INERTIAL UPPER STAGE
 
The inertial upper stage will be used with the space shuttle to
transport NASA's Tracking and Data Relay satellites to geosynchronous
orbit, 22,300 statute miles from Earth.  The IUS was also selected by
NASA for the Magellan, Galileo and Ulysses planetary missions.
 
The IUS was originally designed as a temporary stand-in for a
reusable space tug and was called the interim upper stage.  Its name
was changed to inertial upper stage (signifying the satellite's
guidance technique) when it was realized that the IUS would be needed
through the mid-1990s.
 
The IUS was developed and built under contract to the Air Force
Systems Command's Space Division.  The Space Division is executive
agent for all Department of Defense activities pertaining to the
space shuttle system and provides the IUS to NASA for space shuttle
use.  In August 1976, after 2.5 years of competition, Boeing
Aerospace Company, Seattle, Wash., was selected to begin preliminary
design of the IUS.
 
The IUS is a two-stage vehicle weighing approximately 32,500 pounds.
Each stage is a solid rocket motor.  This design was selected over
those with liquid-fueled engines because of its relative simplicity,
high reliability, low cost and safety.
 
The IUS is 17 feet long and 9.5 feet in diameter.  It consists of an
aft skirt, an aft stage solid rocket motor with 21,400 pounds of
propellant generating 45,600 pounds of thrust, an interstage, a
forward stage solid rocket motor with 6,000 pounds of propellant
generating 18,500 pounds of thrust and using an extendable exit cone,
and an equipment support section.  The equipment support section
contains the avionics that provide guidance, navigation, telemetry,
command and data management, reaction control and electrical power.
All mission-critical components of the avionics system and thrust
vector actuators, reaction control thrusters, motor igniter and
pyrotechnic stage separation equipment are redundant to ensure better
than 98-percent reliability.
 
AIRBORNE SUPPORT EQUIPMENT.
 
The IUS ASE is the mechanical, avionics and structural equipment
located in the orbiter.  The ASE supports and provides services to
the IUS and the spacecraft in the orbiter payload bay and positions
the IUS/spacecraft in an elevated position for final checkout before
deployment from the orbiter.
The IUS ASE consists of the structure, batteries, electronics and
cabling to support the IUS and spacecraft combination.  These ASE
subsystems enable the deployment of the combined vehicle and provide
or distribute and control electrical power to the IUS and spacecraft
and provide communication paths between the IUS, spacecraft and the
orbiter.
 
The ASE incorporates a low-response spreader beam and torsion bar
mechanism that reduces spacecraft dynamic loads to less than
one-third what they would be without this system.  In addition, the
forward ASE frame includes a hydraulic load leveler system to provide
a balanced loading at the forward trunnion fittings.
 
The ASE data subsystem allows data and commands to be transferred
between the IUS and spacecraft and the appropriate orbiter interface.
 Telemetry data include spacecraft data received over dedicated
circuits via the IUS and spacecraft telemetry streams.  An
interleaved stream is provided to the orbiter to transmit to the
ground or transfer to ground support equipment.
 
The structural interfaces in the orbiter payload bay consist of six
standard non-deployable attach fittings on each longeron that mate
with the ASE aft and forward support frame trunnions and two payload
retention latch actuators at the forward ASE support frame.  The IUS
has a self-contained, spring-actuated deployment system that imparts
a velocity to the IUS at release from the raised deployment attitude.
 Ducting from the orbiter purge system interfaces with the IUS at the
forward ASE.
 
 
IUS STRUCTURE.
 
The IUS structure is capable of transmitting all of the loads
generated internally and also those generated by the cantilevered
spacecraft during orbiter operations and IUS free flight.  In
addition, the structure supports all of the equipment and solid
rocket motors within the IUS and provides the mechanisms for IUS
stage separation.  The major structural assemblies of the two-stage
IUS are the equipment support section, interstage and aft skirt.  The
basic structure is made from aluminum skin-stringer construction with
six longerons and ring frames.
 
EQUIPMENT SUPPORT SECTION.
 
The ESS houses the majority of the IUS avionics and control
subsystems.  The top of the ESS contains the 10-foot-diameter
interface mounting ring and electrical interface connector segment
for mating and integrating the spacecraft with the IUS.  Thermal
isolation is provided by a multilayer insulation blanket across the
interface between the IUS and spacecraft.  All line replaceable units
mounted in the ESS can be removed and replaced via access doors even
when the IUS is mated with the spacecraft.
 
IUS AVIONICS SUBSYSTEM.
 
The avionics subsystem consists of the telemetry, tracking and
command; guidance and navigation; data management; thrust vector
control; and electrical power subsystems.  This includes all of the
electronic and electrical hardware used to perform all computations,
signal conditioning, data processing and software formatting
associated with navigation, guidance, control, data management and
redundancy management.  The IUS avionics subsystem also provides the
communications between the orbiter and ground stations and electrical
power distribution.
 
Data management performs the computation, data processing and signal
conditioning associated with guidance, navigation and control; safing
and arming and ignition of the IUS two-stage solid rocket motors and
electroexplosive devices; command decoding and telemetry formatting;
and redundancy management and issues spacecraft discretes.  The data
management subsystem consists of two computers, two signal
conditioner units and a signal interface unit.
 
Modular general-purpose computers use operational flight software to
perform in-flight calculations and to initiate the vehicle thrust and
attitude control functions necessary to guide the IUS and spacecraft
through a flight path determined on board to a final orbit or planned
trajectory.  A stored program, including data known as the onboard
digital data load, is loaded into the IUS flight computer memory from
magnetic tape through the memory load unit during prelaunch
operations.  Memory capacity is 65,536 (64K) 16-bit words.
 
The signal conditioner unit provides the interface for commands and
measurements between the IUS avionics computers and the IUS
pyrotechnics; power; reaction control system; thrust vector control;
telemetry, tracking and command; and the spacecraft.  The SCU
consists of two channels of signal conditioning and distribution for
command and measurement functions.  The two channels are designated A
and B.  Channel B is redundant to channel A for each measurement and
command function.
 
The signal interface unit performs buffering, switching, formatting
and filtering of TT&C interface signals.
 
Communications and power control equipment are mounted at the orbiter
aft flight deck payload station and operated in flight by the orbiter
flight crew mission specialists.  Electrical power and signal
interfaces to the orbiter are located at the IUS equipment
connectors.  Cabling to the orbiter equipment is provided by the
orbiter.  In addition, the IUS provides dedicated hardwires from the
spacecraft through the IUS to an orbiter multiplexer/demultiplexer
for subsequent display on the orbiter cathode ray tube of parameters
requiring observation and correction by the orbiter flight crew.
This capability is provided until IUS ASE umbilical separation.
 
To support spacecraft checkout or other IUS-initiated functions, the
IUS can issue a maximum of eight discretes.  These discretes may be
initiated either manually by the orbiter flight crew before the IUS
is deployed from the orbiter or automatically by the IUS
mission-sequencing flight software after deployment.  The discrete
commands are generated in the IUS computer either as an
event-scheduling function (part of normal onboard automatic
sequencing) or a command-processing function initiated from an uplink
command from the orbiter or Air Force Consolidated Satellite Test
Center to alter the onboard event-sequencing function and permit the
discrete commands to be issued at any time in the mission.
 
During the ascent phase of the mission, the spacecraft's telemetry is
interleaved with IUS telemetry, and ascent data are provided to
ground stations in real time via orbiter downlink.  Telemetry
transmission on the IUS RF link begins after the IUS and spacecraft
are tilted for deployment from the orbiter.  Spacecraft data may be
transmitted directly to the ground when the spacecraft is in the
orbiter payload bay with the payload bay doors open or during IUS and
spacecraft free flight.
 
IUS guidance and navigation consist of strapped-down redundant
inertial measurement units.  The redundant IMUs consist of five
rate-integrating gyros, five accelerometers and associated
electronics.  The IUS inertial guidance and navigation subsystem
provides measurements of angular rates, linear accelerations and
other sensor data to data management for appropriate processing by
software resident in the computers.  The electronics provides
conditioned power, digital control, thermal control, synchronization
and the necessary computer interfaces for the inertial sensors.  The
electronics are configured to provide three fully independent
channels of data to the computers.  Two channels each support two
sets of sensors and the third channel supports one set.  Data from
all five gyro and accelerometer sets are sent simultaneously to both
computers.
 
The guidance and navigation subsystem is calibrated and aligned on
the launch pad.  The navigation function is initialized at lift-off,
and data from the redundant IMUs is integrated in the navigation
software to determine the current state vector.  Before vehicle
deployment, an attitude update maneuver may be performed by the
orbiter.
 
If for any reason the computer is powered down before deployment, the
navigation function is reinitialized by transferring orbiter
position, velocity and attitude data to the IUS vehicle.  Attitude
updates are then performed as described above.
 
The IUS vehicle uses an explicit guidance algorithm (gamma guidance)
to generate thrust steering commands, solid rocket motor ignition
time and RCS vernier thrust cutoff time.  Before each SRM ignition
and each RCS vernier, the vehicle is oriented to a thrust attitude
based on nominal performance of the remaining propulsion stages.
During SRM burn, the current state vector determined from the
navigation function is compared to the desired state vector, and the
commanded attitude is adjusted to compensate for the buildup of
position and velocity errors caused by off-nominal SRM performance
(thrust, specific impulse).
 
Vernier thrust compensates for velocity errors resulting from SRM
impulse and cutoff time dispersions.  Residual position errors from
the SRM thrusting and position errors introduced by impulse and
cutoff time dispersions are also removed by the RCS.
 
Attitude control in response to guidance commands is provided by
thrust vector control during powered flight and by reaction control
thrusters during coast.  Measured attitude from the guidance and
navigation subsystem is compared with guidance commands to generate
error signals.  During solid motor thrusting, these error signals
drive the motor nozzle actuator electronics in the TVC subsystem.
The resulting nozzle deflections produce the desired attitude control
torques in pitch and yaw.  Roll control is maintained by the RCS
roll-axis thrusters.  During coast flight, the error signals are
processed in the computer to generate RCS thruster commands to
maintain vehicle attitude or to maneuver the vehicle.  For attitude
maneuvers, quarternion rotations are used.
 
TVC provides the interface between IUS guidance and navigation and
the SRM gimbaled nozzle to accomplish powered- flight attitude
control.  Two complete electrically redundant channels minimize
single-point failure.  The TVC subsystem consists of two controllers,
two actuators and four potentiometers for each IUS SRM.  Power is
supplied through the SCU to the TVC controller that controls the
actuators.  The controller receives analog pitch and yaw commands
that are proportioned to the desired nozzle angle and converts them
to pulsewidth-modulated voltages to power the actuator motors.  The
motor drives a ball screw that extends or retracts the actuator to
position the SRM nozzle.  Potentiometers provide servoloop closure
and position instrumentation.  A staging command from the SCU allows
switching of the controller outputs from IUS first-stage actuators to
the IUS second-stage actuators.
 
The IUS's electrical power subsystem consists of avionics batteries,
IUS power distribution units, power transfer unit, utility batteries,
pyrotechnic switching unit, IUS wiring harness and umbilical, and
staging connectors.  The IUS avionics subsystem distributes
electrical power to the IUS and spacecraft interface connector for
all mission phases from prelaunch to spacecraft separation.  The IUS
system distributes orbiter power to the spacecraft during ascent and
on-orbit phases.  ASE batteries supply power to the spacecraft if
orbiter power is interrupted.  Dedicated IUS and spacecraft batteries
ensure uninterrupted power to the spacecraft after deployment from
the orbiter.  The IUS will also accomplish an automatic power-down if
high-temperature limits are experienced before the orbiter payload
bay doors are opened.  Dual buses ensure that no single power system
failure can disable both A and B channels of avionics.  For the IUS
two-stage vehicle, four batteries (three avionics and one spacecraft)
are carried in the IUS first stage.  Five batteries (two avionics,
two utility and one spacecraft) supply power to the IUS second stage
after staging.  The IUS battery complement can be changed to adapt to
mission-unique requirements and to provide special spacecraft
requirements.  Redundant IUS switches transfer the power input among
spacecraft, ground support equipment, ASE and IUS battery sources.
 
Stage 1 to stage 2 IUS separation is accomplished via redundant
low-shock ordnance devices that minimize the shock environment on the
spacecraft.  The IUS provides and distributes ordnance power to the
IUS/spacecraft interface for firing spacecraft ordnance devices in
two groups of eight initiators: a prime group and a backup group.
Four separation switches or breakwires provided by the spacecraft are
monitored by the IUS telemetry system to verify spacecraft separation.
 
IUS SOLID ROCKET MOTORS.
 
The two-stage IUS vehicle incorporates a large SRM and a small SRM.
These motors employ movable nozzles for thrust vector control.  The
nozzles are positioned by redundant electromechanical actuators
permitting up to 4 degrees of steering on the large motor and 7
degrees on the small motor.  Kevlar filament-wound cases provide high
strength at minimum weight.  The large motor's 145-second thrusting
period is the longest ever developed for space.  Variations in user
mission requirements are met by tailored propellant off-loading or
on-loading.  The small motor can be flown either with or without its
extendible exit cone, which provides an increase of 14.5 seconds in
the delivered specific impulse of the small motor.
 
IUS REACTION CONTROL SYSTEM.
 
The IUS RCS is a hydrazine monopropellant positive-expulsion system
that controls the attitude of the IUS and spacecraft during IUS coast
periods, roll during SRM thrustings and delta velocity impulses for
accurate orbit injection.  Valves and thrusters are redundant, which
permits continued operation with a minimum of one failure.
 
The IUS baseline includes two RCS tanks with a capacity of 120 pounds
of hydrazine each.  Production options are available to add a third
tank or remove one tank if required.  To avoid spacecraft
contamination, the IUS has no forward-facing thrusters.  The system
is also used to provide the velocities for spacing between multiple
spacecraft deployments and for a collision/contamination avoidance
maneuver after spacecraft separation.
 
The RCS is a sealed system that is serviced before spacecraft mating.
 Propellant is isolated in the tanks with pyrotechnic squib-operated
valves that are not activated until 10 minutes after IUS deployment
from the orbiter.  The tank and manifold safety factors are such that
no safety constraints are imposed on operations in the vicinity of
the serviced tanks.
 
IUS-TO-SPACECRAFT INTERFACES.
 
The spacecraft is attached to the IUS at a maximum of eight
attachment points.  They provide substantial load-carrying capability
while minimizing thermal transfer across the interface.
 
Power and data transmission to the spacecraft are provided by several
IUS interface connectors.  Access to these connectors can be provided
on the spacecraft side of the interface plane or through the access
door on the IUS equipment bay.
 The IUS provides a multilayer insulation blanket of aluminized Kapton
with polyester net spacers and an aluminized beta cloth outer layer
across the IUS and spacecraft interface.  All IUS thermal blankets
are vented toward and into the IUS cavity.  All gases within the IUS
cavity are vented to the orbiter payload bay.  There is no gas flow
between the spacecraft and the IUS.  The thermal blankets are
grounded to the IUS structure to prevent electrostatic charge buildup.
 
FLIGHT SEQUENCE.
 
After the orbiter payload bay doors are opened in Earth orbit, the
orbiter maintains a preselected attitude to fulfill payload thermal
requirements and constraints except during those operations that
require special attitudes (e.g., orbiter inertial measurement unit
alignments, RF communications and deployment operations).
 
On-orbit predeployment checkout is followed by an IUS command link
check and spacecraft RF command check, if required.  The state vector
is uplinked to the orbiter for trim maneuvers the orbiter performs.
The state vector is transferred to the IUS.
 
The forward ASE payload retention latch actuator is released and the
aft frame ASE electromechanical tilt actuator tilts the IUS and
spacecraft combination to 29 degrees.  This extends the spacecraft
into space just outside the orbiter payload bay, which allows direct
communication with Earth during systems checkout.  The orbiter is
then maneuvered to the deployment attitude.  If a problem develops
within the spacecraft or IUS, they can be restowed.
 
Before deployment, the flight crew switches the spacecraft's
electrical power source from orbiter power to IUS internal power.
Verification that the spacecraft is on IUS internal power and that
all IUS and spacecraft predeployment operations have been
successfully completed will be ascertained by evaluating data
contained in the IUS and spacecraft telemetry.  IUS telemetry data
are evaluated by the IUS Mission Control Center at Sunnyvale, Calif.,
and the spacecraft data by the spacecraft Control Center.  Analysis
of the telemetry will result in a go/no-go decision for IUS and
spacecraft deployment from the orbiter.
 
When the orbiter flight crew is given a go decision, the orbiter
flight crew will activate the ordnance that separates the IUS and
spacecraft's umbilical cables.  The flight crew will then command the
electromechanical tilt actuator to raise the tilt table to a
59-degree deployment position.  The orbiter's RCS thrusters are
inhibited, and the Super*zip ordnance separation device physically

separates the IUS and spacecraft combination from the tilt table.
Compressed springs provide the force to jettison the IUS and
spacecraft from the orbiter payload bay at approximately 0.4 foot per
second.  The IUS and spacecraft are deployed in the shadow of the
orbiter or in Earth eclipse.  The tilt table is lowered to minus 6
degrees after deployment.  Approximately 19 minutes after deployment,
the orbiter's orbital maneuvering system engines are ignited to
separate the orbiter from the IUS and spacecraft.
 
The IUS and spacecraft are now controlled by computers on board the
IUS.  Approximately 10 minutes after the IUS and spacecraft are
ejected from the orbiter, the IUS onboard computers send out discrete
signals that are used by the IUS or spacecraft to begin mission
sequence events.  All subsequent operations are sequenced by the IUS
computer from transfer orbit injection through spacecraft separation
and IUS deactivation.  Following RCS activation, the IUS maneuvers to
the required thermal attitude and performs any required spacecraft
thermal control maneuver.
 
Approximately 45 minutes after IUS and spacecraft ejection from the
orbiter, the SRM-1 ordnance inhibits are removed.  At this time, the
bottom of the orbiter is oriented toward the IUS and spacecraft to
protect the orbiter windows from the IUS SRM-1 plume.  The IUS will
then recompute SRM-1 ignition time and maneuver to the proper
attitude for the SRM-1 thrusting period.  When the transfer orbit or
planetary trajectory injection opportunity is reached, the IUS
computer will enable and apply ordnance power, arm the safe and arm
devices and ignite the first-stage SRM.  The SRM-1 thrusting period
lasts approximately 145 seconds to provide sufficient thrust for the
orbit transfer phase of a geosynchronous mission or to provide the
predetermined contribution of thrust for planetary trajectory for
planetary missions.  The IUS first stage and interstage are separated
from the second stage before the IUS reaches the apogee point of its
trajectory for geosynchronous missions.
 
If sufficient coast time is available during the coast phase, the IUS
can perform the maneuvers required by the spacecraft for thermal
protection or communication reasons.
 
For geosynchronous missions, the second-stage motor is ignited at
apogee and its thrusting period lasts approximately 103 seconds,
which provides the final injection to geosynchronous orbit.
Planetary missions use SRM-2 thrust to obtain proper injection
energy.  The IUS then supports spacecraft separation and performs a
final collision and contamination avoidance maneuver before
deactivating its subsystems.
 
Boeing's propulsion team member, Chemical Systems Division of United
Technologies, designed and tests the two solid rocket motors.
Supporting Boeing in the avionics area are TRW, Cubic and the
Hamilton Standard Division of United Technologies.  TRW and Cubic
provide IUS telemetry, tracking and command subsystem hardware.
Hamilton Standard provides guidance system hardware support.  Delco,
under subcontract to Hamilton Standard, provides the avionics
computer.
 
In addition to the actual flight vehicles, Boeing is responsible for
the development of ground support equipment and software for the
checkout and handling of the IUS vehicles from factory to launch pad.
 
Boeing also integrates the IUS with various satellites and joins the
satellite with the IUS, checks out the configuration and supports
launch and mission control operations for both the Air Force and
NASA.  Boeing also develops airborne support equipment to support the
IUS in the space shuttle and monitors it while it is in the orbiter
payload bay.
 
The IUS, without the two SRMs, is fabricated and tested at the Boeing
Space Center, Kent, Wash.  SRMs are shipped directly from Chemical
Systems Division in California to the eastern launch site at Cape
Canaveral, Fla.  Similarly, the Boeing-manufactured IUS subsystems
are shipped from Washington to the eastern launch site.  IUS/SRM
buildup is done in the Solid Motor Assembly Building and the IUS and
spacecraft are mated in the Vertical Processing Facility at the
Kennedy Space Center.  The combined IUS and spacecraft payload is
installed in the orbiter at the launch pad.  Boeing is building 22
IUS vehicles under its contract with the Air Force.