STS-121 Press Kit 

December 2006 CONTENTS i

CONTENTS 

Section Page

STS-116 MISSION OVERVIEW: POWER RECONFIGURATION HIGHLIGHTS 

STATION ASSEMBLY MISSION ...............................................................................................1

STS-116 TIMELINE OVERVIEW ...............................................................................................10

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

LAUNCH AND LANDING ...........................................................................................................15

LAUNCH...............................................................................................................................................15

ABORT-TO-ORBIT (ATO)......................................................................................................................15

TRANSATLANTIC ABORT LANDING (TAL) .............................................................................................15

RETURN-TO-LAUNCH-SITE (RTLS).......................................................................................................15

ABORT ONCE AROUND (AOA)...............................................................................................................15

LANDING .............................................................................................................................................15

MISSION PROFILE...................................................................................................................16

STS-116 DISCOVERY CREW ...................................................................................................17

MISSION PERSONNEL .............................................................................................................27

RENDEZVOUS AND DOCKING ..................................................................................................28

UNDOCKING, SEPARATION AND DEPARTURE .......................................................................................31

INTERNATIONAL SPACE STATION ELECTRIC POWER SYSTEM (EPS) ...................................32

ACTIVE THERMAL CONTROL SYSTEM (ATCS) OVERVIEW.....................................................50

SPACEWALKS .........................................................................................................................65

EVA 1...................................................................................................................................................66

EVA 2 ..................................................................................................................................................67

EVA 3 ..................................................................................................................................................69

PAYLOAD OVERVIEW ..............................................................................................................71

INTEGRATED TRUSS SEGMENT P5.......................................................................................................71

INTEGRATED CARGO CARRIER (SPACEHAB)........................................................................................77

LOGISTICS SINGLE MODULE (SPACEHAB LSM) ...................................................................................78
December 2006 CONTENTS ii

Section Page

EXPERIMENTS .........................................................................................................................79

DETAILED TEST OBJECTIVES ...............................................................................................................79

SHORT-DURATION BIOASTRONAUTICS INVESTIGATION (SDBI) .........................................................79

SHORT-DURATION RESEARCH AND STATION EXPERIMENTS ...............................................................80

EUROPEAN SPACE AGENCY EXPERIMENTS...........................................................................................84

SPACE SHUTTLE MAIN ENGINE ADVANCED HEALTH MANAGEMENT SYSTEM ........................89

SHUTTLE REFERENCE DATA ....................................................................................................90

ACRONYMS AND ABBREVIATIONS .........................................................................................103

MEDIA ASSISTANCE ...............................................................................................................118

PUBLIC AFFAIRS CONTACTS ..................................................................................................119
December 2006 MISSION OVERVIEW 1

STS-116 MISSION OVERVIEW: 

POWER RECONFIGURATION HIGHLIGHTS 

STATION ASSEMBLY MISSION 

With its crane still attached, the orbiter Discovery was mated  

to the external tank and solid rocket boosters on the mobile  

launcher platform in high bay 3 of the Kennedy Space  

Center Vehicle Assembly Building.
December 2006 MISSION OVERVIEW 2

Space Shuttle Discovery launches in December

on its 33rd mission to deliver another truss

segment of the International Space Station and

begin the intricate process of reconfiguring and

redistributing the power generated by two

pairs of U.S. solar arrays.

The shuttle launches with seven astronauts—

six shuttle and one long‐duration station crew

member.  This is the first crew member rotation

in four years involving a shuttle rather than a

Russian Soyuz.

The primary assembly hardware Discovery will

deliver to the space station is the $11‐million

Integrated Truss Segment P5, which measures

11 feet long by 15 feet wide by 14 feet high (3.3

x 4.5 x 3.2 meters).  It will serve as a spacer and

be mated to the P4 truss that was attached in

September during the STS‐115 mission of

Atlantis.

Attachment of the 4,000‐pound

(1,800‐kilogram) P5 sets the stage for the

relocation to its final assembly position of the

P6 truss and the pair of solar arrays that have

been located temporarily atop the station’s

Unity module for six years.

A computer‐generated artistʹs rendering of the International Space Station after flight  

STS‐116/12A.1, following the delivery and installation of the third port truss segment  

(P5) and the retraction of the P6 port solar array wing and two radiators.
Three spacewalks (Extravehicular Activities or

EVAs) spread across the seven days of docked

operations will involve P5 installation and

reconfiguration of cables so that flight

controllers in Mission Control, Houston, can

send commands to swap power generation and

distribution from half of the P6 arrays to the

newest P4 pair (power channel 2/3 moves to

P4‐2A and 1/4 moves to power channel P4‐4A).

In addition to the P5 spacer, Discovery’s

payload bay also houses a small pressurized

logistics module holding supplies and an

integrated carrier delivering space station

hardware and three small satellites to be

deployed after the shuttle has undocked from

the space station.

The 20th shuttle mission to the International

Space Station represents the most

choreographed assembly flight to date between

December 2006 MISSION OVERVIEW 3

the shuttle and station crew members and flight

controllers in Mission Control, who will send

all commands to carefully redistribute power

and thermal management from one location to

another.  The STS‐118 mission in the summer of

2007 will deliver an identical short spacer (S5)

to the opposite end of the station’s truss.

Discovery will launch with seven crew

members, including Commander Mark

Polansky, Pilot William (Bill) Oefelein

(Commander, USN), and Mission Specialists

Nicholas Patrick, Robert (Bob) Curbeam

(Captain, USN), Joan Higginbotham, Christer

Fuglesang representing the European Space

Agency, and Sunita Williams.  Williams will

replace crew member Thomas Reiter (ESA) who

will return to Earth aboard Discovery in her

place.  Williams will return home next summer

following Endeavour’s STS‐118 mission.

Attired in their training versions of the shuttle launch and entry suit, astronauts  

Mark L. Polansky (left), STS‐116 commander, and William A. Oefelein, pilot,  

occupy the commander and pilotʹs station during a training session in the fixed‐ 

base shuttle mission simulator (SMS) in the Jake Garn Simulation and  

Training Facility at Johnson Space Center.
December 2006 MISSION OVERVIEW 4

The launch from complex 39B at the Kennedy

Space Center, Florida, is timed precisely to

occur within the same launch plane (similar to a

lane on a highway) as the space station to

maximize propellant savings and minimize

rendezvous time.

Because of the excellent performance of the

shuttleʹs external tank in minimizing foam

shedding during ascent, and the ability to

perform a 100 percent inspection of the orbiter

thermal protection system for unlikely damage,

the restriction for daylight‐only launches has

been lifted.  This allows Discovery’s launch to

take place at night for the first 10 days of its

“window,” which opens no earlier than Dec. 7

and closes on or about Dec. 26 based on a sun

beta angle constraint.

The first three days of the mission nearly mirror

those of the previous three shuttle flights to

inspect the thermal protection system tiles and

the wing leading edge reinforced carbon‐carbon

panels, and rendezvous and dock with the

International Space Station.

Patrick is the prime shuttle remote manipulator

system (robotic arm) operator and will lead the

inspection effort using the Remote Manipulator

System (RMS) extension – the Orbiter Boom

Sensor System.  Polansky and Oefelein serve as

backup shuttle arm operators.

The highest priority tasks of the flight will be to

transfer one station crewmember for another,

install the new P5 short spacer, reconfigure the

electrical power system and thermal control

system and transfer extra oxygen to storage

tanks on the outside of the U.S. Quest Airlock.

Astronaut Nicholas J. M. Patrick, STS‐116 mission specialist, participates  

in a training session in the crew compartment trainer (CCT‐2) in the  

Space Vehicle Mockup Facility at Johnson Space Center.
December 2006 MISSION OVERVIEW 5

In the Kennedy Space Center Space Station Processing Facility, an overhead crane  

moves the P5 truss for mission STS‐116 to the payload canister.  The third port  

truss segment, the P5 will be attached to the P3/P4 truss on the International  

Space Station during the 11‐day mission.

After docking, the first priority is to transfer

form‐fitting seat liners in the Soyuz spacecraft

making Williams an official member of the

Expedition 14 crew along with Commander

Michael Lopez‐Alegria and Flight Engineer

Mikhail Tyurin.  Reiter then becomes a member

of the shuttle crew with which he will return

home after a six‐month stay on the station.

On flight day 3, Patrick will carefully lift the P5

spacer with the shuttle RMS and hand it to the

waiting station arm.  Higginbotham and

Williams will control the station arm at the

station’s robotic work station in the Destiny

Laboratory.  The spacer will remain on the

station’s arm overnight in preparation for

installation the next day during the first of three

planned spacewalks.
December 2006 MISSION OVERVIEW 6

Astronauts Robert L. Curbeam, Jr. and Christer Fuglesang, STS‐116 mission specialists,  

wearing training versions of the Extravehicular Mobility Unit (EMU) spacesuit,  

participate in an underwater simulation of extravehicular activity (EVA).  

Curbeam and Fuglesang are dwarfed by station truss segments in this  

overall view of the simulation conducted in the Neutral Buoyancy  

Laboratory (NBL) near the Johnson Space Center.

The day after docking (flight day 4), Curbeam

and Fuglesang will leave the Quest Airlock on a

six‐hour spacewalk to assist with installation

and utility connections between the P5 short

spacer and the P4 truss.

The connection tasks include removal of four

launch locks with the two truss segments

approximately 6‐12 inches apart.  The

spacewalkers then will serve as on‐the‐scene

observers for alignment and installation of P5 to

P4.  The installation is completed with the

mating of six utility cables. Astronaut Robert L. Curbeam, STS‐116

mission specialist, attired in a training version

of the Extravehicular Mobility Unit (EMU)

spacesuit, awaits a training session in the

waters of the Neutral Buoyancy Laboratory

(NBL) near Johnson Space Center.
December 2006 MISSION OVERVIEW 7

European Space Agency (ESA) astronaut Christer Fuglesang, STS‐116 mission specialist,  

participates in an Extravehicular Mobility Unit (EMU) spacesuit fit check in the Space  

Station Airlock Test Article (SSATA) in the Crew Systems Laboratory at the Johnson  

Space Center.  Astronaut William A. Oefelein, pilot, assisted Fuglesang.

Between the first and second spacewalks begins

a process of power and thermal reconfiguration

that has never been attempted before.  Ground

commanding removes power from one half of

the P6 solar array followed by the retraction of

its port array.  A minimum of 40 percent of the

array must be retracted to provide enough

clearance for activation of the P4 solar array

tracking rotary joint tested during the previous

shuttle mission (STS‐115).  The starboard solar

array of P6 will be retracted during STS‐117

next March before installation of the S3/S4 set of

solar arrays on the integrated truss structure of

the station.

Though full retraction of the P6 port array is

not necessary, it is planned in a three‐step

process budgeted for five hours, with retraction

of three “bays” first.  Then the port array will

continue to be retracted to approximately 40

percent, and finally to one bay.

The next four hours includes filling one of the

thermal control systems with ammonia before

the final retraction of the solar array into its

canister.

Once automatic sun tracking is confirmed for

the new P4 arrays, the stage is set for the next

day’s second spacewalk to reconfigure power

to the outboard arrays.  This requires precise

coordination between the ground and crew to

ensure electrical power is not flowing.

The United States Orbital Segment (USOS)

electrical power system (EPS) is divided into

three main subsystems:  primary, secondary

and support systems.  The goal while Discovery

is docked is to reconfigure the stationʹs power

system from the current temporary status to its

“assembly complete” configuration.

The S0 truss segment sits in the middle position

on the truss structure on top of the U.S. Destiny
Laboratory, flanked by the S1 and P1 truss

elements.  That truss and the S1 and P1 trusses

contain the major electrical components of the

permanent electrical system.  Those are the

Main Bus Switching Units (MBSUs) and large

transformers called DC‐to‐DC converter units

(DDCUs) that serve to modulate solar array

power to the proper levels required to operate

station systems.

S1 and P1 also house the stationʹs two

independent cooling systems, each of which

include large ammonia tanks, a nitrogen gas

pressurization system and a massive pump

module to enable ammonia to flow through

plumbing lines to radiators that will dissipate

heat from the avionics systems on the station.  

There are three such radiators on S1 and three

December 2006 MISSION OVERVIEW 8

on P1.  To facilitate that heat rejection, the

radiators are mounted on a rotating beam that

can point them toward deep space and away

from the sun.

For the second time in three days, Curbeam and

Fuglesang will head out of the Quest Airlock on

the mission’s second spacewalk on the sixth

day of the mission to reconfigure part of the

power channel (2/3) by routing primary power

through the MBSUs.

Williams joins Curbeam on the third spacewalk

on the eighth day of the mission to do the same

reconfiguration on the other half of the power

channel (1/4).  Oefelein will serve as the in‐

cabin choreographer for all spacewalks and

spacesuit checkout.

Astronauts Sunita L. Williams (left), Expedition 14 flight engineer, and  

Joan E. Higginbotham, STS‐116 mission specialist, use the virtual reality  

lab at the Johnson Space Center to train for their duties aboard the space  

shuttle.  This type of computer interface, paired with virtual reality  

training hardware and software, helps to prepare the entire team  

for dealing with space station elements.
December 2006 MISSION OVERVIEW 9

Mounted on the central truss segment (S0)

launched in April 2002, these four MBSUs have

never been activated, but were checked out

during that mission and again in December

2002.

While the crew members prepare for the second

and third spacewalks, a lengthy set of power‐

down commands will be executed by flight

controllers to place all associated equipment in

a safe configuration before opening the Direct

Current Switching Unit remote bus isolators.

Because the MBSUs generate heat when current

is flowing, they require cooling via the

ammonia loops.  Estimates show the MBSUs

can run without cooling for about one hour, so

well‐choreographed commanding is planned to

activate the ammonia pump module to provide

cooling to the MBSUs within that timeframe.

Once the power reconfiguration is complete,

the station’s newest pair of solar arrays on the

port side of the truss will be brought to life to

provide electrical power to the station.

Throughout the mission, transfer of cargo from

the pressurized module (SPACEHAB) in

Discovery’s payload bay takes place to

resupply the station.

Hardware stowed in the module includes a

Video Baseband Signal Processor, a Rotary Joint

Motor Controller Assembly, an External TV

Camera Group (ETVCG), Oxygen Generation

System, Adjustable Grapple Bar; Remote Power

Control Module(s), Nickel Removal Assembly

Kit, Charcoal Bed Assembly,

Desiccant/Adsorbent replacement unit, Control

Moment Gyro Electrical Assembly and an

Avionics Air Assembly.

After the outside work is completed and before

Discovery departs, the station’s Mobile

Transporter will be relocated to the starboard

end of the truss and will undergo a checkout in

preparation for the next shuttle visit, scheduled

for March 2007, to deliver another truss

segment and the third pair of solar arrays – a

mirror‐image flight to that of Atlantis on the

STS‐115 mission in September.

On flight day 9, (the day after EVA 3), the

Mobile Transporter will be moved to the

starboard end of the truss and undergo a

checkout in preparation for its support of the

next assembly mission next March.

Once transfers are complete, the shuttle will

undock from the station, conduct a flyaround

and move to a station‐keeping distance of about

40 miles.

Back on its own, Discovery’s crew will oversee

a final inspection of the orbiter’s thermal

protection system to ensure it has not sustained

any damage from micrometeoroid debris before

the shuttle is cleared for entry.  The crew will

also remotely deploy three small technology‐

demonstration satellites mounted inside

canisters along an equipment carrier in the

payload bay.

The carrier also holds the Service Module

Debris Panels, 15 Adjustable Mass Plates and

an ISS Passive Flight Releasable Attachment

Mechanism.

Activities on the day before landing include

stowage of gear and checkout of orbiter entry

and landing systems, including the flight

control surfaces and thruster jets used for on‐

orbit and entry steering.

Discovery is scheduled to land the following

day at the Kennedy Space Center, completing

the 117th shuttle mission.

December 2006 TIMELINE OVERVIEW 10

STS-116 TIMELINE OVERVIEW 

FLIGHT DAY 1: 

• Launch

• Payload Bay Door Opening

• Spacehab Module Activation

• 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)

• OBSS Berthing

• 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 Discovery

• Docking to the International Space Station

• Hatch Opening and Welcoming by

Expedition 14 Crew

• Suni Williams joins Expedition 14 crew with

Soyuz seatliner transfer; Thomas Reiter

joins shuttle crew

• Shuttle robot arm grapple of P5 spacer truss

and handoff to station robot arm

• Curbeam and Fuglesang sleep in Quest

Airlock for spacewalk pre‐breathe campout

protocol

FLIGHT DAY 4: 

• Station robot arm installs P5 spacer truss

installation on P4 truss attachment

• Curbeam and Fuglesang EVA # 1 to connect

P5 / P4 power cables, release launch

restraints and to change out TV camera on

S1 truss

• Mobile Transporter moves from Worksite 7

to Worksite 3

FLIGHT DAY 5: 

• P6 truss port array is retracted to enable

Solar Alpha Rotary Joint activation and

rotation on P4 truss

• P4 Solar Alpha Rotary Joint activation and

autotracking of the sun

• Port side loop of External Active Thermal

Control System is filled with ammonia

• Curbeam and Fuglesang sleep in Quest

Airlock for spacewalk pre‐breathe

campout protocol
December 2006 TIMELINE OVERVIEW 11

FLIGHT DAY 6: 

• ISS power down of electrical channels 2

and 3

• Curbeam and Fuglesang EVA # 2 to

reconfigure electrical channels 2 and 3,

relocate Crew Equipment Translation Aid

(CETA) carts 1 and 2

• ISS Treadmill Vibration Isolation System

gyroscope replacement and maintenance

• Port side loop of the External Active

Thermal Control System is activated to

allow ammonia to flow

• ISS power up of electrical channels 2 and 3

FLIGHT DAY 7: 

• Shuttle to ISS transfer work

• Joint Crew News Conference

• Crew off duty time

• Starboard side loop of External Active

Thermal Control System is filled with

ammonia

• Curbeam and Williams sleep in Quest

Airlock for spacewalk pre‐breathe campout

protocol

FLIGHT DAY 8: 

• ISS power down of electrical channels 1

and 4

• Curbeam and Williams EVA # 3 to

reconfigure electrical channels 1 and 4 and

transfer Service Module Debris Panels to

Pressurized Mating Adapter 3

• Starboard side loop of the External Active

Thermal Control System is activated to

allow ammonia to flow

• ISS power up of electrical channels 1 and 4

FLIGHT DAY 9: 

• Shuttle to ISS transfer work

• Mobile Transporter moves to Worksite 2 for

S3 / S4 survey for STS‐117, then returns to

Worksite 4

• Rendezvous tool checkout in preparation

for undocking

FLIGHT DAY 10: 

• Final transfer work

• Farewells and Hatch Closing

• Undocking and ISS flyaround

• Final separation from ISS

• MEPSI pico‐satellite deploy

• ANDE pico‐satellite deploy

FLIGHT DAY 11: 

• Flight Control System Checkout

• Reaction Control System Hot‐Fire Test

• Cabin Stowage

• RAFT pico‐satellite deploy

• Deorbit Timeline Review

• Recumbent Seat Set Up for Reiter in

middeck

• Ku‐Band Antenna Stowage

FLIGHT DAY 12: 

• Deorbit Preparations

• Payload Bay Door Closing

• Deorbit Burn

• KSC Landing
December 2006 MISSION PRIORITIES 12

MISSION PRIORITIES 

1. Perform inspection of space shuttle

reinforced carbon‐carbon (RCC) and

downlink sensor data for evaluation on the

ground.

2. Document space shuttle tile during

rendezvous with station using ISS imagery

resources during the rendezvous pitch

maneuver (RPM), followed by docking.

3. Complete ISS crew member swap

(Expedition 14 Flight Engineer Suni

Williams for Expedition 13 Flight Engineer

Thomas Reiter).

• Install the Soyuz seat liner, known as

the Individual Equipment Liner Kit

(IELK)

• Check out the Russian launch/entry suit,

known as the Sokol suit

• ISS safety briefing

• Transfer mandatory crew rotation items

(a) Transfer required oxygen to ISS (~100

pounds).

(b) Transfer and return Elektron.

4. Transfer water.

5. Install the P5 truss segment onto P4 using

the shuttle and station robotic arms.

• Remove P5 inboard launch locks

(required for mating with P4)

• Install four truss attachment bolts to

structurally mate P5 to P4

• Remove P5 grapple fixture and relocate

to P5 keel (will allow P4 beta gimbal

assembly to rotate)

6. Deactivate P6 2B loads and reconfigure U.S.

segment loads to receive power distribution

from P4 2A and P6 EB via main bus

switching units 2 and 3.  This includes

establishment of active cooling for channel

2/3 MBSUs and DC‐to‐DC converter units

via external active thermal control system

loop B.

• Retract P6 4B solar array wing to one

bay and initiate P3/P4 solar alpha rotary

joint tracking.

• Remove P1‐3A DC‐to‐DC converter

unit‐E thermal covers.

7. Deactivate P6 4B loads and reconfigure U.S.

segment loads to receive power distribution

from P4 4A main buss switching unit 1 and

4.  This includes establishment of active

cooling for channel 1⁄4 MBSUs/DDCUs via

external active thermal control system loop

A.  (P6 4B channel configured to

dormant/parachute mode.

• Remove S1‐4B and S0‐4B DC‐to‐DC

converter unit‐E thermal covers.

• Uplink the D1 patch to portable

computer system R9.

8. Transfer critical cargo items per transfer

priority list.

9. Transfer Zvezda Service Module debris

panels and adapter to pressurized mating

adapter‐3 aft grapple fixture.
December 2006 MISSION PRIORITIES 13

10. Relocate both Crew and Equipment

Translation Aid (CETA) carts from the

starboard side to the port side.

(a) Perform contingency spacewalk to

complete primary mission objectives.

(b) Perform late inspection of Discovery’s

wing leading edge and nosecap.

11. Perform minimum crew handover

(12 hours) for rotating crew members.

(a) Perform the Oxygen Recharge

Compressor Assembly and Carbon

Dioxide Removal Assembly removal

and replacement and return removed

hardware via shuttle.

12. Perform utilization activities to support

experiments, including midodrine, ALTEA,

Latent Virus, Sleep Short, and PMDIS.

13. Perform daily ISS payload status checks as

required.

14. Transfer remaining cargo items per mission

rules.

15. Perform external wireless instrumentation

system power connections between P5 and

P4.

16. Remove and replace External Television

Camera Group (ETVCG) at Camera Port 3,

Starboard 1 Outboard Lower.

17. Transfer the adjustable grapple bar from

inside the station to the flex hose rotary

coupler on external stowage platform‐2.

18. Perform P6 4B final retraction and latching

of the solar array blanket box.

19. Install power cables for S0 channels 1/4 2/3.

20. Perform payload operations to support STP‐

H2 (ANDE, MEPSI, RAFT).

21. Perform the following to allow return of

on‐orbit hardware:

• Treadmill gyro removal and

replacement

• Charcoal bed assembly

• Respiratory support pack checkout

22. Transfer nitrogen from the shuttle to the ISS

Quest Airlock high pressure tanks.

23. Perform U.S. and Russian payload research

operation tasks.

24. Perform an additional four hours of ISS

crew handover (16 hours total).

25. Perform imagery survey of the ISS exterior

from shuttle after undocking.

26. Perform payload operations to support

Maui Analysis of Upper Atmospheric

Injections (MAUI) and Ram Burn

Observations (RAMBO)

27. Reboost ISS (altitude TBD based on

available shuttle propellant).

28. The following tasks fit within the existing

spacewalk timelines; however, they may be

deferred if the spacewalk is behind

schedule.  The EVA will not be extended to

complete these tasks.

• Install station robot arm force moment

sensor (FMS) insulation

• Install the starboard and port fluid

quick disconnect bags on the Quest

Airlock
December 2006 MISSION PRIORITIES 14

• Install S0/Unity Node primary power

cable (S0 side only) and reconfigure Z1

patch panels and Russian power to

operate from primary power (i.e.,

MBSU)

29. Perform program‐approved spacewalk get‐

ahead tasks.  The following get‐ahead tasks

do not fit in the existing spacewalk

timelines; however, the team will be trained

and ready to perform any of these tasks

should the opportunity arise.

• Connect P5 to P4 umbilicals (6)

• Open P5 capture latch assembly (CLA)

and partially close (~ 1 turn)

• Remove P5 to P6 truss attachment

system launch locks

• Install the pump module jumper bag on

the Quest Airlock

• Install the vent tool extension bag on

Quest

30. Perform:

• Development Test Objective (DTO) 257

– Structural Dynamics Model Validation

Flight Test and Supplementary

Objectives Document (Internal Wireless

Instrumentation System, known as

IWIS, is required)

• Perform ISS Structural Life Validation

and Extension for the shuttle undocking

(IWIS required).

December 2006 LAUNCH & LANDING 15

LAUNCH AND LANDING 

LAUNCH 

As with all previous Space Shuttle launches,

Discovery on STS‐116 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 9i0 minutes after liftoff.

LANDING 

The primary landing site for Discovery on

STS‐116 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.

December 2006 MISSION PROFILE 16

MISSION PROFILE 

CREW 

Commander: Mark Polansky

Pilot: Bill Oefelein

Mission Specialist 1: Nicholas Patrick

Mission Specialist 2: Bob Curbeam

Mission Specialist 3: Christer Fuglesang

Mission Specialist 4: Joan Higginbotham

Mission Specialist 5: Suni Williams/

Thomas Reiter

LAUNCH 

Orbiter: Discovery (OV‐103)

Launch Site: Kennedy Space Center

Launch Pad 39B

Launch Date: No Earlier Than

December 7, 2006

Launch Time: 9:36 p.m. EST (Preferred

In‐Plane launch time for

12/7)

Launch Window: 5 Minutes

Altitude: 123 Nautical Miles (142

Miles) Orbital Insertion;

190 NM (218 Miles)

Rendezvous

Inclination: 51.6 Degrees

Duration: 10 Days 18 Hours 40

Minutes

VEHICLE DATA 

Shuttle Liftoff Weight: 4,521,350

pounds

Orbiter/Payload Liftoff Weight: 265,466

pounds

Orbiter/Payload Landing Weight: 225,350

pounds

Software Version: OI‐30

Space Shuttle Main Engines: 

SSME 1: 2050

SSME 2: 2054

SSME 3: 2058

External Tank: ET‐123

SRB Set: BI‐128

RSRM Set: 95

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

December 18, 2006

Landing Time: 4:16 p.m. EST

Primary landing Site: Kennedy Space Center

Shuttle Landing Facility

PAYLOADS 

lntegrated Truss Segment (ITS)‐P5, SPACEHAB

Single Module

December 2006 CREW 17

STS-116 DISCOVERY CREW 

The STS‐116 patch design signifies the

continuing assembly of the International Space

Station.  The primary mission objective is to

deliver and install the P5 truss element.  The P5

installation will be conducted during the first of

three planned spacewalks, and will involve use

of both the shuttle and station robotic arms.  

The remainder of the mission will include a

major reconfiguration and activation of the

stationʹs electrical and thermal control systems,

as well as delivery of Zvezda Service Module

debris panels, which will increase protection

from potential impacts of micro‐meteorites and

orbital debris.  In addition, a single expedition

crew member will launch on STS‐116 to remain

on board the station, replacing an expedition

crew member Thomas Reiter, who will fly

home with the shuttle crew.  The crew patch

depicts the space shuttle rising above the Earth

and the station.  The United States and Swedish

flags trail the orbiter, depicting the

international composition of the STS‐116 crew.  

The seven stars of the constellation Ursa Major

are used to provide direction to the North Star,

which is superimposed over the installation

location of the P5 truss on the station.
December 2006 CREW 18

These seven astronauts take a break from

training to pose for the STS‐116 crew portrait.  

Scheduled to launch aboard the Space Shuttle

Discovery are, front row (from the left),

astronauts William A. Oefelein, pilot; Joan E.

Higginbotham, mission specialist; and Mark L.

Polansky, commander.  On the back row (from

the left) are astronauts Robert L. Curbeam,

Nicholas J.M. Patrick, Sunita L. Williams and

the European Space Agencyʹs Christer

Fuglesang, all mission specialists.  Williams will

join Expedition 14 in progress to serve as a

flight engineer aboard the International Space

Station.  The crew members are attired in

training versions of their shuttle launch and

entry suits.

Short biographical sketches of the crew follow

with detailed background available at:

http://www.jsc.nasa.gov/Bios/
December 2006 CREW 19

Commander Mark Polansky

A former Air Force test pilot, Mark Polansky

will lead the crew of STS‐116 on the 20th shuttle

mission to the space station.  Polansky served

as the pilot on STS‐98 in 2001.  Making his

second spaceflight, he has logged more than

309 hours 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, Polansky will fly the shuttle in a

procedure called the rendezvous pitch

maneuver while Discovery is 600 feet below the

station before docking to enable the station

crew to photograph the orbiterʹs heat shield.  

He will then dock Discovery to the station.  

Polansky will also be heavily involved in

shuttle robotic arm operations for inspecting

the orbiter’s heat shield, and transferring cargo

to the station during the docked phase of the

mission.
December 2006 CREW 20

Pilot Bill Oefelein

William Oefelein (Ohʹ‐fe‐line), who has logged

more than 3,000 hours flying more than 50

aircraft, will make his first journey into space as

the pilot for the STS‐116 mission.  Selected by

NASA in June 1998, Oefelein reported to the

Johnson Space Center in Houston in August

1998.  He has served in the Astronaut Office

Advanced Vehicles Branch and CAPCOM

(capsule communicator) Branch.  He will be

responsible for orbiter systems operations and

assisting Polansky in the rendezvous and

docking to the International Space Station.  He

will also serve as the choreographer inside

Discovery and the station for the missionʹs

three planned spacewalks, helping to suit up

and direct the spacewalkers through their

activities.  Oefelein will undock Discovery from

the station at the end of the docked phase of the

mission and conduct a flyaround to enable his

crewmates to photograph the stationʹs

configuration and assess its condition.
December 2006 CREW 21

Mission Specialist Nicholas Patrick

A member of the 1998 astronaut class and a

former flight instructor, Nicholas Patrick is

assigned to STS‐116 as mission specialist 1

(MS‐1).  He reported to NASAʹs Johnson Space

Center in Houston for astronaut training in

August 1998.  His initial training included

scientific and technical briefings, intensive

instruction in shuttle and International Space

Station systems, physiological, survival and

classroom training in preparation for T‐38

flight.  Making his first spaceflight, Patrick will

be the primary operator of the shuttleʹs robotic

arm, using it to unberth the orbiter boom sensor

system to survey Discoveryʹs thermal

protection system on flight day 2 and to

grapple the station’s P5 truss for a handoff to

the station robotic arm operated by Mission

Specialist Joan Higginbotham on flight day 3.  

He will be responsible for the shuttle’s video

and computer networks, and will assist with

the transfer of cargo between the shuttle and

the station.  He will be seated on the flight deck

for launch and on the middeck for landing.
December 2006 CREW 22

Mission Specialist Bob Curbeam

A veteran of two space shuttle flights, Mission

Specialist 2 (MS 2) Bob Curbeam conducted

three spacewalks before being assigned to

STS‐116.  He flew on STS‐85 in August 1997 and

on STS‐98 in February 2001.  He logged more

than 19 hours over the course of three

spacewalks during STS‐98, completing delivery

of the U.S. laboratory Destiny to the space

station.  He will conduct three spacewalks

during STS‐116.  Curbeam, as EV 1, will

conduct the first two spacewalks of the mission

with Christer Fuglesang on flight days 4 and 6.  

During the first spacewalk, the two will install

the P5 truss and attach all mechanical and

electrical interfaces between it and the existing

station truss.  They also will change out a TV

camera on the starboard 1 (S1) truss.  During

the second spacewalk, the duo will unplug

station power channels 2 and 3 from the P6

power truss and connect them to the main truss

(permanent) power system.  They also will

move the Crew and Equipment Translation Aid

carts in preparation the next power

reconfiguration to occur during the third

spacewalk of the mission.  Curbeam will

conduct the third planned spacewalk of the

mission on flight day 8 with Sunita Williams.  

They will unplug station power channels 2 and

3 from the P6 power truss and connect them to

the main truss (permanent) power system.  

They also will transfer Service Module debris

panels from the shuttle to the station.  Curbeam

will be seated on the flight deck for launch and

landing, operating as the flight engineer to

assist Commander Mark Polansky and Pilot Bill

Oefelein.
December 2006 CREW 23

Mission Specialist Christer Fuglesang

Making his first spaceflight, European Space

Agency astronaut Christer Fuglesang

(Fyuʹ‐gel‐sang) joins the crew of STS‐116 as a

mission specialist.  Fuglesang is a member of

ESAʹs European Astronaut Corps, whose home

base is the European Astronaut Centre in

Cologne, Germany.  He entered the mission

specialist class at NASAʹs Johnson Space Center

in August 1996 and qualified for flight

assignment as a mission specialist in April 1998.  

Mission Specialist 3 (MS 3) Fuglesang, as EV 2,

will conduct the first two planned spacewalks

of the mission with Curbeam on flight days 4

and 6.  Fuglesang is the lead for deploying three

small satellites from the payload bay toward

the end of the mission.  Fuglesang will set up

the recumbent seat returning Expedition 14

crew member Thomas Reiter will use for the

trip home aboard the shuttle.  Fuglesang will be

seated on the middeck for launch and landing.
December 2006 CREW 24

Mission Specialist Joan Higginbotham

Mission Specialist 4 (MS 4) Joan Higginbotham

will be making her first flight into space aboard

Discovery.  To assist with the construction of

the space station, Higginbothamʹs primary task

on STS‐116 will be to operate the stationʹs

robotic arm.  Among other robotic tasks, she

will use the station arm to install the P5 truss

onto the P4 truss attachment on flight day 4.  

During the rendezvous, docking and

undocking, she will manage the rendezvous

navigation tools used to guide the shuttleʹs

trajectory relative to the station.  She will serve

as the lead cargo transfer officer, overseeing the

transfer of supplies and equipment between the

shuttle and the station.  She will oversee

payload bay door closing operations.  She will

be seated on the middeck for launch and the

flight deck for landing.
December 2006 CREW 25

Mission Specialist/Expedition 14 Flight Engineer Suni Williams

Making her first spaceflight, Mission Specialist

5 (MS 5) Sunita (Soo‐neeʹ‐tah) Williams will join

Expedition 14 in progress and serve as a flight

engineer after traveling to the station on space

shuttle mission STS‐116.  Williams, who goes by

the name Suni (sunny), will join Expedition 14

on flight day 3, when her Soyuz seatliner is

transferred from the shuttle 3.  The transfer will

mark the beginning of her scheduled six‐month

stay aboard the station.  Williams, as EV 3, will

join Bob Curbeam for the third planned

spacewalk of the mission on flight day 8.  She

will be seated on the middeck for launch.
December 2006 CREW 26

Thomas Reiter

International Space Station Flight Engineer

Thomas Reiter (Toe‐mahsʹ Ryeʹ‐turr) (FE 2) of

the European Space Agency (ESA) flew to the

space station aboard Discovery in July 2006 and

became a member of the Expedition 13 crew.  

He will return to Earth aboard Discovery with

the STS‐116 crew.  In September 2006,

Expedition 13 Commander Pavel Vinogradov

and Flight Engineer and NASA Science Officer

Jeff Williams left the station aboard a Russian

Soyuz spacecraft.  Reiter was joined by

Expedition 14 Commander Mike Lopez‐Alegria

and Flight Engineer Mikhail Tyurin.  Reiter is

the first crew member to serve on two

expeditions.  He spent 179 days in space in

1995‐1996 for a mission to the Russian Mir

space station during which he conducted two

spacewalks and about 40 European scientific

experiments.  Reiter is the first ESA astronaut to

live aboard the International Space Station for a

long‐term mission.  Reiter worked on the

station as part of an agreement between the

Russian Federal Space Agency and ESA.  Reiter

will be on the middeck for landing in a

specially designed recumbent seat to facilitate

his adaptation to a gravity environment for the

first time in six months.
December 2006 MISSION PERSONNEL 27

MISSION PERSONNEL 

KEY CONSOLE POSITIONS FOR STS-116 

Flt. Director CAPCOM PAO

Ascent Steve Stich Ken Ham Kelly Humphries

Chris Ferguson (Wx)

Orbit 1 (Lead) Tony Ceccacci Kevin Ford Kyle Herring (Lead)

Orbit 2 Matt Abbott Megan McArthur Nicole Cloutier

Planning Rick LaBrode Shannon Lucid John Ira Petty

Entry Norm Knight Ken Ham Kyle Herring

Chris Ferguson (Wx)

Shuttle Team 4 Richard Jones N/A N/A

ISS Orbit 1 Derek Hassmann Terry Virts N/A

ISS Orbit 2 (Lead) John Curry Steve Robinson N/A

ISS Orbit 3 Joel Montalbano Hal Getzelman N/A

Station Team 4 Dana Weigel N/A N/A

Mission Control, Kwatsi Alibaruho N/A N/A

Korolev, Russia

JSC PAO Representative at KSC for Launch – Kylie Clem

KSC Launch Commentator – Bruce Buckingham

KSC Launch Director – Mike Leinbach

NASA Launch Test Director – Jeff Spaulding

December 2006 RENDEZVOUS & DOCKING 28

RENDEZVOUS AND DOCKING 

Discovery’s final approach to the International

Space Station during the STS‐116 rendezvous

and docking process will include the

now‐standard back flip pirouette maneuver to

allow station crewmembers to take digital

images of the shuttle’s heat shield.

With shuttle Commander Mark Polansky at the

controls, the shuttle will perform the circular

pitch‐around from a distance of about 600 feet

below the station.  The 9‐minute flip offers

Expedition 14 Commander Mike Lopez‐Alegria

and Flight Engineer Mikhail Tyurin time to

document through digital still photography the

required imagery of Discovery’s thermal

protection system.

The photos then will be transmitted to imagery

experts in the Mission Evaluation Room at

Mission Control, Houston, via the station’s

Ku‐band communications system.

The photography will be performed out of

windows 6 and 8 in the Zvezda Service Module

with Kodak DCS 760 digital cameras and

400 mm and 800 mm lenses.  The Rendezvous

Pitch Maneuver (RPM) is one of several

inspection procedures designed to verify the

integrity of the shuttle’s protective tiles and

reinforced carbon‐carbon wing leading edge

panels.

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

The sequence of events that culminate with

Discovery’s docking to the station actually

begins with the precisely timed launch that

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.

places the orbiter on course for its two‐day

chase to arrive at the station.  The 43‐hour

rendezvous includes periodic thruster firings

that ultimately will place Discovery about 9

statute miles behind the station, the starting

point for final approach.

About 2.5 hours before the scheduled docking

time on flight day 3, Discovery will reach a

point about 50,000 feet behind the station.  

Discovery’s jets will be fired in what is called

the Terminal Initiation (TI) burn to begin the

final phase of the rendezvous.  Discovery will

close the final miles to the station during the

next orbit.
As Discovery 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, Discovery will execute four

small mid‐course corrections with its steering

jets to position the shuttle about 1,000 feet

directly below the station.  From this point,

Polansky will take over the manual flying of the

shuttle up an imaginary line drawn between

the station and the Earth known as the R‐Bar—

or radial vector.

He will slow Discovery’s approach at about 600

feet and, if required, wait for proper lighting

conditions to optimize inspection imagery

gathering as well as crew visibility for the final

rendezvous to docking.

December 2006 RENDEZVOUS & DOCKING 29

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.
bottomside_800mm.cnv

NOTE

1. indicates

critical focus

point.

2. Sequence is

~16 shots;

repeat

sequence, as

time allows.

December 2006 RENDEZVOUS & DOCKING 30

On verbal confirmation by Pilot Bill Oefelein to

alert the station crew, Polansky will command

Discovery to begin a nose forward, three‐

quarters of a degree per second rotational back

flip.  At RPM start, the station crew will begin a

series of precisely‐timed photography for

inspection.  The sequence of mapping

optimizes the lighting conditions.

Both the 400 mm 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 3‐ inch resolution and the

800 mm lens can provide up to 1‐inch

resolution and detect any gap filler protrusions

greater than 1/4 inch.  The imagery includes the

upper surfaces of the shuttle as well as

Discovery’s underside, nose cap, landing gear

door seals and the elevon cove areas with

1‐inch analytical resolution.  The photography

includes detection of any gap filler protrusions

when the orbiter is at 145 and 230‐degree angles

during the flip.  The maneuver and lighting

typically offers enough time for two sets of

pictures.

When Discovery completes its rotation, it will

return to an orientation with its payload bay

facing the station.

Polansky then will move Discovery 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.  Oefelein

will provide navigation information to

Polansky as the shuttle inches toward the

docking port at the forward end of the station’s

Destiny Laboratory.

Oefelein will join Mission Specialists Nicholas

Patrick and Joan Higginbotham in playing key

roles in the rendezvous.  They will operate

laptop computers processing the navigational

data, the laser range systems and Discovery’s

docking mechanism.

Using a camera view from center of Discovery’s

docking mechanism as a key alignment aid,

Polansky will precisely match the docking ports

of the two spacecraft and fly to a point 30 feet

from the station before pausing to verify the

alignment.

For Discovery’s docking, Polansky will close

the final 30 feet at a relative speed of about

one‐tenth of a foot per second (while both

spacecraft are traveling 17,500 mph), and keep

the docking mechanisms aligned within a

tolerance of three inches.
December 2006 RENDEZVOUS & DOCKING 31

At contact, preliminary latches will

automatically attach the two spacecraft.  

Immediately after Discovery docks, the

shuttleʹs steering jets will be deactivated to

eliminate 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.

Once motion between the two spacecraft has

been stopped, Mission Specialists Bob Curbeam

and Christer Fuglesang will secure the docking

mechanism, sending commands for Discovery’s

docking ring to retract and to close a final set of

latches between the two vehicles.

UNDOCKING, SEPARATION AND 

DEPARTURE 

With additional inspections of Discovery’s heat

shield expected to be scheduled after

undocking, the orbiter will depart the station

with the shuttle robotic arm and Orbiter Boom

Sensor System (OBBS) in their stowed

configuration.  The OBSS will be unstowed to

accommodate the inspections.

Once Discovery is ready to undock, Fuglesang

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.  

Discovery’s steering jets will be shut off to

avoid any inadvertent firings during the initial

separation.

Once Discovery is about two feet from the

station, with the docking devices clear of one

another, Oefelein will activate the steering jets

to very slowly move away.  From the aft flight

deck, Oefelein manually will control Discovery

within a tight corridor as it separates from the

station—essentially the reverse of the task

performed by Polansky during rendezvous.

Discovery will continue away to a distance of

about 450 feet, where Oefelein will guide the

shuttle in a circular flyaround of the station.  

Once Discovery completes 1.5 revolutions of

the complex, Oefelein will fire Discovery’s jets

to depart the station’s vicinity for the final time.

Discovery will separate to a distance of about

40 nautical miles and remain there to protect for

a return to the complex in the unlikely event

the late inspection reveals any damage to the

shuttle’s thermal heat shield.
December 2006 ELECTRIC POWER SYSTEM 32

INTERNATIONAL SPACE STATION 

ELECTRIC POWER SYSTEM (EPS) 

The International Space Station (ISS) electrical

power system consists of power generation,

energy storage, power management, and

distribution (PMAD) equipment.  Electricity is

generated in a system of solar arrays.  Besides

the solar arrays on the Russian element, the

station currently has two photovoltaic modules,

a term that refers to a set of solar arrays,

batteries and the associated electronics, on

orbit, with two more scheduled for delivery.

The Electric Power System (EPS) provides all

user loads and housekeeping electrical power

and is capable of expansion as the station is

assembled and grows.  Eight independent

power channels for high overall reliability

supply the electric power.

A photovoltaic (PV) electric power generation

subsystem was selected for the space station.  A

PV system uses solar arrays for power

generation and chemical energy storage

(Nickel‐hydrogen) batteries to store excess solar

array energy during periods of sunlight and

provide power during periods when the station

is in Earth’s shadow (eclipse).  The station

orbits the earth every 90 minutes and for about

35 minutes, the station must run on batteries

while the station is in eclipse.

Flexible, deployable solar array wings that are

covered with solar cells provide power for the

ISS.  Each PV module contains two wings, and

each wing consists of two blanket assemblies.  

The solar array wings are tightly folded inside a

blanket for launch.  They are deployed in orbit

and supported by an extendable mast.

Analogy between municipal utility  

and the station’s EPS

Nominal electrical output of each power

channel is about 11 kilowatts (kW), or 20.9 kW

per PV module.  Four PV modules will supply

approximately 83.6 kW.

The primary purpose of the Energy Storage

Subsystem (ESS) is to provide electrical power

during periods when power from the solar

arrays is not enough to support channel loads.  

The ESS stores energy during periods when

solar arrays can generate more power than

necessary to support loads.  The system consists

of three nickel‐hydrogen (Ni‐H2) batteries per

power channel and each battery consists of two

battery Orbital Replacement Units (ORUs).  

Each battery also has a charge/discharge unit

(BCDU).  The Ni/H2 battery design was chosen

because of its high energy density light weight

and proven heritage in space applications since

the late 1970s to early 1980s.

The entire EPS may be divided into two power

subsystems.  The primary power subsystem

operates at a voltage range of 137 to 173 volts

direct current (Vdc) and consists of power

generation, storage and primary power

distribution.  The secondary power subsystem
December 2006 ELECTRIC POWER SYSTEM 33

operates at a voltage range of 123 to 126 Vdc

and is used to supply power to user loads.  

Direct Current‐to‐Direct Current Converter

Units (DDCUs) are used to convert primary

power to secondary power.

The U.S. power system is also integrated with

Russian power sources, so that power from the

American power bus can be transferred to the

Russian power bus and vice versa.  The Russian

power system operates at a nominal voltage of

28 Vdc.  American to Russian Converter Units

(ARCUs) and Russian to American Converter

Units (RACUs) are used to convert power from

the American secondary power bus to the

Russian power bus and vice versa.

SOLAR POWER 

The most powerful solar arrays ever to orbit

Earth capture solar energy to convert it into

electric power for the ISS.

Eight solar array wings supply power at an

unprecedented voltage level of 137 to 173 Vdc

that is converted to a nominal 124 Vdc to

operate equipment on the ISS.  The Space

Shuttle and most other spacecraft operate at

nominal 28 Vdc, as does the Russian ISS

segment.

The higher voltage meets the higher overall ISS

power requirements while permitting use of

lighter‐weight power lines.  The higher voltage

reduces ohmic power losses through the wires.  

Some eight miles of wire distribute power

throughout the station.

Each PV module contains two solar array

wings.  An individual wing is 110‐feet long by

38‐feet wide.  Each wing consists of two array

blankets that are covered with solar cells.  The

blankets can be extended or retracted by a

telescopic mast which is located between the

two blankets.  Each solar array wing is

connected to the ISSʹs 310‐foot long truss and

extend outward at right angles to it (P4 and P6

are currently on orbit).  A series of 400 solar

cells, called a string, generates electricity at high

primary voltage levels while 82 strings are

connected in parallel to generate adequate

power to meet the power requirement for each

power channel.  There are a total of 32,800 cells

per power channel or 65,600 solar cells on each

PV module.

A solar cell assembly is about three inches

square.  The cells are made of silicon and have a

nominal 14.5 percent efficiency for sunlight‐to‐

electricity conversion.  Cells are welded on to a

flexible printed circuit laminate that connects

cells electrically.  The sun‐facing surface of the

cell is protected by a thin cover glass.  Each

group of eight cells, connected in series, is

protected by a bypass diode to minimize

performance impact of fractured or open cells

on a string.  Solar arrays are designed for an

operating life of 15 years.

Two mutually perpendicular axes of rotation

are used to point solar arrays towards the Sun.  

Each solar array wing is connected to one Beta

Gimbal Assembly (BGA), located on each PV

module, that is used to rotate that solar array
wing.  Another rotary joint, called Solar Alpha

Rotary Joint (SARJ), is mounted on the truss

and rotates the four solar array wings together.  

When the station is complete, there will be

eight BGAs and two SARJs.  These rotary joints

are computer controlled and ensure full sun‐

tracking capability as the ISS goes around the

earth under a wide range of orbits and ISS

orientations.

ELECTRIC POWER SYSTEM OVERVIEW 

Like a cityʹs central power plant, the PV

modules generate primary power at voltage

levels too high for consumer use, ranging from

137‐to‐173 Vdc.  The primary power is routed to

BCDUs for charging batteries and to switching

units that route it to local distribution networks.

The DC‐DC Converter Units, DDCUs,

ʺstep‐downʺ the primary power to a more

tightly regulated secondary power voltage,

nominally 124.5 Vdc that is regulated plus or

minus 1.5 Vdc, and distribute it to ISS loads.  

On Main Street, USA, the users would be shops

and homes.  On the ISS, they are laboratories,

living quarters and other modules.

Even though the Station spends about one‐third

of every orbit in Earthʹs shadow, the electrical

power system continuously provides usable

power (about 84 kW at assembly complete) to

ISS systems and users.  When the ISS is in

eclipse, the batteries that stored energy from

solar arrays during the sunlit portion of the

orbit supply power.

December 2006 ELECTRIC POWER SYSTEM 34

The power system is cooled by a thermal

system through which excess heat is removed

by liquid ammonia coolant in tubes that

ultimately loop through radiator panels that

radiate the heat to space.

Russiaʹs segment of the ISS provides its own

power sources, supplying 28‐volt‐dc to the

Russian modules.  Power is shared between the

two segments when required to support

assembly and operations for all ISS partners.  

Russian‐to‐American Converter Units (RACUs)

and American‐to‐Russian Converter Units

(ARCUs) step‐up and step‐down converters,

respectively, deal with the difference between

U.S. and Russian bus voltage levels.  As ISS

assembly continues, Russian solar arrays

(a 72‐foot pair on Control Module Zarya and a

97‐foot pair on the Russian Service Module)

will receive more shadow, which will diminish

their power generation capability.

The overall design and architecture of the ISS

EPS was managed by NASA’s Glenn Research

Center in the early 1990s.  Boeing’s Rocketdyne

Propulsion and Power division (now Pratt &

Whitney RPP) built most of the hardware for

the electrical power system.  Lockheed Martin

built the solar arrays and the Solar Alpha

Rotary Joint for Rocketdyne.  Boeing, along

with Pratt & Whitney RPP, as a subcontractor,

continues to provide EPS sustaining

engineering to NASA.  Most EPS components

and cargo assemblies undergo final acceptance

testing at Kennedy Space Center before flight to

ISS.
December 2006 ELECTRIC POWER SYSTEM 35

Electrical Power Distribution Overview

EPS BLOCK DIAGRAM OVERVIEW 

This block diagram gives an overview of how

the stationʹs electrical system functions when

assembly is complete.   The Solar Array Wing

(SAW) can generate power at a wide range of

voltage, however, the Sequential Shunt Units

(SSU), located close to the SAW in the

Integrated Equipment Assembly (IEA), regulate

the voltage that comes out of the solar arrays at

an established setpoint of about 160 Vdc.  When

a solar array can produce sufficient power, then

the surplus power is routed to the Battery

Charge/Discharge Units (BCDU), which charge

the batteries.  When a solar array cannot

produce sufficient power to satisfy ISS loads

then the bus voltage starts to drop below the

SSU setpoint, and when it drops below the

BCDU setpoint, then the BCDUs start to

discharge batteries to support ISS loads.  The

primary bus voltage varies between the SSU

and BCDU voltage setpoints plus a small

voltage regulation band.

The primary power is provided to the Main Bus

Switching Units (MBSU) for subsequent

distribution to ISS electrical loads.  Four MBSUs

are located on the S0 truss that is fed by eight

independent power channels and the MBSU

outputs supply all ISS loads.  Under normal

operations, each power channel supplies power

to a specific set of loads.  However, if that

channel fails, the MBSU enables feeding power

to those loads from another channel.  This

greatly enhances the failure tolerance of the

EPS.

All EPS operations are computer controlled and

controls can be exercised by the on‐orbit crew

or by operators on ground.  Operators on the

ground to free up crew time for more important

on‐orbit operations perform most of these

functions.  All control setpoints are stored on

on‐orbit computers and can be changed when

needed.
The MBSUs route power to the DC‐to‐DC

Converter Units (DDCUs).  The DDCUs convert

primary power to secondary power at 123 to

126 Vdc.  Several DDCUs are located inside

pressurized compartments, such as US Lab,

while several are located externally on trusses.  

DDCUs supply regulated secondary power to

Remote Power Controller Modules (RPCMs).  

RPCMs are boxes with multiple switches with

several different load ratings to route power to

user loads.  The RPCMs provide remote

switching of loads and over‐current protection.  

An RPCM can also feed other RPCMs and can

feed Russian power converters, outlet panels,

etc.  There will be thousands of individual

switches in approximately 184 RPCMs on the

station at assembly complete.  There are about

119 RPCMs on the station currently.

The European and Japanese laboratory

modules have their own internal power

distribution system.  Those modules will draw

power from DDCUs, from Node 2.  Their

unique transformers and power control

modules equivalent to U.S. RPCMs will handle

power.  NASA and Boeing have responsibility

for distributing power to those elements, but

the individual international partners will be

responsible for power within their respective

elements.

PRIMARY POWER DISTRIBUTION 

OVERVIEW 

Primary Power Distribution provides a

commandable interface between generated or

stored power to loads that are located down

stream.  Power distribution within a power

channel is performed by a DC Switching Unit

(DCSU) and the power distribution to loads is

performed by the MBSU.  At ISS assembly

complete, there will be eight DCSUs and four

MBSUs involved in primary power

distribution.  The DCSUs and MBSUs use a

December 2006 ELECTRIC POWER SYSTEM 36

network of high power switches called Remote

Bus Isolators (RBIs) to direct the power flow.  

The RBIs do not physically control the direction

of the current flowing through them but they

do provide a means of isolating a current path

in the event of a malfunction or if a repair is

needed on the primary power system.  The

RBIs in both the DCSU and MBSU are fully

commandable by on‐board computers.

Each power channel contains one DCSU to

perform power distribution on the Integrated

Equipment Assembly (IEA).  During insolation,

the DCSU routes power from the arrays to an

MBSU distribution bus, as well as to the BCDUs

for battery charging.  During eclipse, the DCSU

routes battery power to the same MBSU

distribution bus to satisfy power demands, and

it also sends a small amount of power back to

the SSU to keep the SSU firmware functioning

in preparation for the next insolation cycle.  In

addition to primary power distribution, the

DCSU has the additional responsibilities of

routing secondary power to components on the

PV modules (i.e., the Electronics Control Unit

and other support components).  This

secondary power is provided by the DDCU

located on the IEA.  The DDCU receives

primary power from the DCSU, converts it into

secondary power, and sends it to Remote

Power Controller Modules (RPCMs) for

distribution.  The PV module RPCMs are

housed within the DCSU.

The MBSUs act as the distribution hub for the

EPS system.  The four MBSUs onboard the ISS

are all located on the Starboard Zero (S0) truss.  

Each of MBSU receives primary power from

two power channels and distributes it

downstream to the DDCUs and other users

including Service Module (SM) American to

Russian Converter Units (ARCUs).  In addition,

the MBSUs can be used to crosstie power

channels (i.e., feed one power channel loads
December 2006 ELECTRIC POWER SYSTEM 37

with a different power channel source) to assist

in failure recovery and assembly tasks.

The BGAs and SARJs on the ISS also play a role

in primary power distribution.  The BGA

provides for the transmission of primary power

from the solar array wings to the IEA and the

SARJ provides for transmission of primary

power from the DCSUs to the MBSUs.  The

BGAs and SARJs incorporate a roll‐ring design

to provide conduits for power (and data), while

allowing a continuous 360° rotation.

SECONDARY POWER DISTRIBUTION 

The workhorse of the secondary power

distribution system is the RPCM, an Orbital

Replacement Unit (ORU), which contains solid‐

state or electromechanical switches, known as

Remote Power Controllers (RPCs).  RPCs can be

remotely commanded, by on‐board computers,

to control the flow of power through the

distribution network and to the users.  There

are different types of RPCMs, containing

varying numbers of RPCs and varying power

ratings.  As shown above, secondary power

flows from a DDCU and is then distributed

through a network of ORUs called Secondary

Power Distribution Assemblies (SPDA) or

Remote Power Distribution Assemblies

(RPDA).  Essentially, SPDAs and RPDAs are

housings that contain one or more RPCMs.  The

only distinction between SPDAs and RPDAs is

the location downstream of a DDCU.  RPDAs

are always fed from other RPCMs inside

SPDAs.  Note that RPCMs have only one power

input; thus, if power is lost at any level of the

Secondary Power System, all downstream user

loads will be without power.

There is no redundancy in the Secondary Power

System; rather, redundancy is a function of the

user’s loads.  For example, a critical user load

may be able to select between two input power

sources that use different power channels and

thus different secondary power paths.

As with DDCUs, SPDAs and RPDAs may be

located inside pressurized compartments or

outside.  Depending on their specific location,

SPDAs or RPDAs may interface with the Lab

Internal Thermal Control System (ITCS) or use

heat pipes to dissipate heat.  RPCMs are also

located within the DCSU on the IEAs to

provide secondary power‐to‐power channel

components, as required.
December 2006 ELECTRIC POWER SYSTEM 38

REDUNDANCY 

Each of the power channels is preconfigured to

supply power for particular ISS loads; however,

to provide a backup source of power for critical

equipment, the assembly complete design

provides for rerouting (i.e., cross‐tying)

primary power between various power

channels, as necessary.  At assembly complete,

the ISS will have four PV modules containing

eight power channels with full cross‐strapping

capability.  However, it is important to note

that only primary power can be cross‐strapped.  

Once power is converted into secondary power,

power flow through the distribution network

cannot be rerouted.

As a result, if there is a failure within the

Secondary Power System, there is no

redundancy, and the entire downstream path

from the failure is unpowered.  Instead, user

loads generally determine redundancy.  There

are three types of user redundancy schemes as

listed below:

• Components may be wired with multiple

power input sources, providing the

capability of swapping among them.

• Two or more components that perform the

same function can be fed by different power

sources; thus, the responsibilities of one

component can be assumed by another.

• Multiple components can work together to

perform a function; with the loss of a single

component, operational capabilities are

degraded not lost.

• Set of available jumpers that can be used to

temporarily regain power to a load until the

secondary system can be fixed.

SYSTEM PROTECTION 

The EPS is designed to protect equipment from

power surges and overheating at several points

along the power path from the source to the

users.  Current, voltage, and temperature

sensors are located in nearly all the EPS

equipment (ORUs) and are monitored by

firmware located on the hardware or on‐board

computers, or both.  If a voltage, current, or

temperature is out of range, an appropriate

safing action will be initiated either by the

firmware or by computer software.  The safing

action is designed to limit the amount of time

that the box is exposed to high power or high

temperature.  In case of power surges, it is also

designed to limit the impact of that surge on

other equipment along the power path.

The system protection function includes the

architecture’s ability to detect that a fault

condition has occurred, confine the fault to

prevent damaging connecting components, and

execute an appropriate recovery process to

restore functionality, if possible.  This process is

usually referred to as Fault Detection, Isolation,

and Recovery (FDIR).  For example, upon

detection of a fault, components can be isolated,

thereby preventing propagation of faults.  In

response to overcurrent conditions, the

architecture is designed such that each

downstream circuit protection device is set to a

lower current rating and responds more

quickly than the protection device directly

upstream.  This ensures that electrical faults or

shorts in the system do not propagate toward

the power source.  Another function of the

architecture’s system protection shuts down the

production of power when array output voltage

drops below a specified lower‐limit threshold.  

This prevents the PV cells from continuing to

feed a downstream fault.  In summary, all the

various implementations of system protection
work together to isolate faults or shorts at the

lowest level.  This approach minimizes impacts

to the users of the EPS and protects the EPS

from damage by low‐level faults.

KEY EPS COMPONENTS 

December 2006 ELECTRIC POWER SYSTEM 39

SOLAR ARRAY WING (SAW) 

The principle function of the SAW is to produce

electrical power from solar energy.  The SAW

contains 32,800 solar cells, 16,400 per blanket,

which can produce approximately 31 kilowatts

(kW) of electrical power at Beginning of Life

(BOL), and about 26 kW after 15 years, at their

designed End of Life (EOL).  However, it is

important to note that power availability is

influenced by ISS attitude, operational mode

(e.g., proximity operations), Sun alpha and beta

angle, shadowing, etc.

BETA GIMBAL ASSEMBLY (BGA) 

The function of the BGA is to provide minor

array pointing correction along the Beta Angle.   

The beta angle is the angle between the orbit

plane and the solar direction (changes

~±4°/day)) to compensate for apparent solar

motion induced by seasonal variations.  There

is one BGA associated with each SAW.  The

BGA provides one axis of rotation for a solar

array wing.  The BGA is capable of a full

360 degrees of rotation or may be commanded

to a specific location via computer command.  

Electric power generated by the solar array

wing is transferred through the BGA over the

entire range of BGA axis rotation.  The transfer

of power is accomplished by a rotary coupling,

the roll ring subassembly, which is mounted

coaxially with the axis bearing and torque

motor.

The BGA may be commanded to the following

modes of operation:

• Angle Command Mode.  BGA axis of

rotation aligned to a commanded angle

position.

• Latch Mode.  The BGA axis of rotation is

locked at specified location and prevented

from further rotation.

• Manual Operating Mode.  All non‐essential

functions are disabled and the drive motor

is disabled.  BGA axis may be rotated by

manual action from the IEA side.

• Rate Mode.  BGA may be commanded to

rotate at a specified rate.

ELECTRONICS CONTROL UNIT (ECU) 

The ECU is located on the BGA.  It is the

command and control link for the solar array

wing and BGA.  The ECU provides power and

control for extension and retraction of the solar

array mast, latching and unlatching of the

blanket boxes, BGA rotation, and BGA latching.

SOLAR ALPHA ROTARY JOINT 

The purpose of the SARJ is to rotate the PVMs

to provide alpha angle array pointing

capability.  The port SARJ and starboard SARJ

are located at the outboard end of the P3 and S3

truss segments and provide 360° continuous

rotational capability to the segments outboard

of P3 and S3.  The SARJ will normally complete

one complete 360‐degree revolution per orbit.  
The SARJ transfer electrical power through a

set of roll rings, which provide a continuous

rolling electrical connection while rotating.

INTEGRATED EQUIPMENT ASSEMBLY 

(IEA) 

Each IEA, located on P4, S4, P6 and S6, 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), and integrates the Thermal

Control Subsystem which consists of one

Photovoltaic Radiator (PVR) ORU and two

Pump Flow Control Subassembly (PFCS) ORUs

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  (and S3 and

S4 when launched next year) segments is

provided by the Alpha Joint Interface Structure.

December 2006 ELECTRIC POWER SYSTEM 40

The IEA measures 16 feet (4.9 meters) by 16 feet

(4.9 meters) by 16 feet ((4.9 meters), weighs

nearly 17,000 pounds a (7,711.1 kilograms) 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 DCSU used for primary power

distribution; the DDCU used to produce

regulated secondary power; the BCDU used

to control the charging and discharging of

the storage batteries; and the batteries used

to store power.

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.

3. The computers used to control the P4

module ORUs consist of two Photovoltaic

Controller Unit (PVCU)

Multiplexer/Demultiplexers (MDMs).

The IEA power system is divided into two

independent and identical channels.  Each

channel is capable of control (fine regulation),

storage and distribution of power to the ISS.  

The two power modules are attached outboard

of the AJIS.
December 2006 ELECTRIC POWER SYSTEM 41

SEQUENTIAL SHUT UNIT (SSU) 

The SSU is the primary power regulation device

that controls SAW output.  By design, the SSU

provides a consistent source of power (typically

~160 V‐dc), based upon a programmable

setpoint.  Regulation of the array output

voltage is required because array output

current, when illuminated with sunlight, is

often greater than the ISS demands.  To

accomplish this, the SSU receives power

directly from the PV array and maintains

output voltage to a setpoint, by shunting and

un‐shunting solar array strings.  Each string can

be individually connected or disconnected from

the primary bus and the power output from the

SSU is the sum of all connected strings at any

time.  When the SSU power output exceeds the

power demand, then the bus voltage starts to

rise and that triggers SSU to shunt strings to

reduce SSU power output, and vice versa.

The voltage setpoint is provided to the SSU by

the on‐board computer.  The setpoint is

designed to maximize array power capability

(maximum power point) while ensuring control

stability.  As solar arrays age, the voltage

setpoint is adjusted to ensure optimum

performance.

There will be eight SSU’s on orbit when the

station assembly is complete (two each per IEA,

four IEAs when assembly complete).  The SSU

is located on the beta gimbal platform, at the

bottom of the mast canister.

BATTERY SUBASSEMBLY ORBITAL 

REPLACEMENT UNIT 

The battery subassembly consists of 38

lightweight nickel hydrogen cells and

associated electrical and mechanical equipment,

packaged in an ORU enclosure.  When the sun

is behind the Earth, all of the power is provided

off the batteries, providing about a third of the

station’s power daily.  The battery interfaces

with a Battery Charge/Discharge Unit (BCDU),

which provides charge and discharge control of

electric energy.  During isolation, solar electric

energy, regulated by the SSU, will replenish

energy stores in preparation for the next eclipse

cycle.  Two battery ORUs makes a battery set.  

There will be 24 battery sets on ISS at assembly

complete.

The batteries have a design life of about seven

years, but the actual life obtained on orbit is a

function of how deeply the battery is

discharged and the number of charge‐discharge

cycles (nominally 16 cycles per day).
December 2006 ELECTRIC POWER SYSTEM 42

The battery ORUs can be changed out

robotically using a special purpose manipulator

on the end of the station’s robotic arm.   Each

battery measures 41 inches (104.1 cm) by

37 inches (94 cm) by 19 inches (48.3 cm) and

weighs 372 pounds (168.7 kilograms).

BATTERY CHARGE/DISCHARGE UNIT 

(BCDU) 

The BCDU serves a dual function of charging

the batteries during isolation and providing

conditioned battery power to the primary

power busses during eclipse.

The Control Power Remote Bus Isolator

(CPRBI) controls the flow of power to the DC

control power output bus and also functions as

a circuit breaker, limiting the load current

during faults.  The Fault Isolator (FI) limits the

battery discharge current, in the event of a fault,

to 85 to 127 amps.  The BCDU also includes

provisions for battery status monitoring and

protection from power circuit faults.

Each BCDU measures 28 inches (71.1 cm) by

40 inches (101.6 cm) by 12 inches (30.5 cm) and

weights 235 pounds (106.6 kilograms).  The

BCDU has an 8.4 kW battery charge capability

with a 6.6 kW discharge capability.  It provides

70 to 120 volts‐dc control power output and can

regulate power between 130 to 180 volts‐dc.  

The power storage system consists of a BCDU

and two Battery Subassembly ORUs.

MAIN BUS SWITCHING UNIT (MBSU) 

Located on the S0 truss, the four MBSUs

distribute primary power from the power

channels, downstream to the DDCUs, and other

loads.  They also provide the capability to cross

tie Primary Power Channels to feed those

DDCU loads in the event of a Primary Power

Failure.

Command, communication, health monitoring,

and RBI drive functions are provided by the

Switchgear Controller Assembly (SCA).  The

MBSUs have a design life of approximately

fifteen years.  There is a spare MBSU located on

orbit.

The system’s design can accommodate the loss

of PV modules and other problems by remotely

accessing the MBSUs, by either the ground or

on station, and internally redirecting power to

by‐pass faults or failures in the EPS.  The four

MBSUs themselves are not redundant.  All

MBSUs are required to power all station loads.  

However, MBSUs provide redundancy for

power modules upstream.  The MBSU output

voltage range is from 133 to 177 V‐dc.
December 2006 ELECTRIC POWER SYSTEM 43

The MBSU will be first used during

STS‐116/Assembly Mission 12A.1.  All four

MBSUs are activated during this mission.  Each

MBSU box is 28 inches by 40 inches by

12 inches and weighs 220 pounds.

DC SWITCHING UNIT (DCSU) 

The DCSU is the electrical distribution box for a

primary power channel.  It provides fault

protection for the many EPS ORUs, and

numerous EPS support functions.  The DCSUs

primary function is to route power between the

solar arrays, batteries, and downstream MBSUs

and DDCUs.

The DCSU is used for power distribution,

protection and fault isolation within the

Integrated Equipment Assembly.  The DCSU

uses remote‐controlled relays (RBIs) identical to

those on the MBSU to route primary power to

the BCDUs, MBSUs and DDCU.   The DCSU

also routes secondary power (124 V‐dc plus or

minus 1.5 V‐dc) through solid state switches to

the ECU, SSU, and PFCS ORUs.

There will be a total of eight DCSUs on the ISS

once S5/S6 is delivered in June 2008.  To date,

DCSUs have performed very well, with no

failures.  They have a design life of 15 years.  

Each DCSU box is 28 inches by 40 inches by

12 inches and weighs 238 pounds.

DIRECT CURRENT-TO-DIRECT 

CURRENT CONVERTER UNIT (DDCU) 

The secondary power conversion function uses

one type of ORU, the DDCU.  The DDCU

provides electrical isolation between the

primary and secondary EPS.  As the name

implies, the DDCU is responsible for dc power

conversion, in this case primary power into

secondary power, using a transformer.  Each

DDCU has one primary power input and one

secondary power output.  The DDCU converts

the coarsely regulated primary power (115 V‐dc

to 173 V‐dc) to a voltage‐regulated secondary

power (124.5 V‐dc nominal plus or minus

1.5 V‐dc).  The primary power on the ISS is like

the main transmission lines in a city with the

DDCU serving like a transformer on a utility

pole that converts the power so it can be used

in your home.  The primary power voltage is

typically 160 V‐dc; however, voltage can vary

over a wide range although the output is

specified to be 124 V‐dc, which is the prescribed

voltage for all users of the Secondary Power

System.  If any other voltage level is required

by user loads (such as payloads or crew

equipment), it is the user’s responsibility to

perform the conversion from 124 V‐dc to its

required voltage.
DDCU‐I (internal)

December 2006 ELECTRIC POWER SYSTEM 44

DDCUs come in three versions:

• DDCU‐E (for external).  These are used

outside the habitable spaces on the station

on the power module IEA, and Z1, S0, S1

and P1 trusses.  There are four DDCU‐Es on

S0, and one each on S1 and P1.

• DDCU‐I (for internal).  These are used

currently in the U.S. Laboratory and

eventually in the Node 2 modules when

those are delivered over the next several

years.  There are six DDCU‐Is in the U.S.

laboratory.

• DDCU‐HP, The heat‐pipe versions are

currently located on the Zenith One (Z1)

truss.  There are two DDCU‐HPs on Z1.

The only difference between the three types is

the nature of the cooling system that transfers

internal heat build up during operation to the

surrounding space environment.  The DDCU‐Is

are affixed to cold plates that transfer the heat

to a water‐cooled system inside the Lab and

Node modules.  This heat is then transferred by

heat exchangers to the external ammonia

cooling system where the heat is rejected into

space via the large central radiators on the

DDCU‐I (external)

external S1 and P1 trusses (which will become

active during this mission).  The DDCU‐Es are

cooled via cold plates which interface directly

to the external systems.  The DDCU‐HP are

cooled directly by the radiator heat‐pipe panel

on which they are mounted.

To date, DDCUs have performed very well,

with no failures.  They have a design life of

15 years.  There are already spare DDCUs (with

the exception of DDCU‐HP) on orbit.  Each

DDCU‐E is 27 inches by 23 inches by 12 inches

and weighs 129 pounds.  Each DDCU‐I is

27 inches by 18 inches by 10 inches and weighs

112 pounds.  Each DDCU‐HP is 68 inches by

25 inches by 14 inches and weighs 200 pounds.
December 2006 ELECTRIC POWER SYSTEM 45

REMOTE POWER CONTROLLER MODULE 

(RPCM) 

RPCMs are the interface between the EPS and

all non‐EPS equipment onboard the ISS.  The

RPCM is a multi‐channel, high power circuit

breaker.  The RPCM is the workhorse of the

secondary power system.  The RPCM has the

following two purposes:

• To control the distribution of secondary

power to downstream loads by opening or

closing RPCs

• To protect the EPS against downstream

faults by opening RPCs when overcurrents

are sensed

Six types of RPCMs facilitate system protection,

fault isolation and power flow control in the ISS

Electric Power Systems.  There are

approximately 119 RPCMs currently located on

the station and there will be a total of 184 when

assembly is complete.  There are multiple spare

RPCM located on orbit.  Several failed RPCMs

have been replaced on the ISS.

RPCMs are located in the following modules:  

Integrated Equipment Assembly (Two each in

DCSU or a total of 16), P1 (12), P3 (8), S1 (12), S3

(8), S0 (24), Z1 (4), Node 1 (12), Node 2 (18),

Node 3 (27), Quest Airlock (4) and U.S. Destiny

Lab (39).  Each RPCM is 6.8 inches by 8 inches

by 3.5 inches and weigh 10.5 pounds.

ELECTRICAL POWER SYSTEM (EPS) 

RECONFIGURATION

The space station will change from its early

configuration to the beginning of the final

assembly complete configuration.  Since the P6

module has been on top of Z1, its power has

been going directly to the DDCUs in the lab

and Z1.  Four MBSUs have been sitting on the

S0 truss, but have not been getting any input

power and are not feeding anything

downstream.  They have been sitting on a non‐

operating cooling loop.  A series of bypass

jumper cables and reconfigurable connections

were initially installed that have allowed P6 to

directly feed the DDCUs from the DCSUs.  The

initial configuration essentially bypassed the

MBSUs.  During STS‐116, power will flow

through the MBSUs for the first time since they

have been on orbit, so that the main power

from the trusses comes into the MBSUs first

before going to the DDCUs.  Each DCSU will

feed one side of an MBSU.

To complete the reconfiguration, approximately

112 Extravehicular Activity (EVA) power

connectors will be removed and reattached.  

Hundreds of commands from mission control

will be issued to power up and down various

components.  This mission includes the largest

number of EVA power connectors ever

removed in a single assembly mission.
December 2006 ELECTRIC POWER SYSTEM 46

Two EVAs will be conducted to reconfigure the

main power channels or paths.  The purpose of

the reconfigurations is to route power through

the truss MBSUs and DDCUs.  There are two

channel domains over which electrical power

travels across the truss and station elements

called the 1 and 4, and 2 and 3 power domains,

which comprises the basic numbering system

for the power paths or domains.   The four

MBSUs are numbered 1 through 4.  All the

hardware downstream has a similar numbering

system so you know which MBSU it came from.  

For example, when mission controllers refer to

the 2/3 domain, they are referring to all the

loads that go through MBSUs 2 and 3.  In

understanding the station power system,

anything that receives a power feed from the

2/3 domain will have a “2/3” numbering

scheme.  Two separate thermal cooling loops on

the ISS cool the 1/4 and 2/3 hardware.   Each

SAW on the IEA is labeled as A or B, so that

each power channel has a power domain and a

solar array designation.  For example, the two

power channels on the P6 module are 2B and

4B, and those on the P4 module are 2A and 4A.  

Similarly the channels on S4 are 1A and 3A and

S6 are 1B and 3B.

During the first EVA during STS‐116/Assembly

Mission 12A.1, the P5 truss will be installed on

the outboard side of P4.  On the next day,

NASA will remotely retract the 4B solar array,

since there is a structural interference between

the P6 4B array and the P4 arrays if rotated on

the Solar Alpha Rotary Joint.  Solar Array 4B on

P6 will be retracted into its blanket boxes and

the SARJ will be commanded into a solar

tracking mode.  P4 is not currently powering

the rest of the space station, but will be brought

on line during this mission.

During the second EVA during Flight Day 6,

the entire 2/3 domain will be brought down for

several hours while various power connectors

are rewired.  The remaining 1/4 path will

support all the loads on the ISS, which means

there will be no redundant power for the

heaters and most of the equipment and systems

(communications, guidance, navigation and

control, etc.) throughout the station.  They will

power everything back up once the 2/3 domain

is configured properly, and the MBSUs are

brought on line.  The MBSUs were checked on

earlier spacewalks in 2002.  NASA has

provisions in place in case an MBSU fails when

reconfigured.   There will be an entire day,

beginning when the crew begins their sleep

shift following the first EVA, in which mission

controllers will begin their procedures to

transfer all the station systems from 1/4 to the

2/3 domain power path.  Powering the DDCUs

down, many of the systems that are being

powered have been fed by DDCUs in the lab up

to now, but after this reconfiguration, many of

the DDCUs from the trusses will come alive for

the first time.  Two MBSUs and three truss

DDCUs will be activated.  On EVA 2, there are

15 umbilicals connected via EVA, and 6 Intra‐

Vehicular Activity (IVA).  IVA activities consist

of jumpers being connected inside the Lab,

Node, and Airlock by the crew to provide

backup power sources during the

reconfigurations.  There will be about

73 mates/demates of power connectors.

Two days later, a third EVA will be conducted

on Flight Day 8 to reconfigure the 1/4 domain.  

For EVA 3, 20 umbilicals will be connected

during the EVA and 9 IVAs, with a total of

81 mates/demates.  Like on the earlier EVA,

everything will run off the 2/3 domain power

path while various connectors are removed and

mission control powers items up and down.  

Two MBSUs and three truss DDCUs will be

activated.  Like the 2/3 domain reconfiguration,

there will be limited power redundancy for
several hours.  The astronauts will do all of the

reconfigurations on the 1/4 side while

everything is powered down and then once all

the connections are complete, then all four

MBSUs will be on line and both cooling loops

will be running and all the power is brought

back up.  Once NASA starts powering up after

each 2/3 and 1/4 reconfiguration, they will

know right away if any of the 2/3 or 1/4

hardware is not working.  As each item is

powered back up, the electronics will conduct

an internal self‐checking procedure.  All of this

data goes into mission control.  There is a

considerable amount of cleanup following an

EVA to get all the systems back running and in

a nominal configuration.  It takes the entire

night and next day following each EVA to

verify those systems are operating nominally.

There will also be some minor reconfigurations

on the Russian side when power converters are

December 2006 ELECTRIC POWER SYSTEM 47

moved over to the U.S. primary side to where

they are getting a higher voltage input.  It is a

seamless change and there are no changes in

the major configurations on how the U.S. and

Russians sides share power.

NASA conducted a large number of electrical

power system simulations to prepare for this

mission and is confident that electrical

reconfiguration will go as planned.  NASA has

a number of contingency procedures since this

reconfiguration operation is not without risks,

especially if a problem should occur with the

ammonia pumps for the cooling system that

will be brought on line for the first time.  

Boeing and other industry engineers will be

assisting NASA in the ISS Mission Evaluation

Room, one of the primary backroom support

centers to mission control, should any issues or

problems arise.
December 2006 ELECTRIC POWER SYSTEM 48

EQUIPMENT REFERENCE

Image shows locations of the MBSUs and

DDCUs on the SO truss element as part  

of the Electrical Power System

Image shows locations of DDCUs and MBSUs

on the SO truss element to be activated

Location of DDCU to be activated on S1 truss Location of DDCU to be activated on P1
December 2006 ELECTRIC POWER SYSTEM 49

EPS and other systems components on SO truss
December 2006 ACTIVE THERMAL CONTROL SYSTEM 50

ACTIVE THERMAL CONTROL SYSTEM 

(ATCS) OVERVIEW 

Most of the stationʹs many systems produce

waste heat, which needs to be transferred from

the ISS to space to achieve thermal control and

maintain components at acceptable

temperatures.  An Active Thermal Control

System (ATCS) is required to achieve this heat

rejection function when the combination of the

ISS external environment and the generated

heat loads exceeds the capabilities of the

Passive Thermal Control System to maintain

temperatures.  An ATCS uses a mechanically

pumped fluid in closed‐loop circuits to perform

three functions:  heat collection, heat

transportation, and heat rejection.  Waste heat

is removed in two ways, through cold plates

and heat exchangers, both of which are cooled

by a circulating ammonia loops on the outside

of the station.  The heated ammonia circulates

through large radiators located on the exterior

of the space station, releasing the heat by

radiation to space that cools the ammonia as it

flows through the radiators.

The ATCS consists of the Internal Active

Thermal Control System (IATCS), External

Active Thermal Control System (EATCS), the

Photovoltaic Thermal Control System (PVTCS)

and the Early External Active Thermal Control

System (EEATCS).  The IATCS consists of loops

that circulate water through the interior of the

U.S. Destiny Laboratory module to collect the

excess heat from electronic and experiment

equipment and distributes this heat to the

Interface Heat Exchangers for transfer to the

EATCS.  At assembly complete, there will be

nine separate ITCS water loops in the U.S. and

International Partner pressurized modules.

The Photovoltaic Thermal Control System

(PVTCS) consists of ammonia loops that collect

excess heat from the Electrical Power System

(EPS) components in the Integrated Equipment

Assembly (IEA) on P4 and eventually S4 and

transport this heat to the PV radiators (located

on P4, P6, S4 and S6) where it is rejected to

space.  The PVTCS consist of ammonia coolant,

eleven coldplates, two Pump Flow Control

Subassemblies (PFCS) and one Photovoltaic

Radiator (PVR).

Active Thermal Control System Architecture
The External Active Thermal Control System

(EATCS), activated for the first time on this

mission, consists of ammonia loops to collect

heat from the Interface Heat Exchangers and

external electronic equipment mounted on

coldplates and transports it to the S1 and P1

radiators where it is rejected to space.  In lieu of

using the EATCS initially, the station hardware

has been cooled by the Early External Active

Thermal Control System (EEATCS).  The

EEATCS has provided heat rejection capability

rejection capability for the U.S. Laboratory

Interface Heat Exchangers (IFHX) since STS‐98

through STS‐116.

The EEATCS is the temporary system used to

collect, transport, and reject waste heat from

habitable volumes on the International Space

Station (ISS).  The EEATCS collects heat from

the Interface Heat Exchangers (IFHX) located

on the U.S. Laboratory module, circulates the

working fluid, anhydrous ammonia, via the

Pump and Flow Control Subassembly (PFCS),

and rejects heat to space via two orthogonally

oriented stationary radiators.

INTERNAL ACTIVE THERMAL CONTROL 

SYSTEM (IATCS) 

The purpose of the U.S. Destiny Laboratory

ITCS is to maintain equipment within an

allowable temperature range by collecting,

transporting, and rejecting waste heat.  The

ITCS uses water because it is an efficient

thermal transport fluid and is safe inside a

habitable module.  The IATCS is a closed loop

system that provides a constant coolant supply

to equipment, payloads and avionics to

maintain proper temperature.

The U.S. Laboratory contains two independent

loops, a Low Temperature Loop (LTL) and a

Moderate Temperature Loop (MTL).  This

approach allows for segregation of the heat

December 2006 ACTIVE THERMAL CONTROL SYSTEM 51

loads, simplifies heat load management, and

provides redundancy in case of equipment

failure.  The LTL is designed to operate at 40° F

(4° C) and service systems equipment requiring

low temperatures, such as the Environmental

Control and Life Support System (ECLSS)

Common Cabin Air Assembly (CCAA) and

some payload experiments.  The LTL contains

approximately 16.64 gallons (63 liters) of fluid.  

The MTL nominally operates at 63° F (17° C)

and provides most of the cooling for systems

equipment (i.e. avionics) and payload

experiments.  The MTL contains approximately

52.83 gallons (200 liters) of fluid.  The IATCS

loops can be configured and operated as a

single loop.  This capability is used for a variety

of purposes, including the reduction of wear on

the pumps, reduction of pump power usage, or

to compensate for a pump failure.

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 can dissipate

6,000 watts of heat per orbit on average and is

commanded by the IEA computer.  Each PFCS
December 2006 ACTIVE THERMAL CONTROL SYSTEM 52

consumes 275 watts during normal operations

and measures approximately 40 inches

(101.6 cm) by 29 inches (73.7 cm) by 19 inches

(48.3 cm), weighing 235 pounds (106.7

kilograms).

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 (740.7 kilograms) and

when deployed measures 10.24 feet

(3.12 meters) by 44.62 feet (13.6 meters).  When

the station assembly is complete, there will be a

total of four PVRs, one for each PV module (S4,

P4, P6, S6).

EARLY EXTERNAL ACTIVE THERMAL 

CONTROL SYSTEM (EEATCS) 

Function 

Since the U.S. Laboratory became operational

before the permanent External Active Thermal

Control System (EATCS) was assembled, a

temporary external cooling system was needed.  

External cooling from the Russian segment is

not possible because there are no operational

interfaces between the U.S. On‐orbit Segment

(USOS) and the Russian On‐orbit Segment

(ROS) thermal systems.  Instead, a modified

version of the Photovoltaic Thermal Control

System (PVTCS) called the Early External

Active Thermal Control System (EEATCS) acts

as a temporary thermal system.  The EEATCS

consists of two independent, simultaneously

operating ammonia cooling loops (ACL).  These

loops transport heat loads from the Interface

Heat Exchanger (IFHX) located on the

Laboratory moduleʹs aft endcone to the

radiators located on truss segment P6.

The EEATCS is needed until the permanent

EATCS is activated.  Once the permanent

EATCS becomes operational on mission, the

EEATCS will be deactivated.  After

deactivation, portions of the EEATCS will be

used as spare components for the PVTCS loops.

Hardware 

Pump & Flow Control System (PFCS)

Each external loop contains a Pump & Flow

Control System (PFCS) which contains most of

the controls and mechanical systems that drive

the EEATCS.  There are two pumps per PFCS

which circulate ammonia throughout the

external coolant loops and a Flow Control

Valve (FCV) which mixes cold radiator flow

and warm IFHX return flow to regulate the

temperature of the ammonia in the loop.  The

PFCS also contains the primary ammonia

accumulator, which provides limited ammonia

leakage makeup, protection against thermal

expansion of the ammonia, and a net positive

suction head greater than the minimum

required to prevent pump cavitations.  

Additionally, all manner of pressure,

temperature, flow, and quantity sensors used

by the EEATCS are part of the PFCS.

Radiators

The EEATCS radiator ORU is a direct flow,

deployable and retractable radiator system with

two independent cooling loops.  The EEATCS

radiator consists of seven radiator panels, the

deploy/retract mechanism, support structure,

and the necessary plumbing.  The EEATCS

radiator has two channels (A & B) that acquire

heat from the Lab Low Temperature (LT) and

Moderate Temperature (MT) Loop Interface

Heat Exchanger (IFHX) via liquid anhydrous

ammonia.  The ammonia flows from the PFCS

to the associated IFHX, to the EEATCS radiator
December 2006 ACTIVE THERMAL CONTROL SYSTEM 53

manifold tubes, across the radiator panels and

back to the PFCS.  The radiator panels reject the

excess heat to space via two non‐articulating

EEATCS radiator ORUs:  one AFT (Trailing)

and one Starboard (Normal).  The two radiator

ORUs are located on the P6 Long Spacer Truss

Segment.  The radiator measures 10.24 feet

(3.12 meters) by 44.62 feet (13.6 meters).

Interface Heat Exchanger (IFHX)

The Interface Heat Exchanger (IFHX) units

accomplish heat transfer from the IATCS water

coolant loops to the external ammonia coolant

loops.  When the station is complete,

10 interface heat exchangers will be in

operation to provide heat transfer from the

IATCS loops of the various habitable modules

to the two external ammonia coolant loops.  

The IFHX units are located on the U.S.

Laboratory, Node 2, and Node 3.

EXTERNAL ACTIVE THERMAL CONTROL 

SYSTEM (EATCS) OVERVIEW 

The EATCS provides heat rejection capabilities

for all U.S. pressurized modules and the main

power distribution electronics on S0, S1 and P1.  

The system uses a single‐phase anhydrous

ammonia as its working fluid for its high

thermal capacity and wide range of operating

temperatures.  Ammonia has an extremely low

freezing point of ‐107 degrees ° F (‐77 °C) at

standard atmospheric pressure.  The EATCS is

comprised of two independent loops labeled

loop A on S1 (Starboard) and Loop B on P1

(Port).  The independent loops were designed so

that a failure in one would not take down the

entire EATCS system.  Both loops are physically

separated to prevent orbital debris from taking

out the lines and the fluid transport lines are

buried within the truss structure.  If a loop does

go down, the EATCS operates at a reduced

capacity.  Each loop collects heat from up to five

Interface Heat Exchangers (IFHXs) mounted on

the Node 2, U.S. Destiny Laboratory, and Node 3

as well as externally mounted coldplates.   Most

of the cold plates and plumbing to the

pressurized modules are located on the S0 center

truss.  The EATCS is designed to provide 35 kW

of heat rejection per loop for a total capability of

70 kW.  The EATCS also provides ammonia re‐

supply capability to the Photovoltaic Thermal

Control Systems (PVTCS) located on P4, P6, S4

and S6.  All EATCS components are located

outside the pressurized volumes to prevent crew

contact with ammonia.

At assembly complete, each ammonia loop will

supply coolant to five Interface Heat Exchangers

(IFHX) and five cold plates (three Direct

Current‐to‐Direct Current Units (DDCUs) and

two Main Bus Switch Units (MBSUs).  Two

MBSU cold plates, each designed to remove

495 watts at 80 lbs/hr.  Three DDCU cold plates

are each designed to remove 694 watts at

125 lbs/hr.  The cold plate interfaces with the

component base‐plate via radiant fins.  IFHXs

transfer thermal energy from the Internal

Thermal Control System’s (ITCS) water based

coolant to the ETCS anhydrous ammonia

coolant.  Ammonia supply temperature is

currently set at 37 °F (2.8 °C)
December 2006 ACTIVE THERMAL CONTROL SYSTEM 54

The ITCS supply temperature varies as a

function of the modules’ thermal load.  IFHX

can isolate and bypass the IFHX core on the

ammonia side in the event a cold slug is

detected at the pump outlet to prevent ITCS

coolant from freezing.

Key Components

The External Active Thermal Control System

(EATCS) is the primary permanent active heat

rejection system on ISS.  It acquires, transports,

and rejects excess heat from all U.S. and

International Partner modules except the

Russian modules.  The EATCS contains two

ammonia coolant loops, which cool equipment

on the S0, S1, and P1 truss segments.  Capable

of rejecting up to 70kW, the EATCS provides a

substantial upgrade in heat rejection capacity

from the 14kW capability of the Early External

Active Thermal Control System (EEATCS).

Heat Acquisition Subsystem (HAS) 

The HAS consists of the Interface Heat

Exchanger (IFHX) Orbital Replacement Units

(ORU), Main Bus Switch Unit (MBSU) and

DC‐to‐DC Converter Unit (DDCU) cold plates

ORU.

Heat Rejection Subsystem (HRS) 

The HRS consists of the radiator ORU, which is

a deployable, eight‐panel system that rejects

thermal energy via radiation.  The HRS also

consists of the Radiator Beam Valve Module

(RBVM) that provides radiator isolating or

venting, radiator beam which carries three

radiators and connects to the Thermal Radiator

Rotary Joint (TRRJ), which rotates to the

radiator beam to provide radiator articulation.  

The EATCS allows the flow of ammonia

through heat rejection radiators that constantly

rotate to optimize cooling for the station.
December 2006 ACTIVE THERMAL CONTROL SYSTEM 55

Interface Heat Exchanger (IFHX) 

The Interface Heat Exchanger (IFHXs) provides

the interface between a module’s internal TCS

and the EATCS.  The IFHXs transfer heat from

the internal loops of the USOS modules to the

EATCS ammonia loops.  IFHXs are used to

collect heat from USOS modules.  There are five

IFHXs for each EATCS loop.  Some IFHXs are

plumbed in series such that the cool ammonia

flows through a module’s Low Temperature

Loop (LTL) IFHX prior to flowing though

another module’s Moderate Temperature Loop

(MTL) IFHX.

The IFHX units accomplish heat transfer from

the IATCS water coolant loops to the external

ammonia coolant loops.  Each IFHX core

utilizes a counterflow design with 45

alternating layers.  IATCS water flows through

23 of the layers, while EATCS ammonia flows

through the 22 alternate layers in the opposite

direction.  These alternating layers of relatively

warm water and relatively cold ammonia help

to maximize the heat transfer between the two

fluids via conduction and convection.  The heat

exchanger core is a simple flow through device

with no command or telemetry capability.  

IFHXs are mounted on the Node 2, U.S.

Laboratory, and Node 3.  The U.S. Laboratory

IFHXs have been connected to the EEATCS,

until this flight, when the EEATCS ammonia

fluid line quick‐disconnect will be disconnected

and reconnected to the EATCS.  When Node 2

arrives on STS‐120, it is equipped with six

IFHXs designed to provide cooling for itself, the

Columbus and Japanese Experiment Module.  

Node 3 also contains a set of IFHXs, which are

connected to the EATCS when it arrives on a

shuttle flight currently set for launch no earlier

than January 2010.

When the station is complete, 10 interface heat

exchangers will be in operation to provide heat

transfer from the IATCS loops of the various

habitable modules to the two external ammonia

coolant loops.  The IFHX units will be located

on the U.S. Laboratory, Node 2, and Node 3.  

Because of the highly toxic nature of ammonia,

IFHX ORUs are mounted external to the

pressurized modules as a safety precaution.  

Each IFHX measures 25 inches (63.50 cm) by 21

inches (53.34 cm) by 8 inches (20.32 cm) and

weighs about 91 pounds (41.28 kilograms).

Interface Heat Exchanger (IFHX) provide the  

interface between the ITCS and the EATCS
Heaters 

Each EATCS loop has electrically powered

heaters wrapped around the supply and return

fluid lines on the S0 Truss to maintain the

minimum operating temperature.  These

heaters are used during low heat load

conditions and are turned on and off by

software in the Multiplexer/Demultiplexer

(MDMs).  These heaters can be operated in

closed‐loop mode (temperature based) or open‐

loop mode (time based).  Numerous heaters are

located on the EATCS plumbing on the S1

(Loop A) and P1 (Loop B) truss segments to

prevent ammonia freezing and flexible hose

damage during nonoperational periods.  These

heaters are thermostatically controlled and

have no software interface.

Cold Plate 

Each ETCS loop provides cooling to externally

mounted coldplates located on the S0, S1 (Loop

A), and P1 (Loop B) truss segment.  These

coldplates contain Electrical Power System

(EPS) equipment used to convert and distribute

power to downstream ISS loads.  Each

ammonia loop contains four coldplates, two

attached to Direct Current‐to‐Direct Current

Converter Units (DDCUs) and two attached to

Main Bus Switching Units (MBSUs).

December 2006 ACTIVE THERMAL CONTROL SYSTEM 56

Main Bus Switching Unit (MBSU) Coldplate

Direct Current‐to‐Direct Current Converter

Units (DDCU) Cold Plate

Each MBSU coldplate measures 37 inches

(93.98 cm) by 33 inches (83.8 cm) by 20 inches

(50.8 cm) and weighs about 109 pounds

(49.4 kilograms).

Each coldplate ORU is connected to the EATCS

ammonia loop by self‐sealing quick disconnect

(QD) couplings and contains a finned coldplate,

two or three strip heaters and temperature

sensor.  The coldplates are installed such that

the fins of the coldplate are positioned adjacent

to corresponding fins on either the DDCU or

the MBSU to facilitate heat transfer by radiation

between the cooled equipment and the

coldplate.  Each DDCU coldplate measures

35 inches (88.9 cm) by 28 inches (71.12 cm) by

31 inches (78.74 cm) inches and weighs about

96 pounds (43.54 kilograms).

Pump Module (PM) 

Circulation, loop pressurization, and

temperature control of the ammonia is

provided by the Pump Module (PM).  Each

ammonia loop contains a Pump Module

Assembly (PM) ORU to provide flow and

accumulator functions and maintains proper

temperature control at the pump outlet.  Each
December 2006 ACTIVE THERMAL CONTROL SYSTEM 57

PM consists of a single pump, a fixed charge

accumulator, a Pump & Control Valve Package

(PCVP) containing a firmware controller,

startup heaters, isolation valves, and various

sensors for monitoring performance.  The

accumulator within the PM works in concert

with the Ammonia Tank Assembly (ATA) tanks

to compensate for expansion and contraction of

ammonia caused by the temperature changes

and keeps the ammonia in the liquid phase via

a fixed charge of pressurized nitrogen gas on

the backside of its bellows.

The Pump Module (PM) provides fluid

pumping, fluid temperature control and system

pressure control.  The PCVP provides flow

control.  A single pump in the PCVP provides

circulation of the ammonia.  The Flow Control

Valve (FCV) located within the PCVP regulates

the temperature of the ammonia.  The FCV

mixes “cool” ammonia exiting the radiators

with “warm” ammonia that has bypassed the

radiators.

Nominally, loop A will operate at 8,200 lb/hr

and loop B at 8,900 lb/hr at 14,000 and 14,700

revolutions per minute, respectively.

The Pump Module ORU circulates liquid

ammonia at a constant flowrate to a network

of coldplates and heat exchangers located on

the external trusses and U.S. modules,

respectively.

For STS‐116, initial activation with U.S.

Laboratory IFHX where Loop A pump will run

at 11,500 rpm, equivalent to 5,000 lb/hr while

Loop B pump will run at 11,500 rpm which is

equivalent to 5,200 lb/hr.

The accumulator located in the PM provides

auxiliary pressure control.  The accumulator

resides upstream of the PCVP in each PM ORU.  

The accumulator keeps the ammonia in the

liquid phase by maintaining the pressure above

the vapor pressure of ammonia and provides

makeup ammonia in case of a leak.  The

accumulator works in conjunction with the

ATA to absorb fluctuations in the fluid volume

due to varying heat loads through the

expansion and contraction of its internal

bellows.

Nominal operating pressure for the loops is

300 psia at the pump inlet; the pressure will be

brought up to 390 psia for start up.  The

maximum system design pressure is 500 psia.

Each PM measures 69 inches (175.26 cm) by

50 inches (127 cm) by 36 inches (.91 cm) inches

and weighs about 780 pounds (353.8 kilograms)

Low and High Pressure Flow Control 

Monitoring 

Failure Detection, Isolation and Recovery

(FDIR) for high and low pressure conditions are

monitored and issued by the S1/P1

Multiplexer/Demultiplexers (MDMs).  For an

over pressure, gaseous nitrogen pressure is

relieved down to 360 psia when pump inlet

pressure reaches 415 psia (active control).  The

PVCP Inlet pressure, Radiator return pressure,

and Bypass return pressure sensors are part of

this system and two of three pressure readings

are used to determine if an overpressure

condition exists.  The pump will shut down

issued when the pump outlet pressure reaches

480 psia (active control).  Various relief valves
and burst disks at the IFHX, PM, and RBVM

will relieve at approximately 70 psia (passive

control)

Low pressure (current limit set at 170 psia) is

monitored by two methods to determine a low

pressure condition (chooses higher of the two

values to determine the limit).  Low pressure

conditions are monitored using the PVCP inlet

pressure, radiator return pressure, and bypass

return pressure sensors.

Temperature Control 

The PCVP also maintains temperature set point

control of the ammonia supplied to the HAS.  

The PCVP has a temperature control capability

of 36 °F (2.2  °C) to 43 °F (6.1  °C) and it will be

set at 37 °F ± 2 °F (2.8  °C).  The temperature

control method is by three way mixing valve

that mixes flow from the radiators and the HRS

Bypass.  Heaters on the HRS Bypass leg provide

an additional level of control.  Heaters are used

to provide fluid conditioning in the event the

thermal load on the loop is not sufficient to

maintain set point control and to support

temporary transient events.  Total heater power

of 1.8 kW is split across two heater strips

mounted on the HRS bypass lines (900 watts

each).

Pump outlet over temperature protection is

provided by a Firmware Controller (FWC) in

the PCVP that uses three PCVP outlet sensors

to determine an over temp condition and issues

zero pump speed.  The S1/P1 MDMs use the

PM outlet sensor to determine an over temp

condition and pull power from the Solenoid

Driver Output (SDO) card providing power to

the PM.

Current limit is set at 65 °F (18.33 °C).  Freeze

Protection in the IFHX is detected by the PCVP

firmware which shuts down the pump

December 2006 ACTIVE THERMAL CONTROL SYSTEM 58

(first leg).  When an under temperature

condition is detected by the S1/P1 MDMs, it

will pull power from the SDO card providing

power to the PM (second leg).  Under

temperature detected by the S0 1,2 MDMs pulls

power from the utility rail (third leg, leaves

many things without power).  The current limit

is set at 35 °F (1.67 °C).

Fluid Supply 

Each ammonia loop contains an ATA ORU to

contain the heat transfer fluid (liquid ammonia)

used by the EATCS loops.  There is one ATA

per loop located on the zenith side of the S1

(Loop A) and P1 (Loop B) truss segments.  The

ATA ORU will be used to fill the EATCS loop

on startup, to supply makeup fluid to the

system, to act as an accumulator in concert with

the PM accumulator and provide the capability

to vent the ammonia loops by way of a

connection to an external non‐propulsive vent.  

Each ATA primarily consists of two bellows

ammonia tanks pressurized by an external

nitrogen source, two internal survival heaters

and two sets of quantity, differential pressure,

absolute pressure and temperature sensors.  

The ATAs are isolatable and replaceable on

orbit.

Ammonia re‐supply capability for the  

EATCS and the eight PVTCS located  

on P6, P4, S4 and S6 is provided by  

the Ammonia Tank Assembly (ATA).
December 2006 ACTIVE THERMAL CONTROL SYSTEM 59

Multilayer Insulation (MLI) applied to the

exterior surfaces of the ORU is provided to

guard against excessive heat loss.  The ATA

ORU is protected against Micro‐

Meteoroid/Orbital Debris (MM/OD) by

shielding on the exterior of each tank and the

ORU itself.  Each ATA measures 79 inches

(200.66 cm) by 46 inches (116.84 cm) by 55

inches (139.7 cm) inches and weighs about

1,120 pounds (508.02 kilograms).

The ATA in combination with the Nitrogen

Tank Assembly (NTA) provide fluid supply

and primary system pressure control.  A single

ATA was launched on Flights 9A and 11A

(ITS S1 and P1) with approximately 640 lbm

ammonia in each ATA, 320 lbm per tank.  ATA

provides necessary plumbing connection to the

ammonia vent system via the vent panel.  

Supply to outboard trusses is provided through

the vent panel.  The ATA acts as the primary

accumulator for the EATCS in concert with the

NTA.  If required, it can also be used to

replenish the PVTCS fluid lines.

Each ammonia loop contains a NTA ORU to

provide storage for the high pressure nitrogen

used for controlled pressurization of the ATA.   

The NTA mounts to the S1 (Loop A) and P1

(Loop B) truss segments and is connected to the

ATA by self‐sealing QDs.  Each NTA ORU

primarily consists of a nitrogen tank, a gas

pressure regulating valve (GPRV), isolation

valves and survival heaters.  The nitrogen tank

provides a storage volume for the

high‐pressure gaseous nitrogen, while the

GPRV provides a pressure control function as

well as nitrogen isolation and over pressure

protection of downstream components.  The

NTA provides the necessary pressure to move

the ammonia out of the ATA.  The single high‐

pressure tank containing nitrogen at 2,500 psia

(@70 °F, ground fill) and uses the GPRV to

supply continuous pressure up to 390 psia in

one psia increments.  A back‐up mechanical

valve limits the maximum nitrogen pressure to

416 psia.  The GPRV provides pressure control

as well as high‐pressure nitrogen isolation and

overpressure protection of downstream

components.  The NTA has venting capabilities

and over pressure controls.  Each NTA

measures 64 inches (162.56 cm) by 36 inches

(91.44) by 30 inches (76.2 cm) inches and weighs

about 460 pounds (208.65 kilograms).
Fluid Lines and Quick Disconnects (QD) 

Fluid lines and external QDs provide the

transportation path from the truss segments to

the IFHXs.  Connections between segments are

made with flex hoses and QDs.  There are flex

hoses and QDs between each truss, and

between the S0 truss and the various IFHXs.

Heat Rejection System 

Heat collected by the EATCS ammonia loops is

radiated to space by two sets of rotating

radiator wings—each composed of three

separate radiator ORUs.  Each radiator ORU is

composed of eight panels, squib units, squib

unit firmware controller, Integrated Motor

Controller Assemblies (IMCAs),

instrumentation, and QDs.  Each Radiator ORU

measures 76.4 feet (23.3 meters) by 11.2 feet

(3.4 meters) and weighs 2,475 pounds

(1,122.64 kilograms)

Each ammonia loop contains one radiator wing

comprised of three Radiator ORUs mounted on

the Radiator Beam and six Radiator Beam Valve

Modules (RBVM) and one Thermal Radiator

Rotary Joint (TRRJ).  The Radiator ORUs utilize

anhydrous ammonia to reject heat from the

EATCS.

December 2006 ACTIVE THERMAL CONTROL SYSTEM 60

Heat Rejection System (HRS) Radiator during

deployment testing at Lockheed Martin

Missiles and Fire Control.

Each Radiator ORU contains a deployment

mechanism and eight radiator panels.  The

deployment mechanism allows the Radiator

ORU to be launched in a stowed configuration

and deployed on orbit.  Each radiator ORU can

be remotely deployed and retracted.

Each individual radiator has two separate

coolant flow paths.  Each flow path flows

through all eight radiator panels.  Each panel’s

flow path has eleven flow tubes for a total of

22 Inconel flow tubes or passages (11 passages

per flow path) per radiator panel; flow tubes

are freeze tolerant.  Flow tubes are connected

along the edge of each panel by manifolds.  Flex

hoses connect the manifold tubes between

panels.  Each panel has a white (Z‐93) coating

which provides optimum thermo‐optical

properties to maximize heat rejection.  Flow

tube arrangement is designed to minimize

ammonia freezing in the radiator.

Each radiator path contains one Radiator Beam

Valve Module (RBVM) as a part of the radiator

wing.  Six RBVMs are mounted on the radiator

beams on the S1 and P1 truss segments.  Two

RBVMs service each radiator ORU.  Each

RBVM consists of an isolation relief valve, an

isolation valve, an Integrated Motor Controller

Assembly (IMCA), QDs, and pressure and
December 2006 ACTIVE THERMAL CONTROL SYSTEM 61

temperature sensors.  The RBVM controls the

transfer of ammonia between the Radiator

Assembly ORU and the rest of the EATCS loop.  

Each RBVM contains sensors to monitor

absolute pressure, temperature and valve

position within the ORU.  Remote control

venting of the radiator fluid loop is also

available through the RBVM to facilitate

radiator replacement and prevent freezing of

the ATCS coolant during contingency

operations.  The RBVM provides flow path

isolation in the event that a panel suffers

micro‐meteoroid damage.  Leak isolation FDIR

functions are controlled by the S1/P1 MDMs

monitoring large leaks via the STR/PTR MDMs.  

Additionally, the RBVM provides automatic

pressure relief when the EATCS is over

pressurized.  Each RBVM weighs about

50 pounds (22.68 kilograms) and measures

24 inches (60.96 cm) by 20 inches (50.8 cm) x 5.4

inches (13.72 cm).

The rotation capability for each radiator

assembly is provided through a Thermal

Radiator Rotary Joint (TRRJ).  The TRRJ

provides power, data, and liquid ammonia

transfer to the rotating radiator beam while

providing structural support for the radiator

panels.  Each TRRJ is composed of the

following:  a bearing assembly, two Rotary Joint

Motor Controllers (RJMCs), two Drive Lock

Assemblies (DLAs), a Flex Hose Rotary Coupler

(FHRC), and a Power and Data Transfer

Assembly (PDTA).  The bearing assembly is the

rotary interface and primary structural

component of the TRRJ.  The RJMCs provide

control for the DLA system, which provides

joint rotation and joint locking capability.  The

FHRC consists of four flex hoses, two supply

and two return.  The PDTA provides the data

and power paths for transfer to and from the

radiator beam.

TRRJ ORU provides rotation capability to the

Radiator Beam to optimize the thermal

environment of the radiators and to maximize

heat rejection capability and prevent freezing in

the radiator manifolds.  Rotation angles are

determined via the Radiator Goal Angle

Calculation (RGAC) algorithm which

commands the Radiator Beam to put the

radiators either “edge to the sun” during

isolation phase of the orbit or “face to the

Earth” during the eclipse phase.  The RGAC

ensures the radiators stay cold enough so the

heat can be rejected but warm enough so that

the ammonia does not freeze.  There is a

temperature goal of –40 °F at the radiator

outlet.  The FHRC provides the transfer of

liquid ammonia across the rotary joint and is

capable of rotating 230 degrees, at ±115 degrees

from its neutral position. (software command

limit is ±105°); with a variable rotation speed of

There are two RBVMs (one per flow path) that

allow or prevent the transfer of ammonia to

and from the radiator panels.
Thermal Radiator Rotary Joint (TRRJ) provides controlled rotation of the EATCS radiators, allows

the transfer of power, data, and ammonia across the rotating interface, and provides the structural

support between the S1/P1 truss segments and the associated radiator wing assembly.

December 2006 ACTIVE THERMAL CONTROL SYSTEM 62

0 to 45 degrees‐per‐minute.  Each TRRJ

measures approximately 5.6 feet (1.7 meters) by

4.6 feet (1.4 meters) by 4.3 feet (1.3 meters) and

weighs 927 pounds (420.5 kilograms).

System Performance Overview 

Loop A and B operate at slightly different flow

rates mainly due to differing system hydraulic

resistance layout.  Actual heat rejection will

need to be planned and coordinated between

all modules so as not to exceed the EATCS total

heat rejection capability of 70 kW (at assembly

complete), including S1/P1/S0 mounted

electrical equipment.

Software 

Thermal Control System (TCS) software is used

to control and monitor the system.  The TCS

software executes actions such as system

startup, loop reconfiguration, and valve

positioning for flow and temperature control

automatically or via commands from crew

laptops or ground workstations.  Telemetry

from the various temperature, pressure, flow,

and quantity sensors is monitored by TCS

software and displayed on crew laptops or

ground workstations.  In addition, Fault

Detection, Isolation, and Recovery (FDIR)

software is used to monitor the performance of

the TCS and, if there is a problem, alert the

crew and flight controllers.  In some cases,

FDIR software initiates recovery actions.

EATCS Activation 

With the solar arrays attached by the shuttle

Atlantisʹ crew in September, NASA is ready

with STS‐116 to start generating enough power

so that the permanent cooling system can be

brought on line.  NASA will have to bring up

the permanent electrical power distribution

system on line first before activating EATCS.  

The goal is to power up the MBSUs, route

power through them to the cooling system and

gets that system activated before the MBSUs

can overheat.  The EATCS has never been

tested as an entire integrated system in its on‐

orbit configuration.  However, NASA and

Boeing engineers did conduct an entire loop

test, but it did not have a rotating Radiator.

During STS‐116/Assembly flight 12A.1, during

the second and third spacewalks, the EATCS is
December 2006 ACTIVE THERMAL CONTROL SYSTEM 63

activated as part of the entire reconfiguration.  

There are two major power domains, 1/4 and

2/3.  During EVA 2, the 2/3 power domain

reconfiguration takes place and loop B is

activated and then on EVA 3, the 1/4 power

domain reconfiguration takes place and loop A

is activated.  Both power‐on and power‐off are

major reconfigurations, when half the station

power is turned off for periods of time.  NASA

has implemented a large number of procedures

to account for the situation where they are zero‐

fault tolerate.  The key to each EVA is activation

of the pump modules and getting the system

started for the first time.  Assuming a nominal

activation, there are approximately 24

procedure steps to activate the EATCS for the

first time.

Preparations to activate the EATCS start when

the ground filled nitrogen pad is vented from

the center radiators and the rest of the system

(to be vented after STS‐116 reaches orbit).  The

system is presently pressured with about

80 psia of nitrogen.  The next step will be to

introduce ammonia into the system with only

one radiator ORU per loop plumbed, which

will take several hours.  Filling and activating

the system will be one of the major challenges

that mission controllers and engineers will face.  

Activation will take about an hour for each

loop.  If there were an anomaly during filling

such as pressure begins to drop like a leak, then

the entire reconfiguration would be

interrupted.  Mission controllers have

thoroughly rehearsed their procedures to

identify any potential leaks in the system.  

Some minor leaks are possible.  Early activation

will provide cooling to the MBSUs and DDCUs

(their respective cold plates) on S1, S0 and P1.  

The activation sequence also allows for a

through checkout of the EATCS loops prior to

connecting the U.S. Destiny Laboratory IFHX to

the EATCS.  The US Laboratory IFHXs will

continue to be cooled via the Early External

Active Thermal Control System (EEATCS)

located on P6 Long Spacer.

During the 12A.1 stage, EVAs will be

conducted to connect the U.S. Laboratory

IFHXs to the EATCS.  After the IFHXs are

connected to the EATCS, the EEATCS will enter

a dormancy phase.

On STS‐120, also known as Assembly Mission

10A, the two remaining radiator ORUs per wing

are deployed and filled.  Node 2 end‐cones are

connected to the EATCS via the starboard and

port, boom trays located on the forward end of

S0.  Pump shut down will be required.  All six

IFHXs on Node 2 receive cooling from the

EATCS:  two for Node 2, two for Attached

Pressurized Module (APM‐Columbus), and two

for Japanese Experiment Module (JEM‐Kibo).  

The APM and JEM IFHXs remain in a bypassed

and isolated configuration to prevent accidental

freezing of the water side core of the IFHXs.

On Assembly Flight 20A, currently set for

launch no earlier than January 2010, the Node 3

end‐cone is connected to the EATCS via jumpers

from the US Laboratory aft end‐cone.  A bypass

line with an isolation valve on the US Laboratory

Aft end‐cone allows the fluid circuit to be

completed prior to Node 3 arrival.  Upon Node 3

arrival, the end‐cone connection and activation

the bypass‐Isolation valves on US Lab end‐cone

are closed.  This mission completes the EATCS

loop architecture.

Boeing engineers in Huntington Beach, Calif.

designed the EATCS as well as S0, S1 and P1

which contains most of the EATCS hardware.  

Major subcontractors to Boeing were Hamilton

Sundstrand (PCVP, coldplates, core for heat

exchanger), Honeywell (tanks, accumulator,

RBVMs), Lockheed Martin (radiators) and

Marotta (valves).
December 2006 ACTIVE THERMAL CONTROL SYSTEM 64

COMPONENT REFERENCE DRAWINGS 

Component locations on S1/P1

Component locations on S1/P1

December 2006 SPACEWALKS 65

SPACEWALKS 

The major objectives of the three spacewalks on

the STS‐116 mission include:  install the Port 5

(P5) truss segment; reconfigure the space

station power system from its temporary to its

permanent configuration and incorporate

power from the P4 solar arrays for the first

time:  activate ammonia loops for cooling;

relocate the Photovoltaic Radiator Grapple

Fixture (PVRGF) and the Crew Equipment

Translation Aid (CETA) carts; and transfer the

Service Module Debris Panels (SMDPs).

The square‐shaped P5 truss is about the length

of a small compact car.  It will provide

structural spacing and utility connections

between the P4 and P6 solar arrays after the

P6 solar arrays are relocated during a shuttle

mission next year.  The station eventually will

have 11 integrated truss segments joined

together to stretch 356 feet end to end to

support four solar array assemblies and

radiators to power and cool the station.

Astronauts Robert L. Curbeam, Jr. and Christer Fuglesang, STS‐116 mission specialists, wearing

training versions of the Extravehicular Mobility Unit (EMU) spacesuit, participate in an  

underwater simulation of extravehicular activity (EVA).  Curbeam and Fuglesang are  

dwarfed by station truss segments in this overall view of the simulation conducted  

in the Neutral Buoyancy Laboratory (NBL) near the Johnson Space Center.
December 2006 SPACEWALKS 66

During the three spacewalks, the astronauts

and Mission Control Houston will power down

equipment, transfer power to other redundant

power channels and unplug and plug in

electrical connectors.

The extravehicular activities (EVAs) are

planned for flight days 4, 6 and 8, with each

spacewalk estimated to last 6.5 hours.  There

will be several challenges during the

spacewalks.  On flight day 4, as the P5 spacer is

being moved into position, it will come as close

as 2.7 inches from the P4 truss in a move similar

to parallel parking in a snug space.  On flight

day 6 and again on flight day 8, Mission

Control, Houston, will power down half of the

station as the crew prepares to rearrange the

power system from the temporary system it has

used to its permanent configuration.  The

process will activate several pieces of

equipment for the first time.  It also will allow

the station to use power generated by the P4

solar arrays, added to the complex during

STS‐115 in September 2006, for the first time.

Mission specialists Bob Curbeam and Christer

Fuglesang will conduct the first and second

spacewalks.  Curbeam and Mission Specialist

Suni Williams will conduct the third spacewalk.  

Pilot Bill Oefelein will be the intravehicular lead

for the spacewalks, assisting the spacewalkers

from inside the spacecraft with their tasks.  

Williams will operate the station’s robotic arm

during the first two spacewalks and mission

specialist Joan Higginbotham will operate the

stationʹs arm during the third spacewalk.  

Mission Specialist Nicholas Patrick will operate

the shuttleʹs robotic arm.

Curbeam is a veteran of three previous

spacewalks.  This will be the first spaceflight

and spacewalks for Fuglesang and Williams.  

Williams, who joins the station crew for a

Astronaut Bob Curbeam

six‐month stay, is slated to perform three

spacewalks during her station flight.

On spacewalks 1 and 2, Curbeam will wear a

spacesuit marked with red stripes around the

legs and Fuglesang will wear an all‐white

spacesuit.  On spacewalk 3, Curbeam will wear

the spacesuit marked with red stripes, while

Williams will wear all white.  All of the

spacewalks will originate from the station’s

Quest airlock.

EVA 1 

There are three major objectives during EVA 1:  

First, install the P5 spacer; second, remove and

stow the Photovoltaic Radiator Grapple

Fixture—a handle used by the shuttle and

station robotic arms during the P5 installation;

and third, remove and replace a camera on the

S1 truss that is needed to view clearances

during future truss segment installations.

On flight day 3, the crew will use the shuttle’s

robotic arm to lift the P5 truss from the shuttleʹs

payload bay and hand it to the station’s robotic

arm.  It will remain ʺparkedʺ at the end of the

station arm while the crew sleeps that night.  

Then, on flight day 4, the station’s arm—the

Space Station Remote Manipulator System or

SSRMS—will be used to move the P5 spacer to

a pre‐installation position on the left, or port,
side of P4.  Once the P5 is in position, Curbeam

and Fuglesang will remove the locks that

secured the spacer’s hardware during its launch

on the shuttle.  Curbeam will remove the P5

launch locks from corners 3 and 1.  Fuglesang

will remove locks from corners 2 and 4.

Once the locks are removed, the two will

provide verbally guide Williams as she

maneuvers the station arm to align P5 to P4.  

The installation is completed by mating six

utility cables.

Before the P5 is locked in place, Mission

Control must take several measures to ensure

nothing disturbs the operation.  The station’s

thrusters will be turned off and the control

moment gyroscopes will be in a mode that

creates as little disturbance as possible until the

crew has tightened three of the four Modified

Rocketdyne Truss Attachment System

(MRTAS) bolts.  Mission Control also will

configure the Solar Alpha Rotary Joint and Beta

Gimbal Assemblies—used to swivel the solar

arrays to allow them to track the sun—into the

proper positions for P5 installation.  All of the

joints will be locked to ensure there is no

unexpected motion.

After the P5 launch locks are removed, the

station robotic arm will guide P5 into the soft

capture position, with Curbeam and Fuglesang

assisting by monitoring structural clearances.  

Once the P5 coarse alignment cone has

captured the P4 soft capture pin, the

spacewalkers will use a pistol grip tool—similar

to a portable hand drill—to tighten the

attachment bolts.  First the bolts on corner 1

(forward/zenith) and 2 (forward/nadir) will be

tightened, in that order, to an initial torque.  

Next, bolts on corners 3 (aft/zenith) and 4

(aft/nadir) will be driven to final torque.  

Finally, bolts 1 and 2 will be tightened to their

final torque.  After the bolts are driven, the

December 2006 SPACEWALKS 67

crew will attach grounding straps on each

corner and remove the soft capture pins.

Next, the crew will move the Photovoltaic

Radiator Grapple Fixture (PVRGF) from the P5

to a place on the station’s Mobile Base System.  

The grapple fixture will be temporarily stored

there until it can be moved to the P6 aft radiator

on a later spacewalk.  The PVRGF is a handle

that was installed on the top, or zenith, side of

the P5 spacer and used by the shuttle and

station arms to move the truss.  It is being

removed to provide enough clearance for the

solar arrays to rotate and track the sun.

To accomplish this, Curbeam and Fuglesang

will move to the grapple fixture, where

Curbeam will use the pistol grip tool to begin

removing fasteners.  Meanwhile, Fuglesang will

move a foot restraint to a new worksite

interface.  Once the foot restraint is in place,

Fuglesang will begin removing grapple fixture

fasteners.  Once all the fasteners are removed,

Fuglesang and Curbeam will perform an almost

acrobatic transfer of the tethered grapple fixture

from one to the other as they move to the

station’s keel.

The final task for this spacewalk is the removal

and replacement of a camera located at port 3 of

the station’s starboard 1 (S1) truss.  The camera

is needed to view clearances during future

installations and deployments.  Curbeam will

set up the worksite while Fuglesang retrieves

the replacement camera from the airlock.  Both

will work to remove the failed camera and

install the new one.

EVA 2 

During EVA 2, there are two primary tasks and

two secondary tasks for the crew.  The first

primary task is to reconfigure space station
December 2006 SPACEWALKS 68

power to the permanent architecture and

integrate P4 power beginning with the channel

2/3 side of the station’s Electrical Power System.  

The second is to relocate two crew and

equipment translation aid (CETA) carts from

the starboard side of the station to its port side.  

This will clear the way for the station’s Mobile

Transporter to move to worksite 2 on the

starboard 1 truss.  Worksite 2 must be checked

out prior to the STS‐117 mission next year.  On

that mission, it will be used during installation

of the starboard 3 and 4 (S3/S4) truss segments.

During the spacewalk, the crew also will install

a thermal blanket for the force moment sensors

on the station’s robotic armʹs latching end

effector.

First, Curbeam and Fuglesang will reconfigure

power on the Channel 2/3 side of the station’s

Electrical Power System to route it through the

Main Bus Switching Units.  The power

reconfiguration requires that all power be shut

down from Channel 2/3.  To accomplish this, a

lengthy set of power down procedures will be

executed from the ground while the crew

prepares for the spacewalk.

The Channel 2/3 power reconfiguration is

expected to take Curbeam and Fuglesang about

an hour and a half to complete.  During the

spacewalk, primary power from channels 2A

and 2B will be connected to the main bus

switching unit 2.  The work will be done at the

Starboard 0 (S0) truss segment near the forward

S0‐Lab struts.  Curbeam will translate to the S0

to disconnect the lab’s secondary power, install

a cable to route power from the S0 truss to the

P1 truss, and to reconfigure the S0‐2B dc

converter unit, the bypass jumper on the main

bus switching, and the P1‐3A dc converter.

Astronaut Christer Fuglesang

The S0 truss segment sits in the middle position

on the truss structure on top of the U.S. Destiny

Laboratory, flanked by the S1 and P1 truss

elements.  That truss, along with the S1 and P1

trusses, contain the major electrical components

of the permanent electrical system, including

the main bus switching units and the DC‐to‐DC

converter units.

The power produced by the station’s solar

arrays is routed to batteries for storage and then

to the main bus switching units.  The main

buses route power to DC‐to‐DC converter units

(DDCUs) which then adjust the primary,

160‐volt dc electricity it receives from the main

buses to about 125 volts of power.  The dc

converters feed the station through the Remote

Power Controller Modules (RPCMs).  A

simplified comparison is that the main buses

are similar to a power substation, the dc

converters are like transformers, and the

controller modules are like the electrical

switches inside a home.

While Curbeam is configuring power to the

Unit 2 main bus, Fuglesang will be removing

Circuit Interrupt Devices (CIDS) 3, 4, and 5.  

Once the main bus is operating, the circuit

interrupt devices won’t be needed.  The circuit

interrupts served as early “circuit breakers” for
the crew.  They will be returned to Earth.  All

three circuit interrupters are located at the lab’s

aft end cone panel, in an area known as “the

rat’s nest.”

Because the main bus switching unit generates

heat once it is activated, a pump module must

also be activated to enable ammonia to flow

through large cold plate loops that cool the bus.  

Prior to both the second and the third

spacewalks, the station’s External Thermal

Control System (ETCS) ammonia loops will be

filled so they are ready when the buses are

activated.  However, if the pump module

doesn’t work—it is powered by the bus—there

is an estimated five‐to six‐hour window before

the bus must be powered off and allowed to

cool.  If the pump isn’t working because it

failed, it will need to be removed and replaced.  

This procedure will require the crew to remove

and replace the pump on the third spacewalk

instead of completing the power

reconfiguration to the P4 truss.  That could

result in an additional spacewalk to complete

the station’s power reconfiguration.  The crew

has trained for that possibility.

Another consideration is the possibility that the

bus or a dc converter will fail and will need to

be removed and replaced.

Once the EV crewmembers have finished the

Channel 2/3 power reconfiguration, they will

begin other tasks while Mission Control powers

up the station.  Curbeam and Fuglesang will

relocate two crew and equipment translation

aid (CETA) carts from their current locations on

the S1 truss to the S0 truss to clear the way for

Atlantis’ astronauts to install the S3 and S4

truss segments on the STS‐117 mission.

The crew also will install a thermal cover on the

force moment sensors on the latching end

effector on the station’s robotic arm and

December 2006 SPACEWALKS 69

reconfigure power to the Z1 truss electrical

patch panel 6, which provides power to the Z1

truss as well as the Russian segment.

The electrical panel task must be conducted

separately from the main power down because

the S‐band communication antennas powered

by Channel 2/3 are sensitive to cold and should

not be shut down for a protracted period of

time.

While Curbeam is busy with this

reconfiguration, Fuglesang will retrieve the

starboard and port quick disconnect bags—

stowage bags filled with maintenance hardware

and tools—from the airlock and install the bags

on top, or the zenith side, of the airlock.

EVA 3 

The primary objective of the third spacewalk is

to reconfigure power on the Channel 1/4 side of

the station’s Electrical Power System.  The crew

also will stow three Service Module Debris

Panel bundles on a grapple fixture on the

Pressurized Mating Adapter 3—a module

attached to the port side of Node 1; reconfigure

power on the Z1 truss patch panels 1 and 6; and

install an Adjustable Grapple Bar (AGB) on the

flex hose rotary coupler on the airlock’s

External Stowage Platform‐2.

The third spacewalk on flight day 8 is basically

a repeat of the flight day 6 power

reconfiguration activities.  During EVA 3, the

Channel 1/4 side of the station’s Electrical

Power System will be reconfigured to route

primary power through the Main Bus

Switching Units.  As on flight day 6, this

spacewalk requires that all power be shut

down—this time from Channel 1/4.  Again, a

lengthy set of power down procedures will be

executed from the ground while the crew

prepares for the walk.  The reconfiguration task
December 2006 SPACEWALKS 70

is also predicted to take about an hour and a

half.

The work will again be at the Starboard 0 (S0)

truss segment.  Curbeam will reconfigure the S0

forward starboard avionics and the S1‐4B and

S0‐1A dc converters, disconnect the secondary

power Lab 4A dc converters, route the S0/N1

power cable, and reconfigure the bus bypass

jumper.

While he is disconnecting and connecting

cables, Williams will remove the 1 and 2 circuit

interrupt devices and reconfigure the other half

of the Z1 truss electrical patch panel 5.  The

circuit interrupters will be returned on the

shuttle.  Also, during the Z1 patch panel task,

the crew will venture out to the Russian

segment interface at Node 1 and reconfigure

the Russian power feeds from the U.S. segment.  

This reconfiguration will allow the Russians to

draw additional power from the U.S. segment

by moving some of their power feeds to larger

switches, capable of transferring more current,

should that be needed.

As Curbeam and Williams reconfigure external

power connectors, the crew inside will

reconfigure an electrical patch panel in the

Destiny lab.  This new configuration will

provide twice as much power for payload use.

Once power is reconfigured for Channel 1/4,

Curbeam and Williams will move to the

shuttle’s payload bay to attach three Service

Module Debris Panel bundles (SMDPs) on to an

adapter.  The three bundles and adapter were

launched on the Integrated Cargo Carrier (ICC),

a pallet located at the rear of the payload bay.  

Once the bundles are attached to the adapter,

the assembly is referred to as “the Christmas

tree.”

Astronaut Sunita L. Williams prior to being

submerged in the waters of the NBL.  

Astronaut William A. (Bill) Oefelein,  

STS‐116 pilot, assisted Williams.

Williams will work from the shuttle arm to

move the “Christmas tree” to stow it on the

grapple fixture located on the aft side of the

Pressurized Mating Adapter 3.  The bundles are

composed of individual panels that will be

installed on the Service Module to provide

additional micrometeroid debris protection.  

The panels will be moved to the Service

Module during a Russian EVA currently slated

for the summer of 2007.

While Williams completes clean‐up of the task,

Curbeam will move to reconfigure power to the

Z1 truss electrical patch panel 1.

During the third spacewalk, Curbeam and

Williams will also transfer the Adjustable

Grapple Bar (AGB)—a portable handle that can

be installed on objects to make it easier for the

crew to move them around during

spacewalks—to the Flex Hose Rotary Coupler

(FHRC) on the airlock’s External Stowage

Platform‐2, where it will be stowed for future

use.
December 2006 PAYLOAD OVERVIEW 71

PAYLOAD OVERVIEW 

INTEGRATED TRUSS SEGMENT P5 

STS‐116 will deliver the square‐shaped Port 5

(P5) truss segment to the International Space

Station.  P5 is part of the 11‐segment integrated

truss structure and the sixth truss element to be

delivered.  The truss structure forms the

backbone of the International Space Station

with mountings for unpressurized logistics

carriers, radiators, solar arrays, other hardware

and the various elements.  Port 5 will be

attached to the Port 4 truss element via the

Modified Rocketdyne Truss Attachment System

(MRTAS) interface.  P5 is used primarily to

connect power and cooling lines and serve as a

spacer between the P4 photovoltaic module

(PVM), or solar battery, and P6 PVM, which

will be joined during a later assembly mission.  

P5 is very similar in construction to the long

spacer located on P6.  Without the P5 short

spacer, the P4 and P6 solar arrays would not be

able to connect due to the way the photovoltaic

arrays (PVA) are deployed on orbit.

The P5 Short Spacer is shown in the Space Station Processing Facility at Kennedy Space Center.
The girder‐like structure is made of mostly

aluminum and provides several extravehicular

aids, robotic interfaces, ammonia servicing

hardware (as part of the station’s External

Active Thermal Control System that allows

ammonia fluid to transfer from P4 to P6) and

can also accommodate an external storage

platform.  The Enhanced Universal Trunnion

Attachment System (EUTAS) allows platforms

to be attached to P5 for the storage of additional

science payloads or spare Orbital Replacement

Units.  P5 also has white thermal blankets on

the structure, which help shade the P4 Solar

Array Assembly ORUs.

December 2006 PAYLOAD OVERVIEW 72

P5 Specifications 

Dimensions: Length is 132.813 inches 

or 11 feet and 0.813 

inches (3.37 meters) 

Width is 179.014 inches

or 14 feet, 11 inches 

(4.55 meters) 

Height is 167.031 inches

or 13 feet and 11 inches

(4.24 meters) 

Weight: 4,110 lbs or 2 tons 

(1,864.26 kg) 

Cost $10,972,000
December 2006 PAYLOAD OVERVIEW 73

P5 is transferred using the shuttle arm from the

payload bay to the Space Station Robotic

Manipulator System (SSRMS), the stationʹs

robotic arm, where it will be placed into the

install or soft‐dock position.  While being

moved on the SSRMS, P5 will have about three

inches of clearance as it passes the P4

Sequential Shunt Unit (SSU).  The truss element

is installed robotically with a crew assist.  

During the first spacewalk, astronauts will use

the truss attachment system to connect P5 to P4

by using their portable hand tools to drive in

four 3/4 inch diameter primary bolts in each

corner.  If a primary bolt cannot be secured, two

contingency bolts at each corner on P5 can be

tightened into the nut assemblies on P4.

Another feature of P5 is the Photo Voltaic

Radiator Grapple Feature (PVRGF).  For

launch, the PVRGF is stowed on top of P5 and

is used by the shuttle and station robotic arms

to grab P5 to lift if from the shuttle cargo bay

and attach it to the station.  After P5 is attached

to P4, the PVRGF will be relocated to the trussʹ 

keel during the first spacewalk using four

fasteners.

P5 also contains a remote sensor box, two

tri‐axial accelerators and two antenna

assemblies as part of the External Wireless

Instrumentation System (EWIS).  EWIS will

give engineers a better understanding of the

actual response of the truss system on orbit to

vibration and other stresses and help engineers

predict the fatigue life and durability of the

truss structure.

Boeing’s Rocketdyne Power and Propulsion

(now Pratt and Whitney) designed P5.  The

component was constructed in Tulsa, Okla., in

2000.  P5 arrived at Kennedy Space Center July

19, 2001, for final manufacture, acceptance and

checkout.  Boeing will continue to provide

sustaining engineering of P5 and for the entire

310‐foot integrated truss assembly.

Drawing shows the Photo Voltaic Radiator

Grapple Feature (PVRGF) attached to  

a folded photovoltaic radiator (PVR)
December 2006 PAYLOAD OVERVIEW 74

Sequence showing removal of P5 from the payload bay

and transferred to the space station’s robotic arm.

Sequence showing the movement and installation of P5  

to the soft‐dock position at the end of P4
December 2006 PAYLOAD OVERVIEW 75

P5 shown installed on the ISS
December 2006 PAYLOAD OVERVIEW 76
December 2006 PAYLOAD OVERVIEW 77

INTEGRATED CARGO CARRIER 

(SPACEHAB) 

The Integrated Cargo Carrier can carry up to

8,000 pounds of cargo on both faces of a pallet

measuring 8 feet long, 15 feet wide and 10

inches thick.  It carries cargo for transfer from

the shuttle to the station and from the station to

work sites on the truss assemblies.  The carrier

is an unpressurized, flatbed pallet and keel‐

yoke assembly housed in the orbiterʹs payload

bay.  It has no active interfaces (thermal,

electrical or data) with the shuttle.

On STS‐116, the ICC will carry the Service

Module Debris Panels (SMDP), Space Test

Program – H2 (STP‐H2), 15 Adjustable Mass

Plates (AMPs) and an International Space

Station Passive Flight Releasable Attachment

Mechanism (ICC PFRAM).  The STP‐H2 is a

payload which deploys three satellites after the

shuttle undocks from the space station.  The

ICC PFRAM is designed for a contingency

return of a Pump Module Integrated Assembly

(PMIA) should a pump module fail during the

flightʹs second spacewalk.  The SMDPs are

transferred to space station and are temporarily

stowed on PMA3.
December 2006 PAYLOAD OVERVIEW 78

LOGISTICS SINGLE MODULE 

(SPACEHAB LSM) 

In Discoveryʹs payload pay, the small

pressurized module known as the SPACEHAB

Logistics Single Module provides cargo launch

and return transportation.  Some of the

hardware stowed in the module on STS‐116 are

a Video Baseband Signal Processor (VBSP), a

Rotary Joint Motor Controller (RJMC)

Assembly, an External TV Camera Group

(ETVCG), Oxygen Generation System (OGS),

Adjustable Grapple Bar, Remote Power Control

Module(s) (RPCM), Nickel Removal Assembly

(NIRA) Kit, Charcoal Bed Assembly (CBA),

Desiccant/Adsorbent ORU, Control Moment

Gyro (CMG) Electrical Assembly (EA), an

Avionics Air Assembly (AAA) and payloads.

Bulkhead

T-Beams

Bulkhead-to-Bulkhead

Horizontal Tie Rods

Rack Support

Structure

Bulkhead-to-Top Panel

Diagonal Tie Rods

Rack Overhead

Structure

Flex Section

Interface Ring

Forward

Bulkhead

Cylinder

D-Ring

EVA

Handrail

EVA

Handrail NPRV

Opening

Viewport

Opening

Keel

Trunnion

(Not Shown)

Aft Bulkhead
December 2006 EXPERIMENTS 79

EXPERIMENTS 

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.

Such experiments assigned to STS‐116 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. Pre‐launch:  Ensure planning will allow

selection of a runway with Microwave

Scanning Beam Landing System support,

which 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.

SHORT-DURATION BIOASTRONAUTICS 

INVESTIGATION (SDBI) 

Short‐Duration Bioastronautics Investigations

(SDBIs) are shuttle‐based, life science payloads,

experiments and technology demonstrations.

SDBI 1503S Test of Midodrine as a 

Countermeasure against Postflight 

Orthostatic Hypotension 

Presently, there are no medications or treatment

to eliminate orthostatic hypotension, a

condition that often affects astronauts following

spaceflight.  Orthostatic hypotension is a

sudden fall in blood pressure that occurs when

a person assumes a standing position.  

Symptoms, which generally occur after sudden

standing, include dizziness, lightheadedness,

blurred vision and a temporary loss of

consciousness.

Space alters cardiovascular function, and

orthostatic hypotension is one of the alterations

that negatively impacts crew safety.  

Susceptibility to orthostatic hypotension is

individual, with some astronauts experiencing

severe symptoms, while others are less affected.  

This countermeasure evaluation proposal,

sponsored by the Countermeasures Evaluation

and Validation Project, is in its second phase of

the evaluation of midodrine.  It is designed to

give the greatest opportunity of measuring the

maximum efficacy of the drug.

This experiment will measure the effectiveness

of midodrine in reducing the incidence and, or,

the severity of orthostatic hypotension in

returning astronauts.  Its effectiveness will be

evaluated with an expanded tilt test.
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.  It 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 

Subjects will don the Actilight watch 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 Actilight 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.

The experiment examines the effects of

spaceflight on the sleep‐wake cycles of

astronauts during mission.  This information

could be vital in treating insomnia on Earth and

in space.

December 2006 EXPERIMENTS 80

SHORT-DURATION RESEARCH AND 

STATION EXPERIMENTS 

Short-duration Research to Be 

Completed during STS-116 

Incidence of Latent Virus Shielding During

Spaceflight (Latent Virus) will determine the

frequencies of reactivation of latent viruses—

inactive viruses in the body that can be

reactivated, such as cold sores—and clinical

diseases after exposure to the physical,

physiological, and psychological stressors

associated with spaceflight.  Understanding

latent virus reactivation may be critical to crew

health during extended space missions as crew

members live and work in a closed

environment.

Maui Analysis of Upper Atmospheric

Injections (MAUI) will observe the exhaust

plume of the Space Shuttle will lead to

assessment of spacecraft plume interactions

with the upper atmosphere.

Effect of Space Flight on Microbial Gene

Expression and Virulence (Microbe) will

investigate the effects of the space flight

environment on virulence (ability to infect) of

three model microbial pathogens:  Salmonella

typhimurium, Pseudomonas aeruginosa, and

Candida albicans, that have been identified as

potential threats to crew health based upon

previous space flight missions.

Test of Midodrine as a Countermeasure

Against Post‐Flight Orthostatic Hypotension

(Midodrine) measures the ability of the drug

midodrine, as a countermeasure, to reduce the

incidence or severity of orthostatic

hypotension—dizziness caused by the blood‐

pressure decrease that many astronauts

experience upon returning to the Earthʹs

gravity.
Perceptual Motor Deficits in Space (PMDIS)

will investigate why shuttle astronauts

experience difficulty with hand‐eye

coordination while on orbit.  This experiment

will measure the decline of astronauts’ hand‐

eye coordination during space shuttle missions.  

These measurements will be used to distinguish

between three possible explanations:  the brain

not adapting to the near weightlessness of

space; the difficulty of performing fine

movements when floating in space; and stress

due to factors such as space sickness and sleep

deprivation.

Ram Burn Observations (RAMBO) is an

experiment in which the Department of

Defense uses a satellite to observe space shuttle

orbital maneuvering system engine burns.  Its

purpose is to improve plume models, which

predict the direction the plume, or rising

column of exhaust, will move as the shuttle

maneuvers on orbit.  Understanding the

direction in which the spacecraft engine plume,

or exhaust flows could be significant to the safe

arrival and departure of spacecraft on current

and future exploration missions.

Sleep‐Wake Actigraphy and Light Exposure

During Spaceflight ‐ Short (Sleep‐Short) will

examine the effects of spaceflight on the sleep‐

wake cycles of the astronauts during space

shuttle missions. Advancing state‐of‐the‐art

technology for monitoring, diagnosing and

assessing treatment of sleep patterns is vital to

treating insomnia on Earth and in space.

Space Test Program –H2 

Atmospheric Neutral Density Experiment

(ANDE) consists of two microsatellites

launched from the Shuttle payload bay will

measure the density and composition of the

December 2006 EXPERIMENTS 81

low Earth orbit (LEO) atmosphere while being

tracked from the ground.  The data will be used

to better predict the movement of objects in

orbit.

Microelectromechanical System‐Based

(MEMS) PICOSAT Inspector (MEPSI) will

demonstrate the use of tiny (the size of a coffee

cup) low‐power inspection satellites that can be

sent out to observe larger spacecraft.  The small

inspection satellites are enabled by

microelectromechanical systems (MEMS), and

will test the functioning of small camera

systems and gyros.

Radar Fence Transponder (RAFT) is a student

experiment from the United Sates Naval

Academy that uses picosatellites to test the

Space Surveillance Radar Fence and

experimental communications transponders.

Payloads Delivered on STS-116 to Be 

Used in Future Station Research 

Cardiovascular and Cerebrovascular Control

on Return from ISS (CCISS) will allow

scientists to better understand the effects of

spaceflight on cardiovascular and

cerebrovascular functions of long duration

crewmembers.

Commercial Generic Bioprocessing Apparatus

Science Insert – 01 (CSI‐01) is comprised of

two educational experiments that will be

utilized by middle school students.  One

experiment will examine seed germination in

microgravity including gravitropism (plant

growth towards gravity) and phototropism

(plant growth towards light).  The second

experiment will examine how microgravity

affects the model organism, Caenorhabditis

elegans, a small nematode worm.

Elastic Memory Composite Hinge (EMCH)

will study the performance of a new type of
December 2006 EXPERIMENTS 82

composite hinge to determine if it is suitable for

use in space.  The experiment will use elastic

memory hinges to move an attached mass at

one end.  Materials tested in this experiment are

stronger and lighter than current material used

in space hinges and could be used in the design

of future spacecraft.

Threshold Acceleration for Gravisensing

(Gravi) will determine the minimum amount of

artificial gravity needed to cause lentil seedling

roots to start growing in a new direction.  This

work supports future efforts to grow sufficient

edible crops on long‐duration space missions.

Lab‐on‐a‐Chip Application Development‐

Portable Test System (LOCAD‐PTS) is a

handheld device for rapid detection of

biological and chemical substances on board

the space station.  Astronauts will swab

surfaces within the cabin, add swab material to

the LOCAD‐PTS and within 15 minutes, obtain

results on a display screen.  Its purpose is to

effectively provide an early warning system to

enable crew members to take remedial

measures if necessary to protect the health and

safety of those on board the station.

Synchronized Position Hold, Engage,

Reorient, Experimental Satellites (SPHERES)

are bowling‐ball sized spherical satellites.  The

first SPHERE satellite arrived on the station in

April 2006 tucked inside a Russian Progress

supply ship.  Another arrived on STS‐121 in

July 2006 and a third will be carried to orbit by

STS‐116 in December 2006.  They will be used

inside the space station to test a set of well‐

defined instructions for spacecraft to perform

autonomous rendezvous and docking

maneuvers.  Three self‐contained free‐flying

spheres will fly within the cabin of the station,  

performing flight formations.  Each satellite is

self‐contained with power, propulsion,

computers and navigation equipment.  The

results are important for satellite servicing,

vehicle assembly and formation flying

spacecraft configurations.

Test of Reaction and Adaptation Capabilities

(TRAC) will test the theory of brain adaptation

during spaceflight by testing hand‐eye

coordination before, during and after the

mission.

Space Station Research Samples 

Returned on STS-116 

Long Term Microgravity:  A Model for

Investigating Mechanisms of Heart Disease

with New Portable Equipment (Card) will

examine the relationship between salt intake

and the cardiovascular system when exposed to

the microgravity environment.

Neuroendocrine and Immune Responses in

Humans During and After Long Term Stay at

ISS (IMMUNO) will provide an understanding

for the development of pharmacological tools to

countermeasure unwanted immunological side

effects during long‐duration missions in space.

The Optimization of Root Zone Substrates

(ORZS) for Reduced Gravity Experiments

Program is an international collaboration that

will provide direct measurements and models

for plant rooting media that will be used in

future Advanced Life Support (ALS) plant

growth experiments.

Passive Observatories for Experimental

Microbial Systems (POEMS) is a

demonstration of a passive system for growing

microbial cultures in space and to observe

genetic changes that occur in them as a result of

living and growing in the space environment.
December 2006 EXPERIMENTS 83

The Renal Stone experiment tests the

effectiveness of potassium citrate in preventing

renal stone formation during long‐duration

spaceflight.  Kidney stone formation, a

significant risk during long missions, could

impair astronaut functionality.

Space Experiment Module (SEM) allows

students to research the effects of microgravity,

radiation and spaceflight on various materials.  

Some students will test for seed growth after

microgravity exposure, while others will test

how their materials protect against radiation

exposure.  The sample vials will be returned to

Earth on a future mission for students to

analyze further.

Stability of Pharmacotherapeutic and

Nutritional Compounds (Stability) will study

the effects of radiation in space on complex

organic molecules, such as vitamins and other

compounds in food and medicine.  This will

help in developing more stable and reliable

pharmaceutical and nutritional

countermeasures suitable for future long‐

duration missions to the moon and Mars.

A Comprehensive Characterization of

Microorganisms and Allergens in Spacecraft

(SWAB) will use advanced molecular

techniques to comprehensively evaluate

microbes on board the space station, including

pathogens – organisms that may cause disease.  

It also will track changes in the microbial

community as spacecraft visit the station and

new station modules are added.  This study

will allow an assessment of the risk of microbes

to the crew and the spacecraft.

Analysis of a Novel Sensory Mechanism in

Root Phototropism (Tropi) will observe the

growth and collect samples of plants sprouted

from seeds.  By analyzing the samples at a

molecular level, researchers expect to gain

insight on what genes are responsible for

successful plant growth in microgravity.

Additional Space Station Research 

From Now Until the End of Expedition 14 

Anomalous Long Term Effects in Astronautsʹ 

Central Nervous System (ALTEA) integrates

several diagnostic technologies to measure the

exposure of crew members to cosmic radiation.  

It will further our understanding of radiation’s

impact on the human central nervous and

visual systems, and provide an assessment of

the radiation environment in the station.

Crew Earth Observations (CEO) takes

advantage of the crew in space to observe and

photograph natural and human‐made changes

on Earth.  The photographs record the Earth’s

surface changes over time, along with more

fleeting events such as storms, floods, fires and

volcanic eruptions.  Together, they provide

researchers on Earth with vital, continuous

images to better understand the planet.

Space Flight‐Induced Reactivation of Latent

Epstein‐Barr Virus (Epstein‐Barr) performs

tests to study changes in the human immune

function.  Using blood and urine samples

collected from crew members before and after

spaceflight, the study will provide insight for

possible countermeasures to prevent the

potential development of infectious illness in

crew members during flight.

Behavioral Issues Associated with Isolation

and Confinement:  Review and Analysis of

Astronaut Journals, using journals kept by the

crew and surveys, studies the effect of isolation

to obtain quantitative data on the importance of

different behavioral issues in long‐duration

crews.  Results will help NASA design

equipment and procedures to allow astronauts

to best cope with isolation and long‐duration

spaceflight.
December 2006 EXPERIMENTS 84

Microgravity Acceleration Measurement

System (MAMS) and Space Acceleration

Measurement System (SAMS‐II) measure

vibration and quasi‐steady accelerations that

result from vehicle control burns, docking and

undocking activities.  The two different

equipment packages measure vibrations at

different frequencies.

Materials on the International Space Station

Experiment 3 and 4 (MISSE – 3 and 4) are the

third and fourth in a series of five suitcase‐sized

test beds attached to the outside of the space

station.  The beds were deployed during a

spacewalk on STS‐121 in July 2006.  They will

expose hundreds of potential space

construction materials and different types of

solar cells to the harsh environment of space.  

After being mounted to the space station about

a year, the equipment will be returned to Earth

for study.  Investigators will use the resulting

data to design stronger, more durable

spacecraft.

Nutritional Status Assessment (Nutrition) is

the most comprehensive in‐flight study done by

NASA to date of human physiologic changes

during long‐duration space flight; this includes

measures of bone metabolism, oxidative

damage, nutritional assessments, and hormonal

changes.  This study will impact both the

definition of nutritional requirements and

development of food systems for future space

exploration missions to the Moon and Mars.  

This experiment will also help to understand

the impact of countermeasures (exercise and

pharmaceuticals) on nutritional status and

nutrient requirements for astronauts.

Sleep‐Wake Actigraphy and Light Exposure

During Spaceflight‐Long (Sleep‐Long) will

examine the effects of spaceflight on the sleep‐

wake cycles of the crew members during long‐

duration stays on the space station.

EUROPEAN SPACE AGENCY 

EXPERIMENTS 

The majority of Christer Fuglesang’s time at the

station will be taken up with station assembly

tasks.  However, he will still be undertaking a

number of experiments and additional activities

during the mission.  One experiment

(Chromosome‐2), one activity monitoring

radiation dosimetry (EuCPD), and two

education activities (ALTEA Filmed Lesson,

Frisbee Competition) are supported by the

European Space Agency (ESA).  The Particle

Flux Demonstrator is a Swedish National Space

Board education experiment, and ALTEA is

supported by ASI, the Italian Space Agency.

ALTEA (Human Physiology/Radiation 

Dosimetry) 

The ALTEA project investigates the effects of

cosmic radiation on brain function.  The focus

of the program will be on abnormal visual

perceptions (often reported as ‘‘light flashes’’

by astronauts) and the impact that particle

radiation has on the retina and visual structures

of the brain under weightless conditions.  

ALTEA will also provide more information on

the radiation environment in the station.

The ALTEA facility is a helmet‐shaped device,

which covers most of the astronaut’s head.  It

consists of six particle detectors and will permit

a 3‐D reconstruction of cosmic radiation

passing through the brain:  measuring particle

trajectory, energy and particle type.  At the

same time, a 32‐channel EEG will measure the

astronaut’s brain activity and a visual

stimulator and a pushbutton will be used to
ALTEA Flight Hardware.

December 2006 EXPERIMENTS 85

determine visual performance and occurrence

of light flashes.  When not in use by an

astronaut, the ALTEA device will be used to

collect continuous measurements of the cosmic

radiation in the U.S. Destiny laboratory on the

station.

The ALTEA experiment follows on from the

precursor Alteino experiment that took place

during the European Marco Polo and Eneide

missions (2002 and 2005), now continuing with

the European ALTCRISS project.

The ALTEA experiment is particularly

important with the increasing length of human

operation on board the International Space

Station and in the perspective of longer‐term

journeys to Mars.  Results may also hold

benefits on Earth in neuroscience in, for

example, the use of Ion therapies to treat brain

tumors.

Chromosome-2 

(Human Physiology) 

During spaceflights, crew members are exposed

to different types of ionizing radiation.  To

assess the impact of these radiations at a genetic

level, an analysis will be made of abnormalities

in chromosome number and structure in

lymphocytes (white blood cells) of station crew

members.

Venous blood samples will be taken a few days

before launch and directly on return from the

long‐duration and short‐duration missions.  

From these blood samples, whole blood

cultures will be set up with a substance

(phytohemagglutinin), which stimulates the

lymphocytes to enter the cell cycle, i.e., to go

through the cell division process.

Forty‐eight hours after the start of the

incubation period, the samples will be prepared

for analysis.  This consists of adding two

substances to the samples.  One (calyculin A)

will cause the DNA strands to become tightly

packed in the chromosomes, a process that

occurs in cells at the beginning of the cell cycle.  

The other (colchicine) will prevent the

chromosomes from dividing into two for cells

at a later stage in the cell cycle.

Hereafter the preparations will be analyzed by

one of two methods to determine the full

spectrum of abnormalities in the chromosomes.  

This will either be by the fluorescence in vitro

hybridization (FISH) method or classical

Giemsa staining.

Multi‐fluorescent chromosome map of a cell

exposed to cosmic radiation
A small portion of the blood samples will be

irradiated with X‐rays.  The frequency of

abnormalities induced by this irradiation will

provide information about the radiosensitivity

of individual crew members and probable

changes in sensitivity caused by spaceflight.

The Chromosome‐2 experiment is planned to

be carried out using eight subjects:  four

subjects from short‐duration flights and four

Expedition crew members.

European Crew Personal Dosimeters 

(EuCPD) (Radiation dosimetry) 

ESA astronaut Fuglesang from Sweden will be

equipped with passive personal radiation

dosimeters to measure the radiation exposure

during his flight.  ESA astronaut Thomas Reiter

has been using similar dosimeters since

arriving at the station in July 2006.

The personal dosimeters have to be worn

around the left‐hand side of the waist and the

left ankle for astronauts inside the station and

at the same locations above the liquid cooling

garment inside the spacesuit for astronauts

undertaking spacewalks.

Each dosimeter is only 8 mm thick and consists

of a stack of five different passive radiation

sensors, i.e., sensors that provide a measure of

the overall radiation dose rather than a measure

of radiation with relation to time.  The different

sensors will measure different radioactive

particles such as a range of neutrons (thermal,

epithermal and fast), and heavy ions as well as

measuring particle impact angles and energy

transfer from particles.

The different sensors to be used will include a

layer of thermo‐luminescence dosimeters that

are also used in the European Matroshka

radiation dosimetry experiment that has been at

the station since its launch on Jan. 29, 2004.

December 2006 EXPERIMENTS 86

ESA astronaut Thomas Reiter wearing

European Crew Personal Dosimeter  

during sample processing on the ISS

Two European Crew Personal Dosimeters will

be worn by Fuglesang during launch.  The

others will be transported to the station as

shuttle cargo.

Post‐flight analysis of the different radiation

sensors will be carried out in a laboratory

environment according to standardized

methods developed by the science team.

Particle Flux Demonstrator (Education) 

This experiment aims to illustrate to students

the difference in the radiation environment in

space and on Earth.  The Particle Flux

Demonstrator will be used to measure the flux

or flow of charged particles (cosmic rays)

passing through the station.  This will vary to

the greatest degree with the variation of the

latitude at which the station is flying.

Charged particles are detected in the Particle

Flux Demonstrator using a plastic scintillator

material connected to a photomultiplier tube.  

The charged particle generates a very faint flash

of deep blue light in the scintillator material.  

This is converted to an electrical pulse by the

photomultiplier tube.  To trigger the Particle

Flux Demonstrator, the charged particle must

pass through two pieces of scintillator
December 2006 EXPERIMENTS 87

Particle Flux Demonstrator hardware

separated by a few centimeters.  This gives the

Particle Flux Demonstrator some directional

sensitivity.  On the station, the majority of

detected particles will be protons.  However, by

varying a requirement on the amount of light

produced in one of the scintillators, it is

possible to select much rarer charged ions also.

The Particle Flux Demonstrator is a portable

device that can be carried through the station

by the astronaut.  It is expected to show how

the flux varies in different parts of the station,

depending on how much local shielding

material there is.  One will also see the change

in flux with the orbital position:  higher flux at

high latitudes and lower near the equator, due

to the interaction of the Earth’s magnetic field.  

An additional aim is to demonstrate that the

relative abundance of ions increases at high

latitudes.

Special attention will be paid to the South

Atlantic Anomaly, where the external particle

flux is several orders of magnitude larger than

elsewhere in the orbit due a dip in the Earth’s

magnetic field in the vicinity of the Brazilian

coast.  The behavior of the magnetic field

means that a belt of particles approaches close

to the Earth’s surface.  The particle flux inside

the station will also increase drastically at this

location.  Additionally, it is expected to

demonstrate the directionality of particles

within the radiation belt

On ground, the Particle Flux Demonstrator will

be able to detect muons (an elementary particle

similar to the electron but with ~200 times the

mass).  These are produced by cosmic ray

particles interacting with molecules high up in

the atmosphere.  However, the rate will be

small—about 1 muon per minute (with the

“telescope” orientated vertically).

The Particle Flux Demonstrator forms part of an

education experiment at high school and

university levels.  Several education

institutions, principally in Sweden but also in

Norway and Denmark, will be taking part in

this experiment using cosmic ray detectors to

take readings on Earth during station flyovers.  

The readings from the institutions will be

compared against readings from the Particle

Flux Demonstrator on the station.  Video

footage of the on‐orbit experiment with

Fuglesang will be made available to students.

The Particle Flux Demonstrator was designed

and built in the Particle and Astroparticle

Physics research group at The Royal Institute of

Technology (KTH) in Stockholm, Sweden.  As

part of a public relations event in Stockholm,

Fuglesang will demonstrate the Particle Flux

Demonstrator during a live video‐link to the

station.

ALTEA: Filmed Lesson on Radiation 

(Education) 

This activity involves the filming of a mini

documentary about the ALTEA experiment to
NASA astronaut Jeff Williams next to the

ALTEA hardware in the U.S. Laboratory  

of the ISS during Expedition 13

December 2006 EXPERIMENTS 88

form part of a lesson on radiation for 16‐18 year

olds.  While ESA astronaut Fuglesang is

undertaking the ALTEA experiment on the

station as part of his mission, ESA astronaut

Thomas Reiter, currently on the station as an

Expedition 14 flight engineer, will film the

experiment.

The footage will be recorded on video tape and

returned to Earth.  This footage will be

combined with additional footage that was

filmed during Fuglesang’s training with

ALTEA hardware at the Johnson Space Center

in Houston before launch and accompanied

with commentary from a member of the

ALTEA science team and by Fuglesang himself.  

Fuglesang has been previously involved as a

member of the scientific team for a similar

experiment:  SilEye, which investigated light

flashes in astronautsʹ eyes on the Mir Space

Station between 1995 and 1999.

The ISS Education Office will coordinate the

development and content of the lesson together

with the ALTEA science team.  Once the

production is complete, ESA will distribute the

lesson to European secondary school teachers

and their students.

Frisbee Competition (Education)

A little known fact about Fuglesang is that he

was once a Swedish national Frisbee champion,

holding the national title in ‘maximum time

aloft’ in 1978, i.e., the longest time a Frisbee can

stay in the air.  He later competed in the Frisbee

World Championship in 1981.

This activity is a competition for children up to

15 years of age that will stimulate the interest of

children in space and culminate in a live linkup

with the competition winner on the ground and

Fuglesang on the station.  During this in‐flight

call, Fuglesang will try to break the current

world record for maximum time aloft, which

stands at 16.72 seconds.  This should not prove

to be too difficult a prospect as Fuglesang has

the advantage of weightlessness on his side.

The competition will be organized in Sweden

by the national newspaper ʺAftonbladetʺ and

will be based on questions including the topics

of space, space activities and Frisbee games.  

The competition winner will be announced

during the in‐flight call and they will present

Fuglesang with the licence, necessary to be able

to compete in professional competition.  Once

this has occurred, Fuglesang will begin his

world record attempt.

The Frisbee that Fuglesang will take to the

station with him is one of his personal Frisbees,

which marks an Atlantic crossing he made.

The Frisbee that Fuglesang will  

be taking into space
December 2006 SPACE SHUTTLE MAIN ENGINE 89

SPACE SHUTTLE MAIN ENGINE 

ADVANCED HEALTH MANAGEMENT SYSTEM 

The Advanced Health Management System

(AHMS) will make its first flight in monitor‐

only mode on the STS‐116 mission.  When in

monitor‐only mode, the AHMS system collects

and processes the turbopump accelerometer

data, which is a measurement of turbopump

vibration, and continuously monitors

turbopump health but cannot shut down an

engine.  Data from the STS‐116 flight will be

collected and examined to ensure AHMS

operates as intended before its subsequent

flight in active mode.

When a NASA space shuttle lifts off the launch

pad it does so with the help of three reusable,

high performance rocket engines.  Each space

shuttle main engine is fourteen feet long, seven

and one‐half feet in diameter at the nozzle exit,

weighs approximately 7,750 pounds and

generates more than twelve million

horsepower, which is equivalent to the output

of more than seven Hoover Dams.  The engines

operate for about eight‐and‐one‐half minutes

during a shuttle liftoff and ascent – long

enough to burn more than 500,000 gallons of

super‐cold liquid hydrogen and liquid oxygen

propellants stored in the huge external tank

attached to the underside of the orbiter.  Liquid

oxygen is stored at ‐298 degrees Fahrenheit and

liquid hydrogen at ‐423 degrees Fahrenheit.  

The engines shut down just before the orbiter,

traveling at about 17,000 mph, reaches orbit.

NASA is confident that this engine upgrade

will significantly improve space shuttle flight

safety and reliability.  The upgrade, developed

by NASA’s Marshall Space Flight Center in

Huntsville, Ala., is a modification of the

existing main engine controller, which is the

on‐engine computer that monitors and controls

all main engine operations.  The modifications

include the addition of advanced digital signal

processors, radiation‐hardened memory and

new software.  These changes to the main

engine controller provide the capability for

completely new monitoring and insight into the

health of the two most risky components of the

space shuttle main engine – the high pressure

fuel turbopump and the high pressure oxidizer

turbopump.

The high pressure fuel and high pressure

oxidizer turbopumps rotate at approximately

35,000 revolutions per minute and 28,000

revolutions per minute, respectively.  To

operate at such extreme speeds, the high

pressure turbopumps utilize highly specialized

bearings and precisely balanced components.

The AHMS upgrade utilizes data from three

existing sensors (accelerometers) mounted on

each of the high pressure turbopumps to

measure how much each pump is vibrating.  

The output data from the accelerometers is

routed to the new AHMS digital signal

processors installed in the main engine

controller.  These processors analyze the sensor

readings 20 times per second looking for

vibration anomalies that are indicative

of impending failure of rotating turbopump

components such as blades, impellers, inducers

and bearings.  If the magnitude of any vibration

anomaly exceeds safe limits, the upgraded

main engine controller immediately would shut

down the unhealthy engine.

December 2006 SHUTTLE REFERENCE DATA 90

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

heads‐up attitude, pointing toward the launch
December 2006 SHUTTLE REFERENCE DATA 91

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

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
December 2006 SHUTTLE REFERENCE DATA 92

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

displays and display information, to determine

the abort region.
December 2006 SHUTTLE REFERENCE DATA 93

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,

NASA managers set Oct. 2 as the date for

Endeavour’s second launch attempt.
December 2006 SHUTTLE REFERENCE DATA 94

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

engines are throttled back up to normal

operating level at about 60 seconds.  This

reduces stress on the vehicle.
December 2006 SHUTTLE REFERENCE DATA 95

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

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
December 2006 SHUTTLE REFERENCE DATA 96

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.  

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
December 2006 SHUTTLE REFERENCE DATA 97

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

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
December 2006 SHUTTLE REFERENCE DATA 98

(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

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
December 2006 SHUTTLE REFERENCE DATA 99

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.

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
December 2006 SHUTTLE REFERENCE DATA 100

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

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
December 2006 SHUTTLE REFERENCE DATA 101

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

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
December 2006 SHUTTLE REFERENCE DATA 102

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.

December 2006 ACRONYMS/ABBREVIATIONS 103

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

BC Bus Controller
December 2006 ACRONYMS/ABBREVIATIONS 104

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 Macrophages in Spac

e

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

CID Circuit Interrupt Device

CIOB Cargo Integration and Operations Branch
December 2006 ACRONYMS/ABBREVIATIONS 105

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

dp/dt delta pressure/delta time

DPA Digital Preassembly

DPS Data Processing System
December 2006 ACRONYMS/ABBREVIATIONS 106

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 System

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

EVR Extravehicular Robotics

EVSU External Video Switching Unit

EXPRESS EXpedite the PRocessing of Experiments to the Space Station

EXT Experimental Terminal
December 2006 ACRONYMS/ABBREVIATIONS 107

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

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 NavigationalSatellite 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

HCA Hollow Cathode Assembly

HCOR High‐Rate Communication Outage Recorder

HDR High Data Rate

HDRL High Date Rate Link
December 2006 ACRONYMS/ABBREVIATIONS 108

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

ISPR International Standard Payload Rack

ISS International Space Station

ISSSH International Space Station Systems Handbook

IT Integrated Truss

ITCS Internal Thermal Control System
December 2006 ACRONYMS/ABBREVIATIONS 109

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

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

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

December 2006 ACRONYMS/ABBREVIATIONS 110
December 2006 ACRONYMS/ABBREVIATIONS 111

N2 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

NiH2 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‐Orbit Maintainable Item

OMS Orbital Maneuvering System

OPCGA Observable Protein Crystal Growth Apparatus

OPP OSVS Patch Panel

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

OSVS Orbiter Space Vision System

OTD ORU Transfer Device

OV Orbiter Vehicle

P&S Pointing and Support

P‐code Precision Code
December 2006 ACRONYMS/ABBREVIATIONS 112

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

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
December 2006 ACRONYMS/ABBREVIATIONS 113

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

RBVM Radiator Beam Valve  

RCC Reinforced Carbon‐Carbon

RCS Reaction Control System

RDA Retainer Door Assembly

RF Radio Frequency

RFCA Rack Flow Control Assembly

RFG Radio Frequency Group
December 2006 ACRONYMS/ABBREVIATIONS 114

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 Solar Array

SABB Solar Array Blanket Box

SAGE Space Arabidoposis Genomics Experiment

SARJ Solar Alpha Rotary Joint

SARJ_C SARJ Controller

SARJ_M SARJ Manager

SASA S‐band Antenna Support Assembly

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
December 2006 ACRONYMS/ABBREVIATIONS 115

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

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
December 2006 ACRONYMS/ABBREVIATIONS 116

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

UTA Utility Transfer Assembly

VAJ Vacuum Access Jumper

VBSP Video Baseband SignalProcessor

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
December 2006 ACRONYMS/ABBREVIATIONS 117

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

December 2006 MEDIA ASSISTANCE 118

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

December 2006 MEDIA CONTACTS 119

PUBLIC AFFAIRS CONTACTS 

HEADQUARTERS 

WASHINGTON, DC 

Allard Beutel

Public Affairs Officer

Space Operations

202‐358‐4769

Katherine Trinidad

Public Affairs Officer

Space Operations

202‐358‐3749

Grey Hautaluoma

Public Affairs Officer

Space Operations

202‐358‐0668

Melissa Mathews

Public Affairs Officer

International Partners

202‐358‐1272

JOHNSON SPACE CENTER 

HOUSTON, TEXAS 

James Hartsfield

News Chief

281‐483‐5111

Kyle Herring

Public Affairs Specialist

Space Shuttle Program Office

281‐483‐5111

Kylie Clem

Public Affairs Specialist

Mission Operations and

Flight Crew Operations

281‐483‐5111

Rob Navias

Program and Mission Operations Lead

281‐483‐5111

Lynnette Madison

Public Affairs Specialist

International Space Station and Engineering

281‐483‐5111

KENNEDY SPACE CENTER 

FLORIDA 

Bruce Buckingham

News Chief

321‐861‐2468

Jessica Rye

Public Affairs Specialist

Space Shuttle

321‐867‐2468

Tracy Young

Public Affairs Specialist

International Space Station

321‐867‐2468

MARSHALL SPACE FLIGHT CENTER 

HUNTSVILLE, ALABAMA 

Dom Amatore

Public Affairs Manager

256‐544‐0034

June Malone

Public Affairs Specialist

Space Shuttle Propulsion

256‐544‐0034
December 2006 MEDIA CONTACTS 120

STENNIS SPACE CENTER 

MISSISSIPPI 

Linda Theobald

Public Affairs Specialist

228‐688‐3249

Paul Foerman

News Chief

228‐688‐1880

Rebecca Strecker

Public Affairs Specialist

228‐688‐3346

AMES RESEARCH CENTER 

MOFFETT FIELD, CALIFORNIA 

Mike Mewhinney

News Chief

650‐604‐3937

Jonas Dino

Public Affairs Specialist

650‐604‐5612

DRYDEN FLIGHT RESEARCH CENTER 

CALIFORNIA 

Alan Brown

News Chief

661‐276‐2665

Leslie Williams

Public Affairs Specialist

661‐276‐3893

GLENN RESEARCH CENTER 

CLEVELAND, OHIO 

Lori Rachul

News Chief

216‐433‐8806

Katherine Martin

Public Affairs Specialist

216‐433‐2406

LANGLEY RESEARCH CENTER 

HAMPTON, VIRGINIA 

H. Keith Henry

Deputy, Office of Public Affairs

757‐864‐6120

UNITED SPACE ALLIANCE 

Mike Curie

Space Program Operations Contract

Houston Operations

281‐483‐9251

321‐861‐3805

Tracy Yates

Florida Operations

321‐861‐3956

BOEING 

Ed Memi

Media Relations

Boeing NASA Systems

281‐226‐4029

EUROPEAN SPACE AGENCY 

Clare Mattok,

ESA Head Office

Paris, France

Phone + 33 1 5369 7412

Fax + 33 1 5369 7690

Clare.Mattok@esa.int

Brigitte Kolmsee,

ESA Head Office

Paris, France

Phone + 33 1 5369 7299

Fax + 33 1 5369 7690

Brigitte.Kolmsee@esa.int
December 2006 MEDIA CONTACTS 121

Franco Bonacina,

ESA Head Office

Paris, France

Phone + 33 1 5369 7713

Fax + 33 1 5369 7690

Franco.Bonacina1@esa.int

Rosita Suenson,

ESA/ESTEC

Noordwijk, the Netherlands

Phone + 31 71 565 3009

Fax + 31 71 565 5728

Rosita.Suenson@esa.int