The Saturn V Instrument Unit is a ring-shaped structure fitted to the top of the Saturn V rocket's third stage (S-IVB) and the Saturn IB's second stage (S-IVB). It was immediately below the SLA (Spacecraft/Lunar Module Adapter) panels that contained the Lunar Module. The Instrument Unit contains the guidance system for the Saturn V rocket. Some of the electronics contained within the Instrument Unit are a digital computer, analog flight control computer, emergency detection system, inertial guidance platform, control accelerometers and control rate gyros. The instrument unit (IU) for Saturn V was designed by NASA at Marshall Space Flight Center (MSFC) and was developed from the Saturn I IU.[1] NASA's contractor to construct the Saturn V Instrument Unit was International Business Machines (IBM).[2]
One of the unused Instrument Units is currently on display at the Steven F. Udvar-Hazy Center in Chantilly, Virginia. The plaque for the Unit has the following description:
The Saturn V rocket, which sent astronauts to the Moon, used inertial guidance, a self-contained system that guided the rocket's trajectory. The rocket booster had a guidance system separate from those on the command and lunar modules. It was contained in an instrument unit like this one, a ring located between the rocket's third stage and the command and lunar modules. The ring contained the basic guidance system components—a stable platform, accelerometers, a digital computer, and control electronics—as well as radar, telemetry, and other units. The instrument unit's stable platform was based on an experimental unit for the German V-2 rocket of World War II. The Bendix Corporation produced the platform, while IBM designed and built the unit's digital computer.
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There was no Instrument Unit for Saturn I Block I boosters (SA-1 to SA-4). Guidance and control equipment was carried in canisters on top of the S-I first stage, and included the ST-90 stabilized platform, made by Ford Instrument Company and used in the Redstone missile.[3]
The IU made its debut with SA-5, the first Saturn I Block II launch. The first version of the IU was 154 inches (3,900 mm) in diameter and 58 inches (1,500 mm) high, and was both designed and built by MSFC. Guidance, telemetry, tracking and power components were contained in four pressurized, cylindrical containers attached like spokes to a central hub.[4]
MSFC flew version 2 of the IU on SA-8, 9 and 10. Version 2 was the same diameter as version 1, but only 34 inches (860 mm) high. Instead of pressurized containers, the components were hung on the inside of the cylindrical wall, achieving a reduction in weight.[5]
The last version, number 3, was 260 inches (6,600 mm) in diameter and 36 inches (910 mm) tall. It was designed by MSFC but manufactured by IBM in their factory at Huntsville, and flew on all Saturn IB and Saturn V launches. This is the version that is on display in Washington, Huntsville, Houston, and the Apollo/Saturn V Center.
| PROGRAM | VEHICLE | MISSION | LAUNCH DATE | PAD | IU VERSION |
|---|---|---|---|---|---|
| Saturn I | SA-1 | SA-1 | 27-Oct-61 | 34 | - |
| Saturn I | SA-2 | SA-2 | 25-Apr-62 | 34 | - |
| Saturn I | SA-3 | SA-3 | 16-Nov-62 | 34 | - |
| Saturn I | SA-4 | SA-4 | 28-Mar-63 | 34 | - |
| Saturn I | SA-5 | SA-5 | 29-Jan-64 | 37B | 1 |
| Saturn I | SA-6 | A-101 | 28-May-64 | 37B | 1 |
| Saturn I | SA-7 | A-102 | 18-Sep-64 | 37B | 1 |
| Saturn I | SA-9 | A-103 | 16-Feb-65 | 37B | 2 |
| Saturn I | SA-8 | A-104 | 25-May-65 | 37B | 2 |
| Saturn I | SA-10 | A-105 | 30-Jul-65 | 37B | 2 |
| Saturn IB | SA-201 | AS-201 | 26-Feb-66 | 34 | 3 |
| Saturn IB | SA-203 | AS-203 | 5-Jul-66 | 37B | 3 |
| Saturn IB | SA-202 | AS-202 | 25-Aug-66 | 34 | 3 |
| Saturn V | SA-501 | Apollo 4 | 9-Nov-67 | 39A | 3 |
| Saturn IB | SA-204 | Apollo 5 | 22-Jan-68 | 37B | 3 |
| Saturn V | SA-502 | Apollo 6 | 4-Apr-68 | 39A | 3 |
| Saturn IB | SA-205 | Apollo 7 | 11-Oct-68 | 34 | 3 |
| Saturn V | SA-503 | Apollo 8 | 21-Dec-68 | 39A | 3 |
| Saturn V | SA-504 | Apollo 9 | 3-Mar-69 | 39A | 3 |
| Saturn V | SA-505 | Apollo 10 | 18-May-69 | 39B | 3 |
| Saturn V | SA-506 | Apollo 11 | 16-Jul-69 | 39A | 3 |
| Saturn V | SA-507 | Apollo 12 | 14-Nov-69 | 39A | 3 |
| Saturn V | SA-508 | Apollo 13 | 11-Apr-70 | 39A | 3 |
| Saturn V | SA-509 | Apollo 14 | 31-Jan-71 | 39A | 3 |
| Saturn V | SA-510 | Apollo 15 | 26-Jul-71 | 39A | 3 |
| Saturn V | SA-511 | Apollo 16 | 16-Apr-72 | 39A | 3 |
| Saturn V | SA-512 | Apollo 17 | 7-Dec-72 | 39A | 3 |
| Saturn V | SA-513 | Skylab 1 | 14-May-73 | 39A | 3 |
| Saturn IB | SA-206 | Skylab 2 | 25-May-73 | 39B | 3 |
| Saturn IB | SA-207 | Skylab 3 | 28-Jul-73 | 39B | 3 |
| Saturn IB | SA-208 | Skylab 4 | 16-Nov-73 | 39B | 3 |
| Saturn IB | SA-210 | ASTP | 15-Jul-75 | 39B | 3 |
Saturn Apollo flight profiles varied considerably by mission.[7][8][9] All missions began, however, with liftoff under power of the first stage. To more smoothly control engine ignition, thrust buildup and liftoff of the vehicle, restraining arms provided support and hold down at four points around the base of the S-IC stage. A gradual controlled release was accomplished during the first six inches of vertical motion.
After clearing the launch tower, a flight program stored in the launch vehicle digital computer (LVDC) commanded a roll of the vehicle to orient it so that the subsequent pitch maneuver pointed the vehicle in the desired azimuth. The roll and pitch commands were controlled by the stored program, and were not affected by navigation measurements. Until the end of the S-IC burn, guidance commands were functions only of time.
First stage cutoff and stage separation were commanded when the IU received a signal that the tank's fuel level had reached a predetermined point. Guidance during the second and third stage burns depended both on time and navigation measurements, in order to achieve the target orbit using the minimum fuel.
Second stage engine cutoff was commanded by the IU at a pre-determined fuel level, and the stage was separated. By this time, the vehicle had reached its approximate orbital altitude, and the third stage burn was just long enough to reach a circular parking orbit.
During manned Apollo missions, the vehicle coasted in Earth orbit for 2-4 passes as the crew performed checks of systems status and other tasks, and as ground stations tracked the vehicle. During the hour and a half after launch, tracking stations around the world had refined estimates of the vehicle's position and velocity, collectively known as its state vector. The latest estimates were relayed to the guidance systems in the IU, and to the Command Module Computer in the spacecraft. When the Moon, Earth, and vehicle were in the optimum geometrical configuration, the third stage was reignited to put the vehicle into a translunar orbit. For Apollo 15, for example, this burn lasted 5 minutes 55 seconds.
After translunar injection came the maneuver called transposition, docking, and extraction. This was under crew control, but the IU held the S-IVB/IU vehicle steady while the Command/Service Module (CSM) first separated from the vehicle, rotated 180 degrees, and returned to dock with the Lunar Module (LM). When the CSM and LM had "hard docked" (connected by a dozen latches), the rearranged spacecraft separated from the S-IVB/IU.
The last function of the IU was to command the very small maneuver necessary to keep the S-IVB/IU out of the way of the spacecraft. On some missions the S-IVB/IU went into high Earth or Solar orbit, while on others it was crashed into the Moon; seismometers were left on the Moon during Apollo 11, 12, 14, 15, and 16, and the S-IVB/IUs of Apollo 13, 14, 15, 16, and 17 were directed to crash. These impacts provided impulses that were recorded by the seismometer network to yield information about the geological structure of the Moon.
The IU consists of six subsystems: structure, guidance and control, environmental control, emergency detection, radio communications (for telemetry, tracking, and command), and power.
The basic IU structure is a short cylinder, 36 inches high and 260 inches (6,600 mm) in diameter, fabricated of an aluminum alloy honeycomb sandwich material 0.95 inches (24 mm) 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 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.
The Saturn V launch vehicle was guided by navigation, guidance, and control equipment located in the IU. A space stabilized platform (the ST-124-M3 inertial platform at location 21) measured acceleration and attitude. A launch vehicle digital computer (LVDC at location 19) solved guidance equations, and an analog flight control computer (location 16) issued commands to steer the vehicle.
The attitude of the vehicle was defined in terms of three axes:
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 is fixed several components:
The angular positions of gimbals on their axes were measured by resolvers, which sent their signals to the 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-IUB/IU stage.
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:
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.
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.
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.
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.
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 can initiate a manual abort if an angular-overrate or two-engine-out condition arose.
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.
Approximately 200 parameters were measured on the IU and transmitted to the ground, in order to
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 30x120 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.
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.
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:
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.
These images show the development of the IU. The first four Saturn launches did not have an IU, but used guidance, telemetry and other equipment installed on top of the first stage.
The first IU flew on the fifth Saturn launch, SA-5, and was 12 feet 10 inches (3.91 m) in diameter and 4 feet 10 inches (1.47 m) high. The components it carried were in pressurized containers. This version flew on SA-5, SA-6 and SA-7. The IU carried by missions SA-8, -9, and -10 was only 2 feet 10 inches (0.86 m) high, and was not pressurized.[11]
With the Saturn IB and Saturn V launches, a third version was used, 21.6 feet (6.6 m) in diameter and 3 feet (0.91 m) high. Comparison of these photographs of the Instrument Unit shows that the configuration of components carried by this version changed, depending on the mission. Some equipment was deleted (e.g. the Azusa tracking system was deleted from later IUs), some equipment was added (e.g. a fourth battery for longer missions), and other components were moved around.
These images also show that some components (e.g. batteries, the ST-124 inertial platform) were installed in the IU after it had been stacked in the VAB on top of the S-IVB third stage.