The IU consists of six subsystems: structure, guidance and control, environmental control, emergency detection, radio communications (for telemetry, tracking, and command), and power.
Structure The basic IU structure is a short cylinder, 36 inches high and in diameter, fabricated of an aluminum alloy honeycomb sandwich material thick. The cylinder is manufactured in three 120-degree segments, which are joined by splice plates into an integral structure. The top and bottom edges are made from extruded aluminum channels bonded to the honeycomb sandwich. This type of construction was selected for its high strength to weight ratio, acoustical insulation, and thermal conductivity properties. The IU supported the components mounted on its inner wall and the weight of the Apollo spacecraft above (the Lunar Module, the Command Module, the Service Module, and the Launch Escape Tower). To facilitate handling the IU before it was assembled into the Saturn, the fore and aft protective rings, 6 inches tall and painted blue, were bolted to the top and bottom channels. These were removed in the course of stacking the IU into the Saturn vehicle. The structure was manufactured by North American Rockwell in Tulsa, Oklahoma. Edward A. Beasley was the I.U. Program Manager. The IU is divided into 24 locations, which are marked on the interior by numbers 1–24 on the aluminum surface just above the blue flange.
Guidance, navigation, and control The Saturn V launch vehicle's GN&C architecture featured Guidance, Navigation, and Control equipment located in the IU. The
ST-124-M3 inertial platform contains three
gimbals: the outer gimbal (which can rotate 360° about the roll or X axis of the vehicle), the middle gimbal (which can rotate ±45° about the yaw or Z axis of the vehicle), and the inner or inertial gimbal (which can rotate 360° about the pitch or Y axis of the vehicle). The inner gimbal is a platform to which are fixed several components: • Two vertical alignment pendulums sent signals before launch to ground support equipment, which generated signals to the platform gyro torque generators to level the inner gimbal. The vertical alignment system levelled the platform to an accuracy of ±2.5
arc seconds. • Two
prisms, one fixed and one
servo-driven, were used with an external
theodolite which sighted through the viewport outside location 21 to set the
azimuth of the inner gimbal before launch. The azimuth could be set to an accuracy of ±5 arc seconds. • Three single-degree-of-freedom
gyroscopes have their input axes aligned along an
orthogonal inertial coordinate system. Three signal generators, fixed to the output axis of each gyro, generated electrical signals proportional to the
torque disturbances. The signals were transmitted through the servo electronics which terminated in the gimbal pivot servotorque motors. The servoloops maintained the inner gimbal rotationally fixed in inertial space. That is, while the vehicle rolled, pitched, and yawed, the inner gimbal remained in the same attitude to which it was set just before launch. Though it was being translated during the launch and orbit process, it was rotationally fixed. • Three integrating
accelerometers measured the three components of velocity resulting from vehicle propulsion. The accelerometer measurements were sent through the launch vehicle data adapter (LDVA at location 19) to the LVDC. In the LVDC the accelerometer measurements were combined with the computed gravitational acceleration to obtain velocity and position of the vehicle. The angular positions of gimbals on their axes were measured by resolvers, which sent their signals to the
Launch Vehicle Data Adaptor (LVDA). The LVDA was the input/output device for the LVDC. It performed the necessary processing of signals to make these signals acceptable to the LVDC. The instantaneous attitude of the vehicle was compared with the desired vehicle attitude in the LVDC. Attitude correction signals from the LVDC were converted into control commands by the flight control computer. The required thrust direction was obtained by gimbaling the engines in the propelling stage to change the thrust direction of the vehicle. Gimbaling of these engines was accomplished through
hydraulic actuators. In the first and second stages (S-IC and S-II), the four outboard engines were gimbaled to control roll, pitch, and yaw. Since the third (S-IVB) stage has only one engine, an auxiliary propulsion system was used for roll control during powered flight. The auxiliary propulsion system provides complete attitude control during coast flight of the S-IVB/IU stage.
Control/EDS - Rate Gyro Assembly (RGA) Package Mounted at location 15, the nine (triple-redundant) "Rate Gyroscopes" for angular rates (pitch, roll, and yaw), were used for "real time" attitude stabilization and the Emergency Detection System (EDS). The angular rate signals were sent to the Control Signal Processor (CSP) for signal voting and then continuously input (i.e. at a high frequency, in real time) to the analog Flight Control Computer (FCC), a task that would of overwhelmed the digital computer. The FCC's damping computations were necessary for stabilizing the vehicle's angular motion during flight. The final command to the engine actuators was a complex, filtered sum of several analog inputs. The FCC's analog components performed complex weighting, integration, and filtering instantaneously and continuously—at a speed necessary for vehicle stability that the 1960s-era digital computers (like the LVDC) could not match. Thus the digital computer provided the strategic steering while the analog FCC provided the tactical, high-speed stability control. The engine-out sensors and the cables for monitoring structural integrity were distributed throughout the rocket stages, but the final decision-making logic to act on those inputs resided within the EDS components in the IU. This component housed dedicated solid-state logic circuitry that formed the core of the automatic abort system. It functioned on preset abort rules (e.g., "two or more engine failures" or "excessive vehicle rate" of pitch, roll, or yaw).
Emergency detection The emergency detection system (EDS) sensed initial development of conditions in the flight vehicle during the boost phases of flight which could cause vehicle failure. The EDS reacted to these emergency situations in one of two ways. If breakup of the vehicle were imminent, an automatic abort sequence would be initiated. If, however, the emergency condition were developing slowly enough or were of such a nature that the flight crew can evaluate it and take action, only visual indications were provided to the flight crew. Once an abort sequence had been initiated, either automatically or manually, it was irrevocable and ran to completion. The EDS was distributed throughout the vehicle and includes some components in the IU. There were nine EDS rate gyros installed at location 15 in the IU. Three gyros monitored each of the three axes (pitch, roll and yaw), providing triple redundancy. The control signal processor (location 15) provided power to and received inputs from the nine EDS rate gyros. These inputs were processed and sent to the EDS distributor (location 14) and to the flight control computer (location 16). The EDS distributor served as a junction box and switching device to furnish the spacecraft display panels with emergency signals if emergency conditions existed. It also contained relay and diode logic for the automatic abort sequence. An electronic timer (location 17) was activated at liftoff and 30 seconds later energized relays in the EDS distributor which allowed multiple engine shutdown. This function was inhibited during the first 30 seconds of launch, to preclude the vehicle falling back into the launch area. While the automatic abort was inhibited, the flight crew could initiate a manual abort if an angular-overrate or two-engine-out condition arose.
Environmental control The environmental control system (ECS) maintains an acceptable operating environment for the IU equipment during preflight and flight operations. The ECS is composed of the following: • The thermal conditioning system (TCS), which maintains a circulating coolant temperature to the electronic equipment of 59° ± 1 °F (15 ± 0.6 °C). • Preflight purging system, which maintains a supply of temperature- and pressure-regulated mixture of air and gaseous nitrogen (air/GN2) in the IU/S-IVB equipment area. • Gas bearing supply system, which furnishes GN2 to the ST-124-M3 inertial platform gas bearings. • Hazardous gas detection sampling equipment which monitors the IU/S-IVB forward interstage area for the presence of hazardous vapors
Thermal conditioning Thermal conditioning panels, also called cold plates, were located in both the IU and S-IVB stage (up to sixteen in each stage). Each cold plate contains tapped bolt holes in a grid pattern which provides flexibility of component mounting. The cooling fluid circulated through the TCS was a mixture of 60 percent
methanol and 40 percent demineralized
water by weight. Each cold plate was capable of dissipating at least 420 watts. During flight, heat generated by equipment mounted on the cold plates was dissipated to space by a
sublimation heat exchanger. Water from a reservoir (water accumulator) was exposed to the low temperature and pressure environment of space, where it first freezes and then sublimates, taking heat from the heat exchanger and transferring it to the water molecules which escape to space in gaseous state. Water/methanol was cooled by circulation through the heat exchanger. ==== Preflight air/GN2 purge
system ==== Before flight, ground support equipment (GSE) supplies cooled, filtered ventilating air to the IU, entering via the large duct in the middle of the umbilical panel (location 7), and branching into two ducts at the top that are carried around the IU in the cable rack. Downward pointing vents from these ducts release ventilating air to the interior of the IU. During fueling, gaseous nitrogen was supplied instead of air, to purge any propellant gases that might otherwise accumulate in the IU.
Gas bearing supply To reduce errors in sensing attitude and velocity, designers cut friction to a minimum in the platform gyros and accelerometers by floating the bearings on a thin film of dry nitrogen. The nitrogen was supplied from a sphere holding 2 cu ft (56.6 L) of gas at 3,000
psig (pounds per square inch gauge, i.e. psi above one atmosphere) (20,7
MPa). This sphere is 21 inches (0,53 m) in diameter and is mounted at location 22, to the left of the ST-124-M3. Gas from the supply sphere passes through a filter, a pressure regulator, and a heat exchanger before flowing through the bearings in the stable platform.
Hazardous gas detection The hazardous gas detection system monitors the presence of hazardous gases in the IU and S-IVB stage forward compartments during vehicle fueling. Gas was
sampled at four locations: between panels 1 and 2, 7 and 8, 13 and 14, and 19 and 20. Tubes lead from these locations to location 7, where they were connected to ground support equipment (external to the IU) which can detect hazardous gases.
Radio communications The IU communicated by radio continually to ground for several purposes. The measurement and telemetry system communicated data about internal processes and conditions on the Saturn V. The tracking system communicated data used by the Mission Ground Station (MGS) to determine vehicle location. The radio command system allowed the MGS to send commands up to the IU.
Measuring and telemetry Approximately 200 parameters were measured on the IU and transmitted to the ground, in order to • Assist in the checkout of the launch vehicle prior to launch, • Determine vehicle condition and to verify received commands during flight, and • Facilitate postflight analysis of the mission. Parameters measured include
acceleration,
angular velocity,
flow rate,
position,
pressure,
temperature,
voltage,
current,
frequency, and others.
Sensor signals were conditioned by
amplifiers or
converters located in measuring racks. There are four measuring racks in the IU at locations 1, 9, and 15 and twenty signal conditioning modules in each. Conditioned signals were routed to their assigned telemetry channel by the measuring distributor at location 10. There were two telemetry links. In order for the two IU telemetry links to handle approximately 200 separate measurements, these links must be shared. Both frequency sharing and time sharing
multiplexing techniques were used to accomplish this. The two
modulation techniques used were
pulse-code modulation/frequency modulation (PCM/FM) and frequency modulation/frequency modulation (FM/FM). Two Model 270 time sharing
multiplexers (MUX-270) were used in the IU telemetry system, mounted at locations 9 and 10. Each one operates as a 30×120 multiplexer (30 primary channels, each sampled 120 times per second) with provisions for submultiplexing individual primary channels to form 10 subchannels each sampled at 12 times per second. Outputs from the MUX-270 go to the PCM/DDAS assembly model 301 at location 12, which in turn drives the 245.3 MHz PCM VHF transmitter. The FM/FM signals were carried in 28 subcarrier channels and transmitted by a 250.7 MHz FM transmitter. Both the FM/FM and the PCM/FM channels were coupled to the two telemetry antennas on opposite sides of the IU outside locations 10 and 22.
Tracking C-band radar transponders carried by the IU provided tracking data to the ground which were used to determine the vehicle's
trajectory. The transponder received coded or single pulse interrogation from ground stations and transmitted a single-pulse reply in the same frequency band (5.4 to 5.9
GHz). A common
antenna was used for receiving and transmitting. The C-band transponder antennas are outside locations 11 and 23, immediately below CCS PCM omni receive antennas.
Radio command The command communications system (CCS) provided for digital data transmission from ground stations to the LVDC. This communications link was used to update guidance information or command certain other functions through the LVDC. Command data originated in the
Mission Control Center,
Houston, and was sent to remote stations for transmission to the launch vehicle. Command messages were transmitted from the ground at 2101.8 MHz. The received message was passed to the command decoder (location 18), where it was checked for authenticity before being passed to the LVDC. Verification of message receipt was accomplished through the IU PCM telemetry system. The CCS system used five antennas: • A single directional antenna outside location 3–4, • Two omni transmit antennas outside locations 11 and 23, and • Two omni receive antennas outside locations 12 and 24.
Power Power during flight originated with four silver-zinc batteries with a nominal voltage of 28±2 vdc. Battery D10 sat on a shelf at location 5, batteries D30 and D40 were on shelves in location 4, and battery D20 was at location 24. Two power supplies converted the unregulated battery power to regulated 56 vdc and 5 vdc. The 56 vdc power supply was at location 1 and provided power to the ST-124-M3 platform electronic assembly and the accelerometer signal conditioner. The 5 vdc power supply at location 12 provided 5 ±.005 vdc to the IU measuring system. == Gallery ==