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The PAM is used to boost various satellites to geosynchronous transfer orbit (22,300 miles) or other higher energy orbits after deployment from the space shuttle vehicle. There are three versions of the PAM that can be used aboard the space shuttle. The PAM-D is capable of launching satellites weighing up to 2,750 pounds. The PAM-DII is for satellites weighing up to 4,150 pounds. The PAM-A is for satellites weighing up to 4,400 pounds. The PAMs are also used for non-geosynchronous transfer orbits.
The PAM's deployable (expendable) stage consists of a spin-stabilized, solid-fueled rocket motor; a payload attach fitting to mate with the unmanned spacecraft; and the necessary timing, sequencing, power and control assemblies.
The reusable airborne support equipment consists of the cradle structure for mounting the deployable system in the space shuttle orbiter payload bay; a spin system to provide the stabilizing rotation; a separation system to release and deploy the stage and unmanned spacecraft; and the necessary avionics to control, monitor and power the system.
The PAM stages are supported through the spin table at the base of the motor and through restraints at the PAF. The forward restraints are retracted before deployment.
The PAM-Ds and DIIs have a sunshield that provides thermal protection of the PAM/unmanned spacecraft when the space shuttle orbiter payload bay doors are open on orbit.
The PAM's ASE consists of all the reusable hardware elements that are required to mount, support, control, monitor, protect and operate the PAM's expendable hardware and unmanned spacecraft from lift-off to deployment from the space shuttle. It will also provide the same functions for the safing and return of the stage and spacecraft in case of an aborted mission. The ASE is designed to be as self-contained as possible, thereby minimizing dependence on orbiter or flight crew functions for its operation. The major ASE elements include the cradle for structural mounting and support, the spin table and drive system, the avionics system to control and monitor the ASE and the PAM vehicle, and the thermal control system.
The cradle assembly provides the vertical structural mounting support for the PAM/unmanned spacecraft assembly in the orbiter payload bay. The cradle is 15 feet wide. The length of the cradle is 93 inches static and 96 inches dynamic. The open-truss-structure cradle is constructed of machined aluminum frame sections and chrome-plated steel longeron and keel trunnions.
The spacecraft-to-cradle lateral loads are reacted by forward retractable reaction fittings between the PAF and cradle, which are driven by redundant dc electrical motors. After the reaction fittings are retracted, the spin table is free to spin the PAM and unmanned spacecraft when commanded.
The spin table consists of three subsystems-spin, separation and electrical interface. The spin subsystem consists of the spin table, the spin bearing, the rotating portion of the spin table, a gear and gear support ring, two redundant drive motors, a despin braking device, and a rotational index and locking mechanism. The separation subsystem includes four compression springs mounted on the outside of the rotating spin table, each with an installed preload of 1,400 pounds and a Marman-type clamp band assembly.
The electrical interface subsystem consists of a slip ring assembly to carry electrical circuits for the PAM and spacecraft across the rotating spin bearing. The electrical wiring from the slip ring terminates at electrical disconnects at the spin table separation point. The slip ring assembly is used to carry safety-critical command and monitor functions and those commands required before separation from the spin table.
The system provides a capability for spin rates between 45 and 100 rpm. Upon command, the spin table will be spun up to the nominal rpm by two electric motors, either of which can produce the required torque. When the spin table rpm has been verified and the proper point in the parking orbit has been reached, redundant, debris-free, explosive bolt cutters are fired upon command from the electrical ASE to separate the band clamp (which is mechanically retained on the spin table), and the springs provide the thrust to attain a separation velocity of approximately 3 feet per second.
In case of an abort mode after spinup, the multiple-disc stack friction-type braking device will despin the PAM and unmanned spacecraft assembly, and the spin drive motor will slowly rotate the assembly until the solenoid-operated indexing and locking device is engaged. Upon confirmation by the ASE that the spin table is properly aligned and locked, the restraint arms will be re-engaged.
The PAM-D PAF structure is a machined forging and provides the subsystem mounting installations and mounts on the forward ring of the motor case. PAM-DII offers two structural designs for the PAF, one of aluminum monocoque and the other of graphite epoxy. The two cradle reaction fittings provide structural support to the forward end of the PAM stage and unmanned spacecraft and transmit loads to the ASE cradle structure. The forward interface of the PAF provides the spacecraft mounting and separation system. One Marman-type clamp band is preloaded as required to properly support the spacecraft, and separation is achieved by redundant bolt cutters. Four separation springs, mounted inside the PAF, provide the impetus for clear separation. For PAM-D, installed preload for each spring is approximately 200 pounds with a spring stroke of 5.25 inches, providing a spacecraft separation velocity of about 3 feet per second. The electrical interface connectors between the PAM and the spacecraft are mounted on brackets on opposite sides of the PAF. Other subsystems mounted on the PAF include the redundant safe and arm device for motor ignition and telemetry components (if desired) and the optional S-band transmitter.
The electrical ASE minimizes the number of operations to be performed by the flight crew so that greater attention can be paid to monitoring functions that are critical to safety and reliability.
Flight crew control functions include system power on, solid rocket motor arming, deployment ordnance arming, emergency deployment and sequence control assembly control.
The electrical ASE performs sunshield opening and closing, control and monitoring of restraint withdrawal, spin table spin and deployment functions; arms (and disarms, if necessary) the SRM; controls and monitors the PAM vehicle's electrical sequencing system (and telemetry system, when used); and provides wiring to carry required spacecraft functions. And, as a mission option, it provides control and monitoring of spacecraft systems.
The PAM thermal control system alleviates severe thermal stresses on both the unmanned spacecraft and the PAM system. The system consists of thermal blankets mounted on the cradle to provide thermal protection and a sunshield mounted on the cradle to control the solar input to and heat loss from the payload when the orbiter payload bay doors are open.
Thermal blankets consisting of multilayered insulation are mounted to the forward and aft sides of the cradle. The orbiter payload bay liner provides thermal protection on the sides and the bottom.
The sunshield consists of multilayered Mylar lightweight insulation supported on a tubular frame. The sunshield panels on the sides are stationary. The portion of the shield covering the top of the PAM/unmanned spacecraft is a clamshell structure that remains closed to protect against thermal extremes when the orbiter payload bay doors are open. The sunshield resembles a two-piece baby buggy canopy. The clamshell is opened by redundant electric rotary actuators operating a control-cable system.
The PAM-D/unmanned spacecraft sunshield accommodates spacecraft up to 86 inches in diameter. The PAM-DII/unmanned spacecraft sunshield accommodates spacecraft up to 115 inches in diameter.
The PAM expendable hardware consists of the Thiokol solid-fueled rocket motor, the payload attach fitting, and its functional system.
The avionics and other systems are common to both the PAM-D and PAM-DII. The PAM-DII, however, required revision or new design of the cradle and structural parts due to the larger volume required and heavier weight than the PAM-D. The PAM-DII utilizes a Thiokol Star 63D motor for greater performance.
PAM-DII functions in the same manner as PAM-D and utilizes all existing PAM-D parameters of preparation, interface, sequence and operation. PAM-DII's thermal, mechanical, and electrical interface and environment are comparable to the PAM-D's.
The payload assist modules are designed and built by McDonnell Douglas Astronautics Co., Huntington Beach, Calif.
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.
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.
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.
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.
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.
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.
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