotation about the axis

along the mast.  The BGAs also provide

structural support for the SAW and

accommodate power and data transfer through

its rotating interface.  One BGA per power

channel (or SAW) is located between the base of

the mast canister and the Integrated Equipment

Assembly (IEA).  The BGA is comprised of four

main components:  the Bearing Motor and Roll

Ring Module (BMRRM), the Beta Gimbal

Housing Subassembly (BGHS), the Beta Gimbal

Deployment Transition Structure (BGDTS) and

the Beta Gimbal Platform (BGP).

Bearing Motor and Roll Ring Module (BMRRM):  

The BMRRM is the most complex and active

component of the BGA.  Its main purpose is to

provide sun tracking capability for the arrays

while transferring power, commands, and data

bi‐directionally across the rotational joint.  The
August 2006 PAYLOAD OVERVIEW 36

BMRRM consists of module housing, two sets

of angular contact bearings, a brushless direct

current motor drive, a roll ring assembly, and

an anti‐rotation latch assembly (two latches).

The BGA allows transfer of power and data

through its 360‐degree continual rotation joint

by the use of roll rings in the BMRRM.  These

roll rings are metallic flexures, which ensure

that a metal contact point is maintained,

allowing power and data to be transferred, no

matter what the angle of rotation is.

Modified Rocketdyne Truss Attachment 

System 

The P4 element includes the active side of the

P4‐P5 Modified Rocketdyne Truss Attachment

System (MRTAS), which will be used to attach

the P5 element to P4 on the next shuttle

mission, STS‐116.  Unlike the other type of

attachment system that mated the P6 array

truss to the Z1 truss and will be used to attach

P6 to P5, the MRTAS does not include a capture

claw since, due to the P4 BGA locations, there is

no room to accommodate it.  The P5 is soft‐

docked to P4 by positioning the four P5 coarse

cones (one on each corner) into receptacles on

P4.  Once the P5 element is soft‐docked, the

crew secures the primary bolt at each corner on

P5 into the nut assemblies on P4.  If the primary

bolt cannot be secured, two contingency bolts at

each corner on P5 can be tightened into the nut

assemblies on P4.

P3/P4 Preparation 

During Atlantis’ rendezvous with the

International Space Station, the shuttle crew

performs a visual inspection of the P3/P4 truss

segment using payload bay and shuttle robotic

arm cameras.  The crew also will observe

through the aft flight deck windows to verify

that the hardware survived the launch intact.  

The crew has no telemetry insight into the

health and status of the truss hardware until

after it is installed with its connectors mated.  

P3/P4 installation and activation is scheduled to

occur the day after docking.  Installation is

performed using both the shuttle robotic arm

and the station robotic arm, Canadarm2.

On flight day 3 after docking, the shuttle

robotic arm will be used to unberth P3/P4 from

the shuttle payload bay and will hand it off to

the station arm.  The station arm will operate

from the Mobile Transporter Power and Data

Grapple Fixture 3 located at work site 7 on the

far port side of the truss.  P3/P4 will spend the

night grappled by the station arm.  The next

day, the P3/P4 is maneuvered to the P1 install

position.  After the installation, a spacewalk is

performed out of stationʹs Quest airlock.  The

spacewalk is critical to the activation and

survival of the P3/P4 integrated hardware.

After P3/P4 is unberthed, many of its hardware

components are vulnerable to the elements of

space, and it is critical to expedite the

installation and application of keep‐alive power

to those components.  A number of hardware

checkouts are required before proceeding with

P3/P4 unberthing from the payload bay.  The

shuttle and station robotic arms support one

another during all phases of the installation

process.  The shuttle robotic arm provides

camera views and additional Space Vision

System (SVS) targeting functions, while the

station arm manipulates the P3/P4 element.  

Also, a number of station systems are

reconfigured and powered down to support the

P3/P4 zenith tray umbilical connections.

P3/P4 Installation 

During P3/P4 unberthing from the payload bay,

Pilot Chris Ferguson and Mission Specialist
August 2006 PAYLOAD OVERVIEW 37

Dan Burbank first maneuver the shuttle arm to

the P3/P4 grapple fixture located on the orbiter

port side and grapple the truss segment.  When

the cameras and viewing equipment are

properly configured for unberthing, the four

orbiter longeron payload retention latch

assemblies and two active keel assemblies are

opened, and the truss is lifted out of the

payload bay.

During unberthing, another crew member

monitors the clearance between the P3/P4 truss

element, the orbiter aft bulkhead and the

Destiny laboratory as the truss is moved using

a lift and tilt maneuver.  The truss is then

swung out over the port side of the orbiter and

maneuvered to the P3/P4 handoff position.  

Mission Specialist Steve MacLean and station

flight engineer Jeff Williams use Canadarm2 to

grapple P3/P4 at the other grapple fixture on

the starboard side.  When the station arm

grapples P3/P4, the shuttle arm releases its

grapple fixture and maneuvers out of the way.  

The cameras are reconfigured for the

installation.

Working from the station’s robotics work

station, MacLean then maneuvers the truss to

the P3/P4 preinstall position on the port side of

P1.  P3/P4 is now in the proper orientation for

installation.  Just inches away from installation,

all shuttle and station thrusters are inhibited as

mating operations begin.  From the “preinstall”

position, MacLean slowly maneuvers the P3/P4

truss segment in toward the port side of P1

using computer displays and shuttle arm

cameras to assist with alignment and

orientation.  At the same time, the P1 active

attachment power/data interfaces and

hardware are verified, and the attachment

software state is transitioned from standby to

manual.

Within 8 to 10 cm (3 to 4 in.) of mating, the

three ready‐to‐latch (RTL) sensors on the P1

side of the truss interface contact the three RTL

striker plates (passive side), which sends a

signal to computers providing a positive

indication to the crew that the capture latch is

ready to be closed.  Each RTL sensor contains

dual redundant micro‐switches, and only one

switch indication is required to proceed with

capture latch actuation.

The margin for the precise alignment of the

attachment device’s capture latches is plus or

minus 3 inches in either direction.

The capture of the P3/P4 and the P1 segments

occurs in two stages with four huge bolts

driven to form a hard mate between the two

truss components.

Activation and Checkout 

The P3/P4 activation operations are the most

complex of the entire mission and are the most

challenging for both the crew and flight

controllers.  The activation of the newly

installed P3/P4 segments involves powering up

and checking out all integrated hardware

within the P3/P4 truss segment.

Following P3/P4 installation to the P1 truss, a

number of activities are required to apply the

keep‐alive power that protects the P3/P4

hardware from freezing.  The P1‐to‐P3 lower

tray utility cable connections are required to

power P3/P4 component heaters.  These

connections, along with similar connections for

the upper tray utility cables will be

accomplished during the first spacewalk by Joe

Tanner and Heidemarie Stefanyshn‐Piper.
August 2006 PAYLOAD OVERVIEW 38

Lower Tray Utility Connections 

(Channel 2/3) 

In parallel with EVA preparation tasks and

P3/P4 installation operations, the ground

temporarily shuts down power to the S0, S1,

and P1 truss systems on power channel 2/3.  

When all P1‐to‐P3 lower tray connections are

made, the ground reactivates that power

channel 2/3 for the S0, S1 and P1 truss segments

in parallel with other spacewalk activities.

Upper Tray Utility Connections 

(Channel 4/5) 

When the channel 2/3 power‐up is complete

and all critical S0, S1 and P1 systems are

verified, all S0, S1, and P1 systems on power

channel 1/4 are then shut off in preparation for

the upper tray (EVA) utility cable connections

between P1 and P3.  Also, a truss thermal clock

is started due to the loss of truss controlled

heaters.  The second channel is reactivated once

all the cable connections are made and verified.

P4 Initial Activation and Battery 

Conditioning/Charging 

P4 activation can begin after circuit breakers are

turned on following the P1‐P3 lower tray

connections.  P4 channel 4A activation can

begin after circuit breakers 6 and 8 are removed

following the P3/P4 upper tray connections.  

Closing or removing these breakers enables P6

channel 2B to supply P4 channel 2A with

startup power.  The P4 is activated on flight day

4.

P4 Solar Array Deploy 

During the P6 channel 2B SAW deploy on the

STS‐97 mission in 2001, the solar array panels

experienced a phenomenon called “stiction”

during deploy.  The array “stiction,” or

stickiness between blankets, caused the array

tension bar to slam back into the lower half

blanket box intermittently.  It resulted in the

array flapping and waving against its tension

bar.  It was determined post‐flight that the

array “stiction” was caused by an atomic

oxygen coating on the panels that fused

together over time.  The severity of the array

“stiction” is directly proportional to the length

of time the panels are stored inside the blanket

boxes, as well as the temperature of the panels

at deployment.  To avoid this problem for the

P4 solar array deployments, a new deploy

procedure was developed.  The blanket boxes

are deployed in high‐tension mode to avoid

large movements of the tension bar.  The arrays

are also thermally conditioned to minimize

panel “stiction”.  Each solar array is also

deployed during periods of low or no sunlight.  

The arrays are deployed on flight day 6.

P4 Photovoltaic Radiator Deploy 

The P4 photovoltaic radiator (PVR) is deployed

by commands sent by Jett and Ferguson from

computers on the shuttle flight deck on flight

day 7 after the radiator cinches and winches are

released during the third spacewalk.  The

radiator is required to cool the arrays’

electronics when the P4 is providing power to

station components.  The P4 power channels

will be provide power for P4 use only until the

station’s electrical system is rewired during the

STS‐116 mission.

Solar Alpha Rotary Joint Activation and 

Checkout 

The SARJ launch and pre‐activation orientation

is at an angle of 180 degrees to its own axis with

both DLAs locked.  Upon completion of its

activation and checkout, the SARJ will have

rotated slightly to an angle of 0 degrees relative

to its axis with one DLA engaged and one DLA
August 2006 PAYLOAD OVERVIEW 39

locked.  The goals of this flight’s SARJ

activation and checkout are to rotate to the

solar array wing deploy position and lock,

verify the functionality of each SARJ power

system and demonstrate its ability to accurately

position itself to track the sun.

The SARJ checkout begins near the end of the

second spacewalk on flight day 5 and continues

through crew sleep prior to solar array

deployment.  The checkout is scheduled to take

up to 12 hours, but could be completed in less

time, if the systems operate as expected.

External Wireless Instrumentation 

System (EWIS) 

With the launch of the P4 element, NASA will

deploy a new External Wireless

Instrumentation System (EWIS) for the first

time.  The system consists of accelerometers

placed around the outboard truss elements of

the integrated truss structure.  The system

allows engineers to gather real‐time data

during dynamic events that might cause higher

vibrations (loads) on the truss elements.  For

more information on the EWIS, see page 50.

PAYLOADS OF OPPORTUNITY 

Ram Burn Observation (RAMBO) 

Ram Burn Observations is a Department of

Defense experiment that observes shuttle

Orbital Maneuvering System engine burns for

the purpose of improving plume models.  On

manifested missions, sensors on a satellite will

observe selected rendezvous and orbit adjust

burns.

Maui Analysis of Upper-Atmospheric 

Injections (MAUI) 

MAUI will observe the space shuttle exhaust

plumes from the Maui Space Surveillance Site

(MSSS).  The observations will occur when the

shuttle fires its engines at night or twilight over

MSSS.  Spectrally filtered images and spectra of

the radiation resulting from exhaust‐

atmosphere interactions will be taken by the

optical telescope and all‐sky imagers.  This will

reveal the chemical and physical mechanisms

associated with the interaction between the

chemical species in engine exhaust and the

space environment.  The improved models of

this interaction will result in enhanced space

event characterization as well as the

determination of sensor requirements for

effective plume and contamination analysis of

other spacecraft.

P3 AND P4 TO EXPAND STATION 

CAPABILITIES, PROVIDING A THIRD 

AND FOURTH SOLAR ARRAY 

The Port Three (P3) and Port Four (P4) Truss

segments are slated for launch to the

International Space Station aboard Space

Shuttle Atlantis no earlier than August 27, 2006,

from Kennedy Space Center, Fla.  The truss is

the next major addition to the 11‐segment

integrated truss structure that will eventually

span more than 300 feet to carry power, data

and temperature control for the orbital

outpost’s electronics.
August 2006 PAYLOAD OVERVIEW 40

P3/P4 Specifications 

Dimensions: 45.3 ft long by 16 feet 

wide by 15.6 feet high 

Weight: 34,885 lbs 

The integrated truss segments started with

Starboard zero (S0) as the center assignment

and were numbered in ascending order

outward to the port and starboard sides.   At

one time, there was an S2 and P2 planned, but

these were eliminated when the station design

was scaled back.  From S0, the truss segments

are P1, P3, P4, P5 and P6 and S1, S3, S4, S5, and

S6.  P6 is presently on orbit and attached to

segment Z1.  P6 will be eventually relocated

and attached to P5.

The primary cargo delivered on STS‐

115/Mission 12A is the Integrated Truss

Segment P3/P4, which mates to P1 and

provides an attachment point for P5.  These

new segments also provide a second set of Solar

Array Wings (SAWs) and the first alpha joint.  

The segments also support utility routing,

power distribution, and a travel path for the

Mobile Remote Service Base System (MBS).

P3 and P4 will be removed from the space

shuttle payload bay using the shuttle arm

(SRMS – Shuttle Remote Manipulator System)

and it will be handed off to the station arm

(SSRMS – Space Station Remote Manipulator

System), where it will be maneuvered and

attached to P1.

The P3 primary structure is made of a

hexagonal shaped aluminum structure and

includes four bulkheads and six longerons.  The

secondary structure includes brackets, fittings,

attach platforms, EVA equipment and

miscellaneous mechanisms.

Major P3 subsystems include the Segment‐to‐

Segment Attach System (SSAS), Solar Alpha

Rotary Joint (SARJ), and Unpressurized Cargo

Carrier Attach System (UCCAS).  The primary

functions of the P3 truss segment are to provide

mechanical, power and data interfaces to

payloads attached to the two UCCAS

platforms; axial indexing for solar tracking, or

rotating of the arrays to follow the sun, via the

SARJ; movement and work site

accommodations for the Mobile Transporter;

accommodations for ammonia servicing of the

outboard PV modules and two

Multiplexer/Demuliplexers (MDMs) which are

basically computers.  The P3 also provides a

passive attachment point to the P1 segment via

the SSAS and pass through of power and data

to and from the outboard segments.

The UCCAS will allow platforms to be attached

to P3 for the storage of additional science

payloads or spare Orbital Replacement Units.  

It has a capture latch to grip and secure a

payload, a berthing target to align payloads to

the mechanism and an Umbilical Mechanism

Assembly, which has a connector that provides

power and data to the payload.

P3/P4 shown attached to the ISS at the  

far right of this picture
August 2006 PAYLOAD OVERVIEW 41

Major subsystems of the P4 Photovoltaic

Module (PVM), include the Photovoltaic

Radiator (PVR), the Alpha Joint Interface

Structure (AJIS), and Modified Rocketdyne

Truss Attachment System (MRTAS).  The P4

Photovoltaic Module includes all equipment

outboard of the Solar Alpha Rotary Joint (SARJ)

outboard bulkhead, namely the two

Photovoltaic Array Assemblies (PVAAs) and

the Integrated Equipment Assembly (IEA).  

Each PVAA consists of a Solar Array Wing

(SAW) and Beta Gimbal Assembly (BGA).  The

PVR provides thermal cooling for the IEA.  The

Alpha Joint Interface Structure (AJIS) provides

the structural transition between P3 and P4

structures.  P4 contains the passive side of the

MRTAS, which will provide the structural

attachment for P5 on P4.

P3 consists of the Solar Alpha Rotary Joint

(SARJ), which continuously rotates to keep the

Solar Array Wings (SAW) on P4 and P6

oriented towards the sun as the station orbits

the Earth.  Each SAW is also oriented by the

BGA, which can change the pitch of the wing.  

Each wing measures 115 feet by 38 feet and

extends out to each side of the Integrated

Equipment Assembly.  There are two wings on

P4.

The P3/P4 integrated truss structure is the

primary payload for the STS‐115 mission and

contains several discrete elements:  two Solar

Array Wings (SAW), Integrated Equipment

Assembly (IEA), Solar Alpha Rotary Joint

(SARJ), two Beta Gimbal Assemblies (BGA) and

Photovoltaic Thermal Control Subsystem.

Following are specific details on each of the

major elements:  
August 2006 PAYLOAD OVERVIEW 42

P4 Photovoltaic Module (power module)

Electrical power is the most critical resource for

the ISS because it allows astronauts to live

comfortably, safely operate the station, and

perform complex scientific experiments.  Since

the only readily available source of energy for

spacecraft is sunlight, technologies were

developed to efficiently convert solar energy to

electrical power.  One way to do this is by using

large numbers of solar cells assembled into

arrays to produce high power levels.  The cells

are made from purified crystal ingots of silicon

that directly convert light to electricity through

a process called photovoltaics.  Since a

spacecraft orbiting the Earth is not always in

direct sunlight, energy has to be stored.  Storing

power in rechargeable batteries provides a

continuous source of electricity while the

spacecraft is in the Earthʹs shadow.

NASA and Lockheed Martin developed a

method of mounting the solar arrays on a

ʺblanketʺ that can be folded like an accordion

for delivery to space.  Once in orbit, astronauts

deploy the blankets to their full size.  Gimbals

are used to rotate the arrays so that they face

the sun to provide maximum power to the

Space Station.  The solar arrays track the sun in

two axes:  beta and alpha.  The complete power

system, consisting of U.S. and Russian

hardware, will generate 75 to 110 kW

(kilowatts) total power, about as much as 55

houses would typically use.

P4 is the second of four PVMs that will

eventually be brought up to the International

Space Station, to convert sunlight to electricity.  

The first one, named P6, was brought on orbit

by the STS‐97 crew in November 2000.  The

primary functions of the P4 photovoltaic

module are to collect, store, convert and

distribute electrical power to loads within the

segment and to other ISS segments.  The P4

PVM consists of two Beta Gimbal/PV Array

assemblies, two Beta Gimbal Transition

Structures, one integrated Equipment Assembly

and associated cabling and tubing.  The P4

PVM components were assembled by Boeing in

Tulsa, Okla. and Lockheed Martin in

Sunnyvale, Calif. prior to final assembly and

testing by Boeing at Kennedy Space Center, Fla.

There are two solar array wings (SAW)

designed, built and tested by Lockheed Martin

in Sunnyvale, Calif. on the P4 module, each

deployed in the opposite direction.  Each SAW

is made up of two solar blankets mounted to a

common mast.  Prior to deployment, each panel

is folded accordion style into a Solar Array

Blanket Box (SABB) measuring 20 inches high

and 15 feet in length.  Each blanket is only

about twenty inches thick while in this stored

position.  The mast consists of interlocking

battens that are stowed for launch inside a Mast

Canister Assembly (MCA) designed, built and

tested by ATK‐Able.  When deployed by the

astronauts, the SAW deploys like an erector set

as it unfolds.  It has two arms like a human

torso when mounted on P4.  These arms will be

rotated outwards by astronauts during planned

spacewalks so they can be fully deployed.  

Because these blankets were stored for such a

long time, NASA, Boeing and Lockheed Martin
August 2006 PAYLOAD OVERVIEW 43

conducted extensive testing to ensure they

would unfold properly once on orbit to ensure

there would be no problems with the blankets

sticking together.  This testing was completed

in July 2003.

When fully deployed, the SAW extends 115 feet

and spans 38 feet across and extends out to

each side of the Integrated Equipment

Assembly.  Since the second SAW is deployed

in the opposite direction, the total wing span is

over 240 feet.

Each Solar Array Wing weighs over 2,400

pounds and use nearly 33,000 (32,800 per wing)

solar array cells, each measuring 8‐cm square

with 4,100 diodes.  The individual cells were

made by Spectrolab and ASEC.  There are 400

solar array cells to a string and there are 82

strings per wing.  Each SAW is capable of

generating nearly 32.8 Kilowatts (kW) of direct

current power.  There are two SAWs on the P4

module yielding a total power generation

capability approaching 66 kW, enough power

to meet the needs of 30 average homes –

without air conditioning (based on an average

2kW of power.)

P3 consists of the Solar Alpha Rotary Joint

(SARJ), which continuously rotates to keep the

solar array wings on P4 and P6 oriented

towards the sun as the station orbits the earth.  

Located between P3 and P4, the SARJ is a 10

foot diameter rotary joint that tracks the sun in

the alpha axis and turns the entire P4 module

(and eventually the P6 module when it is

relocated).  The SARJ weighs approximately

2,500 pounds.  The SARJ can spin 360 degrees

using bearing assemblies and a servo control

system to turn.    All of the power will flow

through the Utility Transfer Assembly (UTA) in

the SARJ.  Roll ring assemblies allow

transmission of data and power across the

rotating interface so it never has to unwind.  

The SARJ was designed, built and tested by

Lockheed Martin in Sunnyvale, Calif., and its

subcontractors.

The solar array wings are also oriented by the

Beta Gimbal Assembly (BGA), which can

change the pitch of the wings by spinning the

solar array.   The BGA measures 3 x 3 x 3 feet

and provides a structural link between the

Integrated Equipment Assembly (IEA).  The

BGA’s most visual functions are to deploy and

retract the SAW and rotate it around its
August 2006 PAYLOAD OVERVIEW 44

longitudinal axis.  The BGA consist of three

major components all mounted on a platform:  

the Bearing, Motor and Roll Ring Module

(BMRRM), the Electronic Control Unit (ECU)

and the Beta Gimbal Transition Structure.  The

BGA was designed by Boeing Rocketdyne in

Canoga Park, Calif., which has since been

acquired by Pratt and Whitney.  The Sequential

Shunt Unit (SSU) that serves to manage and

distribute the power generated from the arrays

is also mounted on each BGA Platform.  The

SSU was designed by Space Systems Loral.

Both the SARJ and BGA are pointing

mechanisms.  They can follow an angle target,

rotating in two axes to always point the surface

of the arrays at the target, the sun Ground

controllers continuously update the target

information so the joints keep rotating, moving

the arrays moving continuously as the station

orbits the Earth every 90 minutes.  The

movement of the rotation is at a rate that

maintains constant contact of the arrays with

the sun as the station circles Earth.  The SARJ

mechanism will move much more than the

BGA.  The BGA will move about four or five

degrees per day, whereas the SARJ will rotate

360 degrees every orbit or about 4 degrees per

minute.  The SARJ will be the first one to be

installed on station.  It is unique because it

rotates the entire truss element, allowing it to

rotate in the alpha axis rotation, a rotation that

moves the array in the direction of the stationʹs

travel.  The station has been using the P6 BGA

to move as an alpha joint.  Eventually, the SARJ

will provide primary rotation with BGA doing

minor movements and will be tested on this

flight, but won’t be activated until assembly

mission STS‐116.

P4 Integrated Equipment Assembly (IEA) 

The IEA has many components: 12 Battery

Subassembly orbital replacement units (ORUs),

six Battery Charge/Discharge Units (BCDU)

ORUs, two Direct Current Switching Units

(DCSUs), two Direct Current to Direct Current

Converter Units (DDCUs), two Photovoltaic

Controller Units (PVCUs).  It integrates the

Thermal Control Subsystem that consists of one

Photovoltaic Radiator (PVR) ORU and two

Pump Flow Control Subassembly (PFCS)

ORU’s used to transfer and dissipate heat

generated by the IEA ORU boxes.  In addition,

the IEA provides accommodation for ammonia

servicing of the outboard PV modules as well

as pass through of power, data to and from the

outboard truss elements.  The structural

transition between the P3 and P4 segments is

provided by the Alpha Joint Interface Structure.

The Integrated Equipment Assembly measures

16 x 16 x 16 feet, weighs nearly 17,000 pounds

and is designed to condition and store the

electrical power collected by the photovoltaic

arrays for use on board the Station.

The IEA integrates the energy storage

subsystem, the electrical distribution

equipment, the thermal control system, and

structural framework.  The IEA consists of three

major elements:

1. The power system electronics consisting of

the Direct Current Switching Unit (DCSU)

used for primary power distribution; the

Direct Current to Direct Current Converter

Unit (DDCU) used to produce regulated

secondary power; the Battery

Charge/Discharge Unit (BCDU) used to

control the charging and discharging of the

storage batteries; and the batteries used to

store power.
August 2006 PAYLOAD OVERVIEW 45

2. The Photovoltaic Thermal Control System

(PVTCS) consisting of:  the coldplate

subassembly used to transfer heat from an

electronic box to the coolant; the Pump

Flow Control Subassembly (PFCS) used to

pump and control the flow of ammonia

coolant; and the Photovoltaic Radiator

(PVR) used to dissipate the heat into deep

space.  Ammonia has significantly greater

heat transfer properties than other chemical

coolants.

3. The computers used to control the P4

module ORUs consisting of two

Photovoltaic Controller Unit (PVCU)

Multiplexer/Demultiplexers (MDMs).

Power received from each PVAA is fed directly

into the appropriate Direct Current Switching

Unit (DCSU).  The DCSU is a high‐power,

multi‐path remotely controlled unit that is used

for primary and secondary power distribution,

protection and fault isolation within the IEA.  It

also distributes primary power to the ISS.  

During periods of isolation, or sunlight, the

DCSU routes primary power directly to the ISS

from its PVAA and also routes power to the

power storage system for battery charging.  

During periods of eclipse, the DCSU routes

power from the power storage system to the

ISS.  The DCSU measures 28” by 40” by 12” and

weighs 238 pounds.

Primary power from the DCSU is also

distributed to the Direct Current to Direct

Current Converter Unit (DDCU).  The DDCU is

a power processing system that conditions the

coarsely regulated power from the PVAA to

123 +/‐ 2 VDC.  It has a maximum power output

of 6.25 kW.  This power is used for all P4

operations employing secondary power.  By

transmitting power at higher voltages and

stepping it down to lower voltages where the

power is to be used, much like municipal

power systems, the station can use smaller

wires to transmit this electrical power and thus

reduce launch loads.  The converters also

isolate the secondary system from the primary

system and maintain uniform power quality

throughout the station.  The DDCU measures

27.25” by 23” by 12” and weighs 129 pounds.

Primary power from the DCSU is also

distributed to the three power storage systems

located within each channel of the IEA.  The

power storage system consists of a Battery

Charge/Discharge Unit (BCDU) and two

Battery Subassembly ORUs.

The BCDU serves a dual function of charging

the batteries during solar collection periods,

and providing conditioned battery power to the

primary power busses (via the DCSU) during

eclipse periods.  The BCDU has a battery

charging capability of 8.4 kW and a discharge

capability of 6.6 kW.  The BCDU also includes

provisions for battery status monitoring and

protection from power circuit faults.  The

PVCU commands the BCDU.  The BCDU

measures 28” by 40” by 12” and weighs 235

pounds.

Each Battery Subassembly ORU consists of 38

lightweight Nickel Hydrogen cells and

associated electrical and mechanical equipment.  

Two battery Subassembly ORUs connected in

series are capable of storing a total of 8 kW of

electrical power.  This power is fed to the ISS

via the BCDU and DCSU respectively.  The

batteries have a design life of 6.5 years and can

exceed 38,000 charge/discharge cycles down to

a 35% level of charge.  Each battery measures

41” by 37” by 19” and weighs 372 pounds.  

Because of delays in launching the P3/P4

elements, the batteries were replaced in March

2005 for the lower deck and August 2005 for the

upper deck.
August 2006 PAYLOAD OVERVIEW 46

In order to maintain the IEA electronics at safe

operating temperatures in the harsh space

environment, they are conditioned by the

Photovoltaic Thermal Control System (PVTCS).  

The PVTCS consist of ammonia coolant, eleven

coldplates, two Pump Flow Control

Subassemblies (PFCS) and one Photovoltaic

Radiator (PVR).

The coldplate subassemblies are an integral

part of IEA structural framework.  Heat is

transferred from the IEA orbital replacement

unit (ORU) electronic boxes to the coldplates

via fine interweaving fins located on both the

coldplate and the electronic boxes.  The fins add

lateral structural stiffness to the coldplates in

addition to increasing the available heat

transfer area.

The PFCS is the heart of the thermal system.  It

consists of all the pumping capacity, valves and

controls required to pump the heat transfer

fluid to the heat exchanges and radiator, and

regulate the temperature of the thermal control

system ammonia coolant.  The PVTCS is

designed to dissipate 6,000 Watts of heat per

orbit on average and is commanded by the IEA

computer.  Each PFCS consumes 275 Watts

during normal operations and measures

approximately 40 x 29 x 19 inches, weighing 235

pounds.

The PVR – the radiator – is deployable on orbit

and comprised of two separate flow paths

through seven panels.  Each flow path is

independent and is connected to one of the two

PFCSs on the IEA.  In total, the PVR can reject

up to 14 kW of heat into deep space.  The PVR

weighs 1,633 pounds and when deployed

measures 44 x 12 x 7 feet.

P3/P4 Facts in brief: 

Manufacturer:  Boeing

Dimensions:  45.3 feet long by 16 feet wide by

15.6 feet high

Weight:  34,885 lbs

Structure:  Primarily aluminum

Major components:  The P3 primary structure

is made of a hexagonal shaped aluminum

structure and includes four bulkheads and six
August 2006 PAYLOAD OVERVIEW 47

longerons.  The secondary structure includes

brackets, fittings, attach platforms, EVA

equipment and miscellaneous mechanisms.  

The P4 Photovoltaic module includes all

equipment outboard of the Solar Alpha Rotary

Joint (SARJ) outboard bulkhead, namely the

two Photovoltaic Array assemblies and the

Integrated Equipment Assembly (IEA).

Purpose:  To carry power, data and

environmental services along the integrated

truss structure.  Also, to provide active thermal

protection to electrical components throughout

the station and to allow the connection of

platforms to store spare parts.

Construction:  P3 was designed by the Boeing

design team at Huntington Beach, Calif.  Boeing

(now Pratt and Whitney) Rocketdyne Power

and Propulsion in Canoga Park, Calif. designed

P4.  Assembly of P3 and P4 in Tulsa, Okla.

started in 1997.  P3 was delivered to the Space

Station Processing Facility at Kennedy Space

Center on November 17, 1999 and P4 was

delivered on July 29, 2000.    P3 and P4 were

handed off to NASA on September 26, 2002.

Major Subcontractors:  Lockheed Martin,

Honeywell, Hamilton Sundstrand, Pratt and

Whitney Rocketdyne

Installation:  P3/P4 is to be installed during

STS‐115 to the P1 truss via the Segment to

Segment Attachment System (SSAS).

P3/P4 SRMS unberth and SSRMS handoff operations is shown above.
August 2006 PAYLOAD OVERVIEW 48

P3/P4 SSRMS maneuver and installation is shown above
August 2006 PAYLOAD OVERVIEW 49

Additional drawings to use 
August 2006 PAYLOAD OVERVIEW 50

EXTERNAL WIRELESS 

INSTRUMENTATION SYSTEM (EWIS) 

With the launch of the P4 element, NASA will

deploy a new External Wireless

Instrumentation System (EWIS) for the first

time.  The system consists of accelerometers

placed around the outboard truss elements of

the integrated truss structure.  The system is

currently installed on P4 and P5 and will also

be installed on the S4 and S6 truss structures

later this year at Kennedy Space Center, Fla.  

The six quarter‐size accelerometers on P4

measure accelerations in all directions, the x‐, y‐ 

and z‐ axis orientations.

The vibration data seen by the accelerometers

will be compared with their loads models so

they can be further refined by engineers with

actual on‐orbit data to better predict the

durability of the station’s integrated truss

structure.  NASA and Boeing engineers will use

these measurements, using refined models, to

analytically extend the 15‐year design life of the

truss elements.  EWIS will also give engineers a

better understanding of the actual response of

the truss system on orbit.

The space station has 33 hardwired

accelerometers currently installed on the

inboard truss elements (13 on S0, four on S1, six

on P1, five on S3 and five on P3).  The data from

these accelerometers is transferred through the

station Command and Data Handling System

where the data is relayed to the ground for

engineers to analyze.  The wireless system was

created because there was no easy way to

hardwire the accelerometers into the outboard

truss elements, outside of the Solar Alpha

Rotary Joint.

The wireless system uses a spread spectrum

radio system that transmits at 900 MHz, similar

to an older cordless phone.  International Space

Station power is supplied to the Remote Sensor

Unit (RSU) on the P4 truss element, which

contains the memory for the measurements,

computer controls and the radio transmitter.  

The data will be collected as soon as P4 is

powered up and will be stored in the RSU.  

When requested, it transmits the data to a

briefcase‐sized Network Control Unit (NCU),

which will be located in the avionics two rack

in Destiny.

In addition to the NCU, two eight‐inch‐long

antennas will be installed on the aft end cone

on the nadir side of Destiny.  The spacewalk to

install these antennas is considered a get‐ahead

task for this mission, but more than likely will

be installed on assembly STS‐118 in 2007.  The

NCU will be installed on station assembly

STS‐117.  EWIS is expected to be fully

operational upon completion of STS‐119.

The new system was designed and tested by

NASA and Boeing engineers, with work

beginning in May 2003.  It allows engineers to

gather real‐time data during dynamic events

that might cause higher vibrations (loads) on

the truss elements, such as docking operations

with the space shuttle or progress vehicles,

SARJ rotations or during thruster firings during

re‐boosts.  The data is expected to be retrieved

several times a month.

NASA and Boeing engineers believe these

measurements will allow some of the

conservatism and uncertainty to be removed

from their models.  Engineers also believe this

data will be helpful in better understanding the

kinds of loads exerted on future large structure

exploration vehicles.
August 2006 EXPERIMENTS 51

EXPERIMENTS 

EXPERIMENTAL PAYLOADS 

There are two experimental payloads that will

be carried to the station on STS‐115.  Currently,

the crew timeline does not support performing

the experiments.  Therefore the experiments

will be left for the station crew to perform.  

However, should an extra day be added to the

mission, the shuttle crew will perform the

experiments.

Effect of Spaceflight on Microbial Gene 

Expression & Virulence (MICROBE) 

The objective of the MICROBE payload is to

characterize the effects of space on the virulence

and physiology of three common microbes

previously isolated from the space shuttle

environment and to facilitate development of

effective antimicrobials/therapeutics and

effective systems for maintaining water

potability.

The experiment consists of 12 Group Activation

Packs (GAPs) that will be loaded on the shuttle

for sortie operations.  On orbit, the GAPs will

be manually activated using a small hand

crank, which will introduce fresh growth

medium into the cultures.  Following about 24

hours of passive ambient incubation, the crew

will manually terminate growth using the hand

crank.

Currently the timeline does not support

scheduling MICROBE.  If an extra day is added

to the mission then eight of the 12 GAPs may be

scheduled depending on the required activities

on that day.  The remaining four GAPs require

activation as late as possible in the mission.  

Scheduling the four GAPs is dependent on the

post undock timeline.

Yeast-GAP 

This experiment will study how individual

genes respond to microgravity conditions.  The

results could help scientists understand how

mammalian cells will respond when they are

grown in microgravity as well as improve

culturing techniques of mammalian cells on

Earth.

The experiment consists of four GAPs that will

be loaded on the shuttle for sortie operations.  

On orbit, the GAPs will be manually activated

using a small hand crank, which will introduce

fresh growth medium into the cultures.  

Following 24 hours of passive ambient

incubation, the crew will manually terminate

growth using the hand crank

Currently the timeline does not support

scheduling Yeast‐GAP.  It is to be transferred to

the station to be performed by the station crew.  

However, if an additional day is added to the

mission then the experiment may be scheduled.

DETAILED TEST OBJECTIVES 

Detailed Test Objectives (DTOs) are aimed at

testing, evaluating or documenting space

shuttle systems or hardware, or proposed

improvements to the space shuttle or space

station hardware, systems and operations.  

Station Detailed Test Objectives (SDTOs) are

aimed at testing, evaluating or documenting

space station systems or hardware or proposed

improvements to the station hardware, systems

and operations.
August 2006 EXPERIMENTS 52

Such experiments assigned to STS‐115 are listed

below.

DTO 805 Crosswind Landing 

Performance (If Opportunity) 

The purpose of this DTO is to demonstrate the

capability to perform a manually controlled

landing in the presence of a crosswind.  The

testing is done in two steps.

1. Prelaunch:  Prelaunch planning must allow

selection of a runway with Microwave

Scanning Beam Landing System (MSBLS)

support.  MSBLS is a set of dual

transmitters located beside the runway

providing precision navigation vertically,

horizontally and longitudinally with respect

to the runway.  This precision navigation

subsystem helps provide a higher

probability of a more precise landing with a

crosswind of 10 to 15 knots as late in the

flight as possible.

2. Entry:  This test requires that the crew

perform a manually controlled landing in

the presence of a 90‐degree crosswind

component of 10 to 15 knots steady state.

During a crosswind landing, the drag chute will

be deployed after nose gear touchdown when

the vehicle is stable and tracking the runway

centerline.

STATION DEVELOPMENT TEST 

OBJECTIVE (SDTO) 

SDTO 12004-U: Shuttle Booster Fan 

Bypass 

The purpose of this SDTO is to increase shuttle

on orbit cryogenic margin by bypassing and

deactivating the shuttle booster fan—airlock

fan—whenever possible on‐orbit.  The booster

fan will not be activated in post insertion and

will remain deactivated for most of the mission.  

It will however be activated for docking and

undocking to provide cooling to the powered

docking avionics and to prevent condensation

on the Orbiter Docking System hatch window.  

The fan will also be activated to provide good

airflow when the crew is working in the airlock,

such as during checkout of the spacesuits.  

While docked, sufficient air circulation will be

achieved through use of the lab forward

intermodular ventilation fan so that carbon

dioxide (CO2) levels on both the station and

shuttle can be controlled during the day by the

space station.  A real‐time call may instruct the

crew to replace lithium hydroxide on the

middeck to supplement the carbon dioxide

removal provided by the station.  Acceptable

flow rate through the IMV fan and proper CO2

levels on the shuttle throughout the docked

timeframe must be confirmed on‐orbit to prove

this cryogenic reduction method a success.

SDTO 13005-U ISS Structural Life 

Validation and Extension 

The overall purpose of this SDTO is to

guarantee safety of the station structure and

crew.  Specific objectives are to accurately

determine structural life usage of the P3/P4

truss, to expand station operations and to

increase the life of the structure.  This

reconstruction requires actual or educated

estimates of input (forcing function) and actual

output (on‐orbit sensor measurements) of the

station response.  Measurement of the force

input (i.e., thruster firing sequences, video of

crew activity, etc.) and station response will aid

reconstruction of station loads and structural

life usage over the life of the station, and thus

allowing life extension of the structure.  This

will be performed during P3/P4 installation

operations.
August 2006 EXPERIMENTS 53

SDTO 15003-U Microgravity 

Environment Definition 

The stationʹs microgravity instrumentation will

be used to measure the microgravity

environment and enable prediction of

acceptable microgravity experiment times for

similar station structural configurations.  The

measurements will be performed before

experimentation times to assure station

operations are consistent with the microgravity

requirements.  Measurements will be made for

specific expected attitudes and dynamic

disturbances—in the case of STS‐115 to

characterize the Solar Array Rotation Joint

checkout.

Short Duration Bioastronautics 

Investigation (SDBI) 

SDBI 1490B Bioavailability and 

Performance Effects of Promethazine 

(PMZ) During Spaceflight 

This investigates the effects of Promethazine, an

antihistamine that can be used as a medication

for motion sickness.  All participants will don

the Actiwatch activity monitor, a wristwatch‐

like device that tracks body motion, as soon as

possible on orbit.  They will wear the Actiwatch

and record sleep times throughout the mission.  

Before the first Promethazine (PMZ) dose,

participating crew members will collect a saliva

sample; thereafter, saliva samples will be

collected and a Stanford Sleepiness Score

completed at 1, 2, 4, 8, 24, 36 and 48 hours post‐

PMZ.  This protocol will be repeated each time

PMZ is taken.  Participants who do not take

PMZ will wear the Actiwatch and record sleep

times throughout the mission.  However, if a

participant elects to take PMZ before sleep,

saliva samples will be collected before the dose,

just before sleep, immediately upon wake‐up,

one hour after wake‐up, and at 24, 36, and 48

hours post‐dose.

SDBI 1493 Monitoring Latent Virus 

Reactivation and Shedding in 

Astronauts 

The objective of this SDBI is to determine the

frequency of induced reactivation of latent

viruses, latent virus shedding, and clinical

disease after exposure to the physical,

physiological, and psychological stressors

associated with spaceflight.

Induced alterations in the immune response

will become increasingly important on long‐

duration missions, with one focus being the

potential for reactivation and dissemination or

shedding of latent viruses.  An example of a

latent virus is herpes simplex type‐1, which

infects 70 to 80 percent of adults.  Its

manifestation is classically associated with the

presence of cold sores, pharyngitis, and

tonsillitis.  Herpes simplex type‐1 is usually

acquired through contact with the saliva, skin,

or mucous membranes of an infected

individual.  However, many recurrences are

asymptomatic, resulting in shedding of the

virus.

SDBI 1634 Sleep-Wake Actigraphy and 

Light Exposure during Spaceflight 

Participating crew members will don the

Actiwatch activity monitor as soon as possible

upon entering orbit and will wear it

continuously throughout the mission on their

non‐dominant wrists outside of their

clothing/sleeve.  The Actiwatch watch can be

temporarily removed for activities such as

spacewalks.  Subjects will also complete a short

log within 15 minutes of final awakening every

morning in‐flight.
August 2006 SHUTTLE REFERENCE DATA 54

SHUTTLE REFERENCE DATA 

SHUTTLE ABORT MODES 

RSLS Aborts 

These occur when the on‐board shuttle

computers detect a problem and command a

halt in the launch sequence after taking over

from the ground launch sequencer and before

solid rocket booster ignition.

Ascent Aborts 

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 engine.  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).

Return to Launch Site 

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 liftoff.

The RTLS profile is designed to accommodate

the loss of thrust from one space shuttle main

engine between liftoff 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 external

tank 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

three‐engine RTLS is selected at the last

moment, about 3 minutes, 34 seconds into the

mission; whereas an RTLS chosen due to an

engine out at liftoff is selected at the earliest

time, about 2 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

toward 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
August 2006 SHUTTLE REFERENCE DATA 55

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 maneuver

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 maneuver has

been completed, the glide phase of the RTLS

begins.  From then on, the RTLS is handled

similarly to a normal entry.

Transoceanic Abort Landing 

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 about 45 minutes after launch.  The

landing site is selected near the normal ascent

ground track of the orbiter 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.  The three

landing sites that have been identified for a

launch are Zaragoza, Spain; Moron, Spain; and

Istres, France.

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 vehicleʹs

landing weight.  TAL is handled like a normal

entry.

Abort to Orbit 

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
August 2006 SHUTTLE REFERENCE DATA 56

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 

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 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 about

90 minutes after liftoff.

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 Aborts 

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 inflight

crew escape system would be used before

ditching the orbiter.

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.

Mission Control Houston is prime for calling

these aborts because it has a more precise

knowledge of the orbiterʹs position than the

crew can obtain from on‐board 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 on‐

board methods, such as cue cards, dedicated
August 2006 SHUTTLE REFERENCE DATA 57

displays and display information, to determine

the 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 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 about

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.

SHUTTLE ABORT HISTORY 

RSLS Abort History: 

(STS-41 D) June 26, 1984 

The countdown for the second launch attempt

for Discovery’s maiden flight ended at T‐4

seconds when the orbiter’s computers detected

a sluggish valve in main engine No. 3.The main

engine was replaced and Discovery was finally

launched on Aug. 30, 1984.

(STS-51 F) July 12, 1985 

The countdown for Challenger’s launch was

halted at T‐3 seconds when on‐board

computers detected a problem with a coolant

valve on main engine No. 2.  The valve was

replaced and Challenger was launched on

July 29, 1985.

(STS-55) March 22, 1993 

The countdown for Columbia’s launch was

halted by on‐board computers at T‐3 seconds

following a problem with purge pressure

readings in the oxidizer preburner on main

engine No. 2 Columbia’s three main engines

were replaced on the launch pad, and the flight

was rescheduled behind Discovery’s launch on

STS‐56.  Columbia finally launched on

April 26, 1993.

(STS-51) Aug. 12, 1993 

The countdown for Discovery’s third launch

attempt ended at the T‐3 second mark when

onboard computers detected the failure of one

of four sensors in main engine No. 2 which

monitor the flow of hydrogen fuel to the

engine.  All of Discovery’s main engines were

ordered replaced on the launch pad, delaying

the shuttle’s fourth launch attempt until

Sept. 12, 1993.

(STS-68) Aug. 18, 1994 

The countdown for Endeavour’s first launch

attempt ended 1.9 seconds before liftoff when

on‐board computers detected higher than

acceptable readings in one channel of a sensor

monitoring the discharge temperature of the

high pressure oxidizer turbopump in main

engine No. 3.  A test firing of the engine at the

Stennis Space Center in Mississippi on

September 2nd confirmed that a slight drift in a

fuel flow meter in the engine caused a slight

increase in the turbopump’s temperature.  The

test firing also confirmed a slightly slower start

for main engine No. 3 during the pad abort,

which could have contributed to the higher

temperatures.  After Endeavour was brought

back to the Vehicle Assembly Building to be

outfitted with three replacement engines,
August 2006 SHUTTLE REFERENCE DATA 58

NASA managers set Oct. 2 as the date for

Endeavour’s second launch attempt.

Abort to Orbit History: 

(STS-51 F) July 29, 1985 

After an RSLS abort on July 12, 1985,

Challenger was launched on July 29, 1985.  Five

minutes and 45 seconds after launch, a sensor

problem resulted in the shutdown of center

engine No. 1, resulting in a safe ʺabort to orbitʺ 

and successful completion of the mission.

SPACE SHUTTLE MAIN ENGINES 

Developed in the 1970s by NASAʹs Marshall

Space Flight Center in Huntsville, Ala., the

space shuttle main engine is the most advanced

liquid‐fueled rocket engine ever built.  Every

space shuttle main engine is tested and proven

flight‐worthy at NASAʹs Stennis Space Center

in south Mississippi, before installation on an

orbiter.  Its main features include variable

thrust, high performance reusability, high

redundancy and a fully integrated engine

controller.

The shuttle’s three main engines are mounted

on the orbiter aft fuselage in a triangular

pattern.  Spaced so that they are movable

during launch, the engines are used – in

conjunction with the solid rocket boosters – to

steer the shuttle vehicle.

Each of these powerful main engines is 14 feet

(4.2 meters) long, weighs about 7,000 pounds

(3,150 kilograms) and is 7.5 feet (2.25 meters) in

diameter at the end of its nozzle.

The engines operate for about 81⁄2  minutes

during liftoff and ascent—burning more than

500,000 gallons (1.9 million liters) of super‐cold

liquid hydrogen and liquid oxygen propellants

stored in the huge external tank attached to the

underside of the shuttle.  The engines shut

down just before the shuttle, traveling at about

17,000 mph (28,000 kilometers per hour),

reaches orbit.

The main engine operates at greater

temperature extremes than any mechanical

system in common use today.  The fuel,

liquefied hydrogen at ‐423 degrees Fahrenheit  

(‐253 degrees Celsius), is the second coldest

liquid on Earth.  When it and the liquid oxygen

are combusted, the temperature in the main

combustion chamber is 6,000 degrees

Fahrenheit (3,316 degrees Celsius), hotter than

the boiling point of iron.

The main engines use a staged combustion

cycle so that all propellants entering the engines

are used to produce thrust or power—more

efficiently than any previous rocket engine.  In

a staged combustion cycle, propellants are first

burned partially at high pressure and relatively

low temperature—then burned completely at

high temperature and pressure in the main

combustion chamber.  The rapid mixing of the

propellants under these conditions is so

complete that 99 percent of the fuel is burned.

At normal operating level, the engines generate

490,847 pounds of thrust (measured in a

vacuum).  Full power is 512,900 pounds of

thrust; minimum power is 316,100 pounds of

thrust.

The engine can be throttled by varying the

output of the pre‐burners, thus varying the

speed of the high‐pressure turbopumps and,

therefore, the flow of the propellant.

At about 26 seconds into launch, the main

engines are throttled down to 316,000 pounds

of thrust to keep the dynamic pressure on the

vehicle below a specified level—about

580 pounds per square foot or max q.  Then, the
August 2006 SHUTTLE REFERENCE DATA 59

engines are throttled back up to normal

operating level at about 60 seconds.  This

reduces stress on the vehicle.

The main engines are throttled down again at

about seven minutes, 40 seconds into the

mission to maintain three g’s – three times the

Earth’s gravitational pull – again reducing

stress on the crew and the vehicle.  This

acceleration level is about one‐third the

acceleration experienced on previous crewed

space vehicles.

About 10 seconds before main engine cutoff or

MECO, the cutoff sequence begins; about three

seconds later the main engines are commanded

to begin throttling at 10 percent thrust per

second to 65 percent thrust.  This is held for

about 6.7 seconds, and the engines are shut

down.

The engine performance has the highest thrust

for its weight of any engine yet developed.  In

fact, one space shuttle main engine generates

sufficient thrust to maintain the flight of 21⁄2

747 airplanes.

The space shuttle main engine is also the first

rocket engine to use a built‐in electronic digital

controller, or computer.  The controller will

accept commands from the orbiter for engine

start, change in throttle, shutdown, and

monitor engine operation.  In the event of a

failure, the controller automatically corrects the

problem or safely shuts down the engine.

NASA continues to increase the reliability and

safety of shuttle flights through a series of

enhancements to the space shuttle main

engines.  The engines were modified in 1988,

1995, 1998 and 2001.  Modifications include new

high‐pressure fuel and oxidizer turbopumps

that reduce maintenance and operating costs of

the engine, a two‐duct powerhead that reduces

pressure and turbulence in the engine, and a

single‐coil heat exchanger that lowers the

number of post flight inspections required.  

Another modification incorporates a large‐

throat main combustion chamber that improves

the engine’s reliability by reducing pressure

and temperature in the chamber.

After the orbiter lands, the engines are removed

and returned to a processing facility at

Kennedy Space Center, Fla., where they are

rechecked and readied for the next flight.  Some

components are returned to the main engine’s

prime contractor, Pratt & Whitney RocketDyne,

West Palm Beach, Fla., for regular maintenance.  

The main engines are designed to operate for

7.5 accumulated hours.

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 about

3,300,000 pounds at launch.  They are ignited

after the three space shuttle main enginesʹ 

thrust level is verified.  The two SRBs provide

71.4 percent of the thrust at liftoff and during

first‐stage ascent.  Seventy‐five seconds after

SRB separation, SRB apogee occurs at an

altitude of about 220,000 feet, or 35 nautical

miles (40 statute miles).  SRB impact occurs in

the ocean about 122 nautical miles (140 statute

miles) downrange.

The SRBs are the largest solid‐propellant

motors ever flown and the first designed for
August 2006 SHUTTLE REFERENCE DATA 60

reuse.  Each is 149.16 feet long and 12.17 feet in

diameter.

Each SRB weighs about 1,300,000 pounds at

launch.  The propellant for each solid rocket

motor weighs about 1,100,000 pounds.  The

inert weight of each SRB is about

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 liftoff.

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.

Additionally, special structural tests were done

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 about

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

about a third 50 seconds after liftoff 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 heavy‐duty rail car with a

specially built cover.

The nozzle expansion ratio of each booster

beginning with the STS‐8 mission is 7‐to‐79.  
August 2006 SHUTTLE REFERENCE DATA 61

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.

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, a transition

piece between the nose cone and solid rocket

motor, 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 ATK 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
August 2006 SHUTTLE REFERENCE DATA 62

only the top nut is frangible.  The top nut

contains two NASA standard detonators

(NSDs), 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 3.5 inches in

diameter.  The frangible nut is captured in a

blast container.

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 signals—arm, 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 on‐board computers at T minus 6.6 seconds

(staggered start—engine three, engine two,

engine one—all about 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 liftoff 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
August 2006 SHUTTLE REFERENCE DATA 63

bending load modes are allowed to initialize

(movement of 25.5 inches measured at the tip of

the external tank, with movement towards the

external tank).

At T minus zero, the two SRBs are ignited

under command of the four on‐board

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 on‐board 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 corresponding 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.

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 generatorʹs 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 lineʹs.  Then all the

fuel is supplied to the fuel pump.
August 2006 SHUTTLE REFERENCE DATA 64

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 servoactuator 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.

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 liftoff 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
August 2006 SHUTTLE REFERENCE DATA 65

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.

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
August 2006 SHUTTLE REFERENCE DATA 66

the redundant NSD pressure cartridge in each

bolt and ignite the BSMs to effect a clean

separation.

SPACE SHUTTLE SUPER LIGHT WEIGHT 

TANK (SLWT) 

The super lightweight external tank (SLWT)

made its first shuttle flight June 2, 1998, on

mission STS‐91.  The SLWT is 7,500 pounds

lighter than the standard external tank.  The

lighter weight tank allows 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 shuttleʹs current tank.  

The tankʹs structural design has also been

improved, making it 30 percent stronger and

5 percent less dense.

The SLWT, like the standard tank, is

manufactured at Michoud Assembly, near

New Orleans, 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 shuttleʹs three

main engines.
August 2006 ACRONYMS/ABBREVIATIONS 67

ACRONYMS AND ABBREVIATIONS 

A/L Airlock

AA Antenna Assembly

AAA Avionics Air Assembly

ABC Audio Bus Coupler

AC Assembly Complete

ACBM Active Common Berthing Mechanism

ACO Assembly and Checkout Officer

ACS Atmosphere Control and Supply

ACSM Attitude Control System Moding

ACU Arm Computer Unit

ADO Adaptation Data Overlay

ADSEP Advanced Separation

ADVASC Advanced Astroculture

ADVASC‐GC Advanced Astroculture—Growth Chamber

AEA Antenna Electronics Assembly

AFD Aft Flight Deck

AJIS Alpha Joint Interface Structure

AKA Active Keel Assembly

APAS Androgynous Peripheral Attachment System

APCU Assembly Power Converter Unit

APDS Androgynous Peripheral Docking System

APFR Articulating Portable Foot Restraint

APM Attached Pressurized Module

APPCM Arm Pitch Plane Change Mode

APS Automated Payload Switch

AR Atmosphere Revitalization

ARCU American‐to‐Russian Converter Unit

ARIS Active Rack Isolation System

ARS Air Revitalization System

ASCR Assured Safe Crew Return

ATA Ammonia Tank Assembly

ATCS Active Thermal Control System

ATU Audio Terminal Unit

AUAI Assemble Contingency System/UHF Audio Interface

AVU Artificial Vision Unit

AVV Accumulator Vent Valve

BA Bearing Assembly

BBC Bus Bolt Controller
August 2006 ACRONYMS/ABBREVIATIONS 68

BC Bus Controller

BCDU Battery Charge/Discharge Unit

BCU Backup Controller Unit  

BDU Backup Drive Unit

BG Beta Gimbal

BGA Beta Gimbal Assembly

BGDTS Beta Gimbal Deployment Transition Structure

BGHS Beta Gimbal Housing Subassembly

BIT Built‐In Test

BITE Built‐In Test Equipment

BMRRM Bearing Motor and Roll Ring Module

BONEMAC Bone Marrow Macroph

ages in Space

BRS Bottom Right Side

BSP Baseband Signal Processor

BTS Bolt Tight Switch

C&C Command and Control

C&DH Command and Data Handling

C&M Control and Monitor

C&T Communication and Tracking

C&W Caution and Warning

C/A‐code Coarse/Acquisition‐code

C/L Crew Lock

CA Control Attitude

CAS Common Attach System

CBM Common Berthing Mechanism

CBOSS Cellular Biotechnology Operating Science System

CCAA Common Cabin Air Assembly

CCASE Commercial Cassette Experiment

CCD Cursor Control Device

CCMS Concentric Cable Management System

CCS Communication and Control System

CCTV Closed‐Circuit Television

CDDT Common Display Development Team

CDRA Carbon Dioxide Removal Assembly

CDS Command and Data Software

CETA Crew and Equipment Translation Aid

CEU Control Electronics Unit

CFA Circular Fan Assembly

CGBA Commercial Generic Bioprocessing Apparatus

CHeCS Crew Health Care System

CHX Condensing Heat Exchanger
August 2006 ACRONYMS/ABBREVIATIONS 69

CID Circuit Interrupt Device

CIOB Cargo Integration and Operations Branch

CLA Camera and Light Assembly

CLPA Camera Light and Pan/Tilt Assembly

CMG Control Moment Gyroscope

CMG‐TA Control Moment Gyroscope‐Thruster Assist

CO2 Carbon Dioxide

COAS Crew Optical Alignment Sight

COR Communication Outage Recorder

COTS Commercial‐Off‐The‐Shelf

CP Cold Plate

CPCG‐H Commercial Protein Crystal Growth‐High

CR Change Request

CRES Corrosion Resistant Steel

CRIM Commercial Refrigerator Incubator Module

CRIM‐M Commercial Refrigerator Incubator Module‐Modified

CRPCM Canadian Remote Power Controller Module

CSA Computer Systems Architecture

CSA‐CP Compound Specific Analyzer‐Combustion Products

CSCI Computer Software Configuration Item

CSM Cargo Systems Manual

CTB Cargo Transfer Bag

CVIU Common Video Interface Unit

CVT Current Value Table

CVV Carbon Dioxide Vent Valve

CWC Contingency Water Collection

DAIU Docked Audio Interface Unit

DAP Digital Autopilot

DC Docking Compartment

dc direct current

DCP Display and Control Panel

DCSU Direct Current Switching Unit

DDCU DC‐to‐DC Converter Unit

DDCU‐CP DC‐to‐DC Converter Unit‐Cold Plate

DDCU‐E External DDCU

DDCU‐HP DC‐to‐DC Converter Unit‐Heat Pipe

DDCU‐I Internal DDCU

DFL Data Format Load

DLA Drive Locking Assembly

DMCU Docking Mechanism Control Unit

DMS‐R Data Management System‐Russian
August 2006 ACRONYMS/ABBREVIATIONS 70

dp/dt delta pressure/delta time

DPA Digital Preassembly

DPS Data Processing System

DTO Development Test Objective

DTV Digital Television

E/L Equipment Lock

E‐Stop Emergency Stop

EACP EMU Audio Control Panel

EAIU EMU Audio Interface Unit

EAS Early Ammonia Servicer

EATCS External Active Thermal Control Subsystem

ECLSS Environmental Control and Life Support System

ECU Electronics Control Unit

ED Engagement Drive

EDDA External Maneuvering Unit Don/Doff Assembly

EEATCS Early External Active Thermal Control System

EET Experiment Elapsed Time

EETCS Early External Thermal Control System

EFGF Electrical Flight‐releasable Grapple Fixture

EGIL Electrical Generation and Integrated Lighting Systems Engineer

EIA Electrical Interface Assembly

EMPEV Emergency Manual Pressure Equalization Value

EMU Extravehicular Mobility Unit

EOA EVA Ohmmeter Assembly

EPCE Electrical Power Consuming Equipment

EPG Electrical Power Generator

EPS Electrical Power System

ER Edge Router

ESA External Sampling Adapter

ESP External Stowage Platform

ESSMDM Enhanced Space Station Multiplexer/Demultiplexer

ESU End Stop Unit

ETCS External Thermal Control Ssytem

ETI Elapsed Time Indicator

ETRS EVA Temporary Rail Stop

ETSD EVA Tool Storage Device

ETVCG External Television Cameras Group

EUE Experiment Unique Equipment

EV Extravehicular

EV‐CPDS Extravehicular‐Charged Particle Directional Spectrometer

EVA Extravehicular Activity
August 2006 ACRONYMS/ABBREVIATIONS 71

EVR Extravehicular Robotics

EVSU External Video Switching Unit

EXPRESS EXpedite the PRocessing of Experiments to the Space Station

EXT Experimental Terminal

EWIS External Wireless Instrumentation System

FAWG Flight Assignment Working Group

FC Firmware Controller

FCC Flat Controller Circuit

FCT Flight Control Team

FCV Flow Control Valve

FD Flight Day

FDA Fault Detection Annunciation

FDIR Failure, Detection, Isolation and Recovery

FDS Fire Detection and Suppression

FET Field Effect Transistor

FGB Functional Cargo Block (Zarya Module of ISS)

FHRC Flex Hose Rotary Coupler

FI Fault Isolator

FPU Fluid Pumping Unit

FQDC Fluid Quick Disconnect Coupling

FRD Flight Requirements Document

FRGF Flight Releasable Grapple Fixture

FSE Flight Support Equipment

FSS Fluid System Servicer

FWCI Firmware Configuration Item

GAS Get Away Special

GC Growth Cell

GCA Growth Cell Assembly

GFE Government‐Furnished Equipment

GFI Ground Fault Interrupter

GLONASS GLObal Navigational Satellite System

GN&C Guidance, Navigation and Control

GNC Guidance Navigation Computer

GPC General Purpose Computer

GPRV Gas Pressure regulating Valve

GPS Global Positioning System

GUI Graphical User Interface

H2 Hydrogen

HAB Habitat Module

HC Hand Controller
August 2006 ACRONYMS/ABBREVIATIONS 72

HCA Hollow Cathode Assembly

HCOR High‐Rate Communication Outage Recorder

HDR High Data Rate

HDRL High Date Rate Link

HEPA High Efficiency Particulate Air

HGA High Gain Antenna

HHL Handheld Lidar

HP Heat Pipe

HPGT High Pressure Gas Tank

HRF Human Research Facility

HRF‐PUF‐DK Human Research Facility Puff Data Kit

HRF‐Res Human Research Facility Resupply

HRFM High Rate Frame Multiplexer

HRM High Rate Modem

HRS Hand Reaction Switch

I/F Interface

I/O Input/Output

IAC Internal Audio Controller

IAS Internal Audio Subsystem

IATCS Internal Active Thermal Control System

ICC Integrated Cargo Carrier

IDA Integrated Diode Assembly

IDRD Increment Definition Requirements Document

IEA Integrated Equipment Assembly

IFHX Interface Heat Exchanger

IFM In‐flight Maintenance

IMCA Integrated Motor Control Assembly

IMCS Integrated Mission Control System

IMU Impedance Matching Unit

IMV Intermodule Ventilation  

INCO Instrumentation and Communication Officer

INSTM Instrumentation

INT Internal

INTSYS Internal Systems

IOC Input/Output Controller

IOCU Input/Output Controller Unit

IP International Partner

IRU In‐Flight Refill Unit

ISA Internal Sampling Adapter

ISL Integrated Station LAN

ISO Inventory and Stowage Officer
August 2006 ACRONYMS/ABBREVIATIONS 73

ISPR International Standard Payload Rack

ISS International Space Station

ISSSH International Space Station Systems Handbook

IT Integrated Truss

ITCS Internal Thermal Control System

ITS Integrated Truss Segment

IUA Interface Umbilical Assembly

IV Intravehicular  

IVA Intravehicular Activity

IVSU Internal Video Switch Unit

IWIS Internal Wireless Instrumentation System

JEM Japanese Experiment Module

JEU Joint Electronic Unit

kW Kilowatt

LA Launch Aft

Lab Laboratory  

LAN Local Area Network

LB Local Bus

LB‐RWS RWS Local Bus

LCA Lab Cradle Assembly

LCC Launch Commit Criteria

LCD Liquid Crystal Display

LDI Local Data Interface

LDR Low Data Rate

LDU Linear Drive Unit

LED Light‐Emitting Diode

LEE Latching End Effector

LEU LEE Electronic Unit

LFDP Load Fault Detection Protection

LGA Low Gain Antenna

LLA Low Level Analog

LMC Lightweight Multipurpose Carrier

LON Launch On Need

LT Low Temperature

LTA Launch to Activation

LTL Low Temperature Loop

LTU Load Transfer Unit

LVLH Local Vertical Local Horizontal

MA Mechanical Assembly
August 2006 ACRONYMS/ABBREVIATIONS 74

MAM Manual Augmented Role

MBE Metal Bellows Expander

MBM Manual Berthing Mechanism

MBS Mobile Remote Service Base System

MBSU Main Bus Switching Unit

MC Midcourse Correction

MCA Major Constituent Analyzer

MCAS MBS Common Attach System

MCC Mission Control Center

MCC‐H Mission Control Center‐Houston

MCC‐M Mission Control Center‐Moscow

MCDS Multifunction CRT Display System

MCS Motion Control System

MCU MBS Computer Unit

MDA Motor Drive Assembly

MDL Middeck Locker

MDM Multiplexer/Demultiplexer

MED OPS Medical Operations

MEPS Microencapsulation Electrostatic Processing System

MEPSI Micro‐Electromechanical System‐based Pico Satellite Inspector

MER Mission Evaluation Room

MET Mission Elapsed Time

METOX Metal Oxide

MFCV Manual Flow Control Valve

MHS MCU Host Software

MILA Mode Indicating Light Assembly

MIP Mission Integration Plan

MISSE Materials International Space Station Experiment

MLI Multi‐Layer Insulation

MM/OD Micrometeoroid/Orbital Debris

MMT Mission Management Team

MOD Mission Operations Directorate

MPEV Manual Pressure Equalization Valve

MPLM Multipurpose Logistics Module

MPM Manipulator Positioning Mechanism

MRL Manipulator Retention Latch

MRS Mobile Remote Servicer

MSD Mass Storage Device

MSFC Marshall Space Flight Center

MSG Microgravity Science Glovebox

MSS Mobile Servicing System

MT Mobile Transporter
August 2006 ACRONYMS/ABBREVIATIONS 75

MTCL Mobile Transporter Capture Latch

MTL Moderate Temperature Loop

MTS Module‐to‐Truss Segment

MTSAS Module‐to‐Truss Segment Attachment System

MTWsN Move to Worksite Number

N

2 Nitrogen

n.mi. nautical mile

NASA National Aeronautics and Space Administration

NCC Nominal Corrective Combination burn

NCG Non Condensable Gas

NCS Node Control Software

NCU Network Control Unit

NET No Earlier Than

NIA Nitrogen Interface Assembly

NiH

2 Nickel Hydrogen

NIV Nitrogen Introduction Valve

NSI NASA Standard Initiator

NSTS National Space Transportation System

NTA Nitrogen Tank Assembly O2 Oxygen

OCA Orbital Communications Adapter

OCAD Operational Control Agreement Document

OCJM Operator‐Commanded Joint Position Mode

OCPM Operator‐Commanded POR Mode

OCS Operations and Control Software

ODIN Orbital Design Integration System

ODS Orbiter Docking System

OI Operational Increment

OIU Orbiter Interface Unit

OIV Oxygen Isolation Valve

OMI On‐OrbitMaintainable It

em

OMS Orbital Maneuvering System

OPCGA Observable Protein Crystal Growth Apparatus

OPP OSVS

Ops Operations

OPS LAN Operations Local Area Network

ORBT Optimized RBar Targeting Technique

ORCA Oxygen Recharge Compressor Assembly

ORU Orbital Replacement Unit

OSE Orbiter Support Equipment

OSO Operations Support Officer
August 2006 ACRONYMS/ABBREVIATIONS 76

OSVS Orbiter Space Vision System

OTD ORU Transfer Device

OV Orbiter Vehicle

P Port

P&S Pointing and Support

P‐code Precision Code

P/L Payload

P/TV Photo/Television  

P3/P4 Port 3/Port 4

PAS Payload Attach System

PBA Portable Breathing Apparatus

PC Personal Computer

PCA Pressure Control Assembly

PCAM Protein Crystallization Apparatus for Microgravity

PCBM Passive Common Berthing Mechanism

PCC Power Converter Controller

PCG‐STES Protein Crystal Growth‐Single Thermal Enclosure System

PCMCIA Personal Computer Memory Card International Adapter

PCP Pressure Control Panel

PCR Portable Computer Receptacle  

PCS Portable Computer System

PCT Post‐Contact Thrusting

PCU Plasma Connector Unit

PCVP Pump and Control Valve Package

PDGF Power and Data Grapple Fixture

PDI Payload Data Interface

PDIP Payload Data Interface Panel

PDRS Payload Deployment and Retrieval System

PDTA Power Data Transfer Assembly

PDU Power Drive Unit

PEHG Payload Ethernet Hub Gateway

PFCS Pump Flow Control Subassembly

PFE Portable Fire Extinguisher

PFMC Pump/Fan Motor Controller

PGBA‐S Plant Generic Bioprocessing Apparatus‐Stowage

PGSC Portable General Support Computer

PGT Pistol Grip Tool

PHALCON Power, Heating, Articulation, Lighting, and Control Officer

PJAM Pre‐stored Joint Position Autosequence Mode

PLB Payload Bay

PM Pump Module
August 2006 ACRONYMS/ABBREVIATIONS 77

PMA Pressurized Mating Adapter

PMCU Power Management Control Unit

PMDIS Perceptual Motor Deficits In Space

PMP Payload Mounting Panel

POA Payload/ORU Accommodation

POR Point of Reference

POST Power ON Self‐Test

PP Planning Period

PPA Pump Package Assembly

PPAM Pre‐stored POR Autosequence Mode

ppO

2 partial pressure of oxygen

PPRV Positive Pressure Relief Valve

PPT Precipitate  

PRD Payload Retention Device

PRLA Payload Retention Latch Assembly

Prox‐Ops Proximity Operations

PSN Power Source Node

PSP Payload Signal Processor  

PTB Payload Training Buffer

PTCS Passive Thermal Control System

PTR Port Thermal Radiator

PTU Pan/Tilt Unit

PV Photovoltaic

PVCA Photovoltaic Controller Application

PVCE Photovoltaic Controller Element

PVCU Photovoltaic Controller Unit

PVM Photovoltaic Module

PVR Photovoltaic Radiator

PVRGF Photovoltaic Radiator Grapple Fixture

PVTCS Photovoltaic Thermal Control System

PWP Portable Work Platform

PWR Portable Water Reservoir

PYR Pitch Yaw Roll

QD Quick Disconnect

R/F Refrigerator/Freezer

RACU Russian‐to‐American Converter Unit

RAIU Russian Audio Interface Unit

RAM Random Access Memory

RAMV Rheostat Air Mix Valve

RBB Right Blanket Box

RBI Remote Bus Isolator
August 2006 ACRONYMS/ABBREVIATIONS 78

RBVM Radiator Beam Valve Module

RCC Reinforced Carbon‐Carbon

RCS Reaction Control System

RDA Retainer Door Assembly

RF Radio Frequency

RFCA Rack Flow Control Assembly

RFG Radio Frequency Group

RGA Rate Gyro Assemblies

RHC Rotational Hand Controller

RHX Regenerative Heat Exchanger

RIC Rack Interface Controller

RJMC Rotary Joint Motor Controller

RMS Remote Manipulator System

ROS Russian Orbital Segment

RPC Remote Power Controller

RPCM Remote Power Controller Module

RPDA Remote Power Distribution Assembly

RPM Rbar Pitch Maneuver

RPOP Rendezvous and Proximity Operations Program

RS Russian Segment

RSC RMS Sideview Camera

RSP Resupply Stowage Platform

RSR Resupply Stowage Rack

RSTS Rack Standalone Temperature Sensor

RSU Roller Suspension Unit

Remote Sensing Unit

RT Remote Terminal

RT‐Box Reaction Time Box

RTAS Rocketdyne Truss Attachment System

RTD Resistive Thermal Device

RTL Ready to Latch

RWS Robotic Workstation

S Starboard

S&M Structures and Mechanisms

SA SolarArray

SABB SolarArray Blanket Box

SAGE Space Arabidoposis Genomics Experiment

SARJ Solar Alpha Rotary Joint

SARJ_C SARJ Controller

SARJ_M SARJ Manager

SASAS‐band Antenna Support Assembly
August 2006 ACRONYMS/ABBREVIATIONS 79

SAW Solar Array Wing

SCA Switchgear Controller Assembly

SCI Signal Conditioning Interface

SCU Service and Cooling Umbilical

SD Smoke Detector

SDO Solenoid Driver Output

SDS Sample Delivery System

SEM Shunt Electronics Module

SEPS Secondary Electrical Power Subsystem

SFCA System Flow Control Assembly

SFU Squib Firing Unit

SGANT Space‐to‐Ground Antenna

SHOSS Spacehab Oceaneering Space System

SHOT Space Hardware Optimization Technology

SIGI Space Integrated Global Positioning System/Inertial Navigation System

SJRM Single Joint Rate Mode

SLDP Spacelab Data Processing

SLP Spacelab Logistics Pallet

SM Service Module

SMCC Shuttle Mission Control Center

SMDP Service Module Debris Panel

SOC State of Charge

SOV Shutoff Valve

SPCE Servicing Performance and Checkout Equipment

SPD Spool Positioning Device

SPDA Secondary Power Distribution Assembly

SPDM Special Purpose Dexterous Manipulator

SPG Single‐Point Ground

SRMS Shuttle Remote Manipulator System

SSAS Segment‐to‐Segment Attach System

SSBA Space Station Buffer Amplifier

SSC Station Support Computer

SSMDM Space Station Multiplexer/Demultiplexer

SSOR Space‐to‐Space Orbiter Ratio

SSP Standard Switch Panel

SSRMS Space Station Remote Manipulator System

SSSH Space Shuttle Systems Handbook

SSSR Space‐to‐Space Station Radio

SSU Sequential Shunt Unit

STCR Starboard Thermal Control Radiator

STES Single Thermal Enclosure System

STR Starboard Thermal Radiator
August 2006 ACRONYMS/ABBREVIATIONS 80

SVS Space Vision System

TA Thruster Assist

TAA Triaxial Accelerometer Assembly

TAH Tray Actuation Handle

TBA Trundle Bearing Assembly

TC Terminal Computer

TCCS Trace Contaminant Control Subassembly

TCCV Temperature Control and Check Valve

TCS Trajectory Control Sensor

TD Translation Drive

TDRS Tracking and Data Relay Satellite

TDRSS Tracking and Data Relay Satellite System

TEA Torque Equilibrium Attitude

TFR Translation Foot Restraint

THC Temperature and Humidity Control

THOR Thermal Operations and Resources Officer

TI Terminal Phase Initiation

TORF Twice Orbital Rate Flyaround

TORU Teleoperator Control Mode

TORVA Twice Orbital Rate +Rbar to +Vbar Approach

TPL Transfer Priority List

TRAC Test of Reaction and Adaption Capabilities

TRC Transmitter Receiver Controller

TRRJ Thermal Radiator Rotary Joint

TSP Twisted Shielded Pair

TTCR Trailing Thermal Control Radiator

TUS Trailing Umbilical System

TVIS Treadmill Vibration Isolation System

TWMV Three‐Way Mixing Valve

UCCAS Unpressurized Cargo Carrier Attach System

UDG User Data Generation

UF Utilization Flight

UHF Ultrahigh Frequency

UIA Umbilical Interface Assembly

UIP Utility Interface Panel

ULF Utilization Logistics Flight  

UMA Umbilical Mechanism Assembly

UOP Utility Outlet Panel

USL U.S. Laboratory

USOS United States On‐Orbit Segment
August 2006 ACRONYMS/ABBREVIATIONS 81

UTA Utility Transfer Assembly

VAJ Vacuum Access Jumper

VBSP Video Baseband Signal Processor

VCP Video Camera Port

VDS Video Distribution System

VDU Video Distribution Unit

VES Vacuum Exhaust System

VGS Video Graphics Software

VRCV Vent/Relief Control Valve

VRIV Vent/Relief Isolation Valve

VRS VES Resource System

VRV Vent/Relief Valve

VSC Video Signal Converter

VSSA Video Stanchion Support Assembly

W/S Worksite

WETA WVS External Transceiver Assembly

WHS Workstation Host Software

WIF Worksite Interface

WRM Water Recovery Management

WS Water Separator

WVA Water Vent Assembly

XPOP X‐axis Pointing Out of Plane

ZCG‐SS Zeolite Crystal Growth—Sample Stowage

ZSR Zero‐g Stowage Rack
August 2006 MEDIA ASSISTANCE 82

MEDIA ASSISTANCE 

NASA TELEVISION TRANSMISSION 

NASA Television is carried on an MPEG‐2

digital signal accessed via satellite AMC‐6, at 72

degrees west longitude, transponder 17C, 4040

MHz, vertical polarization.  For those in Alaska

or Hawaii, NASA Television will be seen on

AMC‐7, at 137 degrees west longitude,

transponder 18C, at 4060 MHz, horizontal

polarization.  In both instances, a Digital Video

Broadcast (DVB)‐compliant Integrated Receiver

Decoder (IRD) (with modulation of QPSK/DBV,

data rate of 36.86 and FEC 3⁄4) will be needed for

reception.  The NASA Television schedule and

links to streaming video are available at:

http://www.nasa.gov/ntv 

NASA TVʹs digital conversion will require

members of the broadcast media to upgrade

with an ʹaddressableʹ Integrated Receiver

Decoder, or IRD, to participate in live news

events and interviews, press briefings and

receive NASAʹs Video File news feeds on a

dedicated Media Services channel.  NASA

mission coverage will air on a digital NASA

Public Services (ʺFree to Airʺ) channel, for

which only a basic IRD will be needed.  The

schedule for television transmissions from the

orbiter and for mission briefings will be

available during the mission at Kennedy Space

Center, Fla.; Marshall Space Flight Center,

Huntsville, Ala.; Dryden Flight Research

Center, Edwards, Calif.; Johnson Space Center,

Houston; and NASA Headquarters,

Washington.  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 posted at:

http://www.nasa.gov/shuttle 

This site also contains information on the crew

and will be updated regularly with photos and

video clips throughout the flight.

Briefings 

A mission press briefing schedule will be issued

before launch.  The updated NASA television

schedule will indicate when mission briefings

are planned.

Internet Information 

Information on safety enhancements made

since the Columbia Accident is available at:

www.nasa.gov/returntoflight/system/index.html 

Information on other current NASA activities is

available at:

http://www.nasa.gov/home 

Resources for educators can be found at the

following address:

http://education.nasa.gov
August 2006 MEDIA CONTACTS 83

PUBLIC AFFAIRS CONTACTS 

HEADQUARTERS 

WASHINGTON, DC 

Allard Beutel

Public Affairs Specialist

Space Operations

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Katherine Trinidad

Public Affairs Specialist

Space Operations

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Grey Hautaluoma

Public Affairs Specialist

Space Operations

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Melissa Mathews

Public Affairs Specialist

International Partners

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JOHNSON SPACE CENTER 

HOUSTON, TEXAS 

James Hartsfield

News Chief

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Public Affairs Specialist

Space Shuttle Program Office

281‐483‐5111

Kylie Clem

Public Affairs Specialist

International Space Station

& Mission Operations

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Lynnette Madison

Public Affairs Specialist

Engineering

281‐483‐5111

Rob Navias

Program and Mission Operations Lead

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KENNEDY SPACE CENTER 

FLORIDA 

Bruce Buckingham

News Chief

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Jessica Rye

Public Affairs Specialist

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Tracy Young

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MARSHALL SPACE FLIGHT CENTER 

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Dom Amatore

Public Affairs Manager

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June Malone

Public Affairs Specialist

Space Shuttle Propulsion

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August 2006 MEDIA CONTACTS 84

STENNIS SPACE CENTER 

MISSISSIPPI 

Linda Theobald

Public Affairs Specialist

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Paul Foerman

News Chief

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Rebecca Strecker

Public Affairs Specialist

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AMES RESEARCH CENTER 

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News Chief

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Public Affairs Specialist

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DRYDEN FLIGHT RESEARCH CENTER 

CALIFORNIA 

Alan Brown

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GLENN RESEARCH CENTER 

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News Chief

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Katherine Martin

Public Affairs Specialist

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LANGLEY RESEARCH CENTER 

HAMPTON, VIRGINIA 

H. Keith Henry

Deputy, Office of Public Affairs

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UNITED SPACE ALLIANCE 

Mike Curie

Space Flight Operations Contract

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Kari Fluegel

Houston Operations

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Tracy Yates

Florida Operations

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BOEING 

Ed Memi

Media Relations

Boeing NASA Systems

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