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Orbiter operational instrumentation is used to collect, route and process information from transducers and sensors throughout the orbiter and its payloads. This system also interfaces with the solid rocket boosters, external tank and ground support equipment. Over 2,000 data points are monitored, and the data are routed to OI MDMs. The instrumentation system consists of transducers, signal conditioners, two pulse code modulation master units, encoding equipment, two operational recorders, one payload recorder, master timing equipment and onboard checkout equipment.

The OI system senses, acquires, conditions, digitizes, formats and distributes data for display, telemetry, recording and checkout. It provides for PCM recording, voice recording and master timing for onboard systems.

Dedicated signal conditioners convert digital and analog data signals received from the various sensors to a usable form. Some raw sensor data may need to be conditioned for compatibility with a multiplexing system. Signal conditioning provides the multiplexer with compatible inputs. The DSCs provide input from transducer signals, such as frequency, voltage, current, pressure, temperature (variable resistance and thermocouple), displacement (potentiometer), 28- or 5-volt-dc discrete output signals, analog and digital level changes, polarity changes or an ac signal change to a dc signal. The DSCs send these converted signals to the appropriate MDMs and displays and to the C/W system.

MDMs can operate in two ways. As multiplexers, they take data from several sources, convert the data to serial digital signals (a digitized representation of the applied voltage) and interleave the data into a single data stream. As demultiplexers, the MDMs take interleaved serial digital information; separate and convert it to analog, discrete or serial digital; and send each separate signal to its appropriate destination. The payload MDMs generally act as demultiplexers. They take interleaved commands from the orbiter GPCs, separate them and send each command to its appropriate destination, such as payload ground command interface logic.

The OI MDMs generally act only as multiplexers. Upon request from the pulse code modulation master unit, the MDMs send these interleaved streams to the PCMMU through the OI data buses. When the MDM is addressed by the PCMMU, the MDM selects, digitizes and sends the requested data to the PCMMU in serial digital form. The PCMMU/OI MDM interface is based on demand and response: that is, the OI MDMs do not send data to the PCMMU until the PCMMU makes the request.


    The PCMMU receives the data requested from the OI MDMs, downlists data from the GPCs under control of flight software and payload telemetry from the payload data interleaver and Spacelab computers, interleaves the data, formats data according to programmed instructions stored within the PCMMU, and sends the interleaved data to the network signal processor to be mixed with the analog air-to-ground voice data from the audio central control unit for transmission through the S-band PM downlink or Ku-band system return links. Telemetry from the PCMMU also is sent through the NSP to the operational recorders for storage and future payloads to be downlinked on the S-band FM or Ku-band system. OI and payload data collected by the PCMMU are sent to the onboard GPCs for display and monitoring purposes upon request. All data received by the PCMMU is stored in memory and periodically updated.

    The PCMMU has two formatter memories: programmable read only and random access. The read-only memory is programmed only before launch; the random-access memory is reprogrammed several times during flight. The PCMMU uses the format memories to downlink data from the computers and OI MDMs into PCM telemetry data streams.

    Only one of the redundant PCMMUs and NSPs operates at a time. The one used is controlled by the crew through the flight deck display and control panel. The primary port of an MDM operates with PCMMU 1 and the secondary port operates with PCMMU 2.

    The PCMMUs receive a synchronization clock signal from the master timing unit. If this signal is not present, the PCMMU provides its own timing and continues to send synchronization signals to the payload data interleaver and network signal processor.

    The payload data interleaver is programmed onboard from mass memory through the GPCs to select specific data from each payload PCM signal and to store the data within its buffer memory locations.


    The network signal processor is the nucleus of the communication systems. It is responsible for processing and routing commands, telemetry and voice between the orbiter and the ground.

    Commands and voice uplinked to the orbiter are received by the S-band PM uplink or the Ku-band system's forward link. The NSP accepts this digitized, time-division-multiplexed data stream for further processing. Encrypted data are routed to Comsec, decrypted and returned to the NSP. Comsec interfaces with the NSP to provide communication security during Department of Defense and NASA missions. Digitized air-to-ground voice data are demultiplexed from command data and converted to analog signals before being routed to the ACCU. Command data are routed to the GPCs of the data processing system through the flight forward MDM.

    Telemetry and voice are downlinked from the orbiter by the S-band FM, S-band PM and Ku-band systems' return link. The NSP accepts and digitizes analog air-to-ground voice data from the ACCU. The digital voice is multiplexed with telemetry from the PCMMU for real-time transmission to the ground through the S-band PM and Ku-band systems. The multiplexed data also are routed to the operational recorders for later transmission through the S-band FM system or mode 2 of the Ku-band system. Encrypted data are routed to a Comsec encryptor and returned to the NSP before being downlinked.

    Instrumentation equipment, except sensors and selected dedicated signal conditioners, is located in the forward and aft avionics bays. Sensors and dedicated signal conditioners are located throughout the orbiter in areas selected on the basis of accessibility, minimum harness requirements and functional requirements. Effective use of remote data acquisition techniques was considered for optimizing equipment location. The factors that were considered in the determination of equipment location were weight, power, physical size, redundancy and wire density, and length to each compartment and interconnect wiring. Abbreviations used to designate the locations of equipment are as follows: OA refers to operational aft, OF to operational forward, OL to operational left, OR to operational right, OM to operational mid.


    The ground command interface logic, also referred to as the ground command interface logic controller, controls many of the functions of the S-band PM, S-band FM, Ku-band, payload communication and CCTV systems. Commands are sent to the orbiter from the ground through S-band system uplink or Ku-band system forward link. All commands, whether sent on S-band or Ku-band, are routed to the onboard GPC through the NSP and associated MDM. The GCIL takes commands for these systems from the GPCs for ground commands or from the appropriate panel for flight crew commands. The command source is selected at a control panel pnl/cmd switch for each system. If the switch for that system is on pnl, the flight system is in control. If the switch for that system is on cmd, the ground has control.


    The five GPCs in the orbiter are loaded with software (to perform different functions) during different phases of the mission. During ascent and entry, four of the GPCs are loaded with primary avionics software system to provide guidance, navigation and control. One computer is loaded with backup flight system software and can take control of the orbiter if the PASS GN&C computers malfunction. The BFS GPC also contains all software necessary to return the orbiter to Earth, if required. Normal communication commands go through the BFS during ascent and entry phases of the mission. During the on-orbit phase of the mission, one computer is loaded with systems management software to support the communication commands. These two GPCs process antenna-pointing control commands for the S-band and Ku-band antenna systems. They also provide a communication command path from the ground to the required subsystem as well as a telemetry downlist path to the PCMMU to be interleaved with other data for transmission to the ground.


    The master timing unit is a stable crystal-controlled timing source for the orbiter. It sends serial time reference signals to the onboard computers, PCMMUs, payloads and various time display panels. It also provides synchronization for instrumentation payloads and other systems. It includes separate time accumulators for Greenwich Mean Time and mission elapsed time, which can be reset or updated from the ground via uplink through the onboard computer or by the flight crew through the use of the flight deck display and control panel keyboard and CRT time displays.

    The signal flows from the 4.608-MHz oscillators to the output of the GMT and MET accumulators. GMT or MET can be displayed on the mission time displays on panels O3 and A4 by positioning the mission timer switch on the respective panel to GMT or MET . In addition, event time displays are located on panels F7 and A4. Separate time accumulators used for the GMT and MET clocks accumulate time in days, hours, minutes, seconds and milliseconds. The GMT capacity is 366 days, 23 hours, 59 minutes, 59 seconds and 999.875 milliseconds. For MET, the capacity is 365 days, 23 hours, 59 minutes, 59 seconds and 999.875 milliseconds. Before flight, both time displays can be updated and reset by ground equipment or by the flight crew using onboard controls. During flight, the GMT and MET accumulators are updated at a predetermined time by uplink and the onboard computer or by voice command. They are entered through the flight deck display and control panel keyboard and CRT display.


    Two recorders are used for serial recording and dumping of digital voice and PCM data from the OI systems. The recorders normally are controlled by ground command, but they can be commanded by the flight crew through the flight deck display and control panel keyboard or through switches on a recorder panel. Input to the recorders is from the network signal processor, either in the form of 128-kbps PCM data or a 192-kbps composite signal that includes the 128-kbps PCM data and two 32-kbps voice channels. The network signal processor receives the PCM data from the PCMMU and voice signals from the audio control center. In addition, during ascent, operations recorder 1 receives three channels of main engine data at 60 kbps. A single recorder can store and reproduce digital data at many rates.

    The operations recorders can be commanded to dump recorded data from one recorder while continuing to record real-time data on the other. The dump data are sent to the FM signal processor for transmission to the ground station through the S-band FM transmitter on the S-band FM return link or to the Ku-band signal processor. When the ground has verified that the received data are valid, the operations recorders can use that track on the tape to record new data.

    Recorder functions can be summarized as follows:

    Data in, recorder 1

    - Accept three parallel channels of engine data at 60 kbps during ascent.

    - Accept 128 and 192 kbps of interleaved PCM data and voice that serially sequences from track 4 to track 14.

    - Accept real-time data from network signal processor. Recording time is 32 minutes for parallel record and 5.8 hours for serial record on tracks 4 through 14 at a tape speed of 15 inches per second.

    Data in, recorder 2

    - Accept 128 and 192 kbps of interleaved PCM voice and data that serially sequences from track 1 to track 14.

    - Accept real-time data from network signal processor. Recording time is 7.5 hours at a tape speed of 15 inches per second for serial record on 14 tracks.

    Data out, recorder 1

    - Play back in-flight engine interface unit data and network signal processor digital data through S-band FM transponder or Ku-band transmitter.

    - Play back in-flight anomaly PCM data for maintenance recording

    - Play back data serially to ground support equipment to T-0 ground support equipment umbilical

    Data out, recorder 2

    - Play back digital data through S-band FM transponder or Ku-band transmitter in flight.

    - Play back anomaly PCM data in flight for maintenance recording.

    - Play back preflight and postflight data serially to GSE T-0 umbilical.

    Recorder control

    - Recorders are manually controlled from the mission specialist flight deck aft station display and control panel or uplink and onboard computer keyboard.

    - Recorder speeds of 7.5, 15, 24 and 120 inches per second are provided by hardwire program plug direct command.

    The tape recorders contain a minimum of 2,400 feet of 0.5-inch by 1-mil magnetic tape. They operate at a tape speed of 24 inches per second in the recording mode and 120 inches per second in the playback mode.


    The payload recorder records and dumps payload analog and digital data in flight through the S-band or Ku-band transmitter. The recorder can record one of three serial inputs or a maximum of 14 parallel digital or analog inputs or combinations of analog and digital data (up to 14 inputs) from the payload patch panel. In flight, all data dumps are serial; capability for parallel dumps does not exist. The recorded digital data can range from 25.6 kbps to 1.024 Mbps. Analog data inputs can be recorded only in parallel with a bandwidth from 1.9 kHz to 1,600 kHz. There are also 14 selectable tape speeds; however, only four speeds are available per flight. Recorder tape speeds are 15, 30, 60 and 120 inches per second. Total recording time ranges from 56 minutes at 1.024 Mbps to 18 hours 40 minutes at 25.6 kbps.

    The recorder normally is controlled by ground command but can be commanded by the flight crew through the flight deck display and control panel keyboard or through switches on the recorder panel.


    The teleprinter is an interim system designed to transmit written data to the flight crew in orbit until dual TDRSS capability is achieved. It is a modified teletype machine located in a locker in the crew compartment middeck. Electrical power is connected to the dc utility power outlet on panel A15, and audio is connected to the PS comm outlet on panel A15. During launch, the PS comm outlet is used for crew communications; therefore, the audio cable for the teleprinter must be connected on orbit, and the ATU at panel L9 must be reconfigured on orbit. The panel L9 payload station power switch is positioned to aud, A/G 1 is positioned to off , A/G 2 is positioned to t/r , the thumbwheel is positioned to 9, A/A is positioned to off , the two icom switches are positioned to off , and the xmit/icom mode rotary switch is positioned to PTT/PTT. The power switch on the teleprinter locker door is in the on position at launch. Turning the dc utility power switch on panel A15 to on powers up the teleprinter and illuminates a green light on the locker door.

    The teleprinter provides the capability to receive and reproduce text-only data, such as procedures, weather reports and crew activity plan updates or changes, aboard the orbiter from the Mission Control Center.

    The teleprinter uplink requires one to 2.5 minutes per message, depending on the number of lines (up to 66). When the ground has sent a message, a msg rcv yellow light on the teleprinter is illuminated to indicate a message is waiting to be removed. The teleprinter is deactivated before seat ingress on entry day.


    The text and graphics system is used when there is an operational TDRSS. The TAGS consists of a facsimile scanner on the ground that sends text and graphics through the Ku-band communications system to the text and graphics hard copier in the orbiter. The interim teleprinter operates via the S-band PM system when the TDRSS is not operational. The hard copier installed on a dual cold plate in avionics bay 3 of the crew compartment middeck provides an on-orbit capability to transmit text material, maps, schematics, maneuver pads, general messages, crew procedures, trajectory and photographs to the orbiter through the two-way Ku-band link using the TDRSS. It is a high-resolution facsimile system that scans text or graphics and converts the analog scan data into serial digital data. Transmission time for an 8.5- by 11-inch page can vary from approximately one minute to 16 minutes, depending on the hard copy resolution desired.

    The text and graphics hard copier operates by mechanically feeding paper over a fiber-optic CRT and then through a heater-developer. The paper then is cut and stored in a tray, accessible to the flight crew. A maximum of 200 8.5- by 11-inch sheets are stored. The status of the hard copier is indicated by front panel lights and downlink telemetry.

    The hard copier can be powered from the ground or by the crew. The ground can power up the system when the Ku-band control switch is in the command position. When this switch is in the panel position, the crew can power up the system.

    Uplink operations are controlled by the Mission Control Center in Houston. Mission Control powers up the hard copier and then sends the message. In the onboard system, light-sensitive paper is exposed, cut and developed. The message is then sent to the paper tray, where it is retrieved by the flight crew.


    The closed-circuit television system is used primarily to support on-orbit activities that require visual feedback to the crew. The CCTV system also provides the capability to document on-orbit activities and vehicle configurations for permanent record or for real-time transmission to the ground. The CCTV system can be controlled by both onboard and remote uplink commands. The CCTV system is a standard monochrome (black and white) system with an optional color capability by means of interchangeable camera lenses. Color scenes are not available on the onboard monitors because of hardware restrictions; however, color scenes are available on recorded and downlinked video.

    Video inputs to the CCTV system are available from cameras mounted at several locations in the payload bay and on the remote manipulator system arm, which is mission dependent.

    Cameras are also used in the middeck and the flight deck of the crew compartment. The video signals from these cameras can be viewed by the crew on one of two monitors in the flight deck, then sent to a video tape recorder for later viewing or to the ground on the S-band FM or Ku-band systems.

    The video control unit controls power and input/output configuration, including camera control of pan, tilt, zoom, focus, iris operations and synchronization for the CCTV system. It processes video and data inputs from cameras and video tape recorder channels and routes them to monitors, the VTR or the ground by downlink.

    Typical uses of the CCTV system for monitoring and recording mission activities include payload bay door operations, remote manipulator system operations, experiment operations, rendezvous and stationkeeping operations, and onboard crew activities.

    The CCTV system consists of the video control unit, television cameras, VTR and onboard television monitors 1 and 2.

    Audio from orbiter intercom channels A and B can be interleaved with the video output, along with Greenwich Mean Time data from the orbiter master timing unit. The VCU also generates a standard test pattern for use in adjusting the monitors and verifying downlink and provides multiplex capability for two split-screen outputs.

    All television cameras are identical with the exception of accessory hardware. Any camera can be equipped with three types of interchangeable lens assemblies, depending on mission requirements. Both monochrome and color lenses are available with normal field of view. Color lenses are also available with a wide-angle field of view. In addition to focus control for sharp images, several special lens control functions are available. All functions can be controlled by both onboard and ground uplink commands.

    An automatic light control is available for each camera. Its modes compensate for scene brightness by eliminating a fraction of the brightness. It then adjusts the light level of the scene by controlling (1) the silicone intensified target tube high voltage, (2) lens aperture (iris) and (3) automatic gain control. As a result, more detail and contrast are available in the brighter areas of a scene. Light or dark areas of a camera scene can be enhanced or subdued by means of gamma commands. More contrast can be obtained within light or dark areas. Zoom capability magnifies or reduces the size of objects in a camera's field of view by adjusting the focal length of the lens. The minimum focus for standard lenses is 3 feet, the maximum is infinity.

    Although five payload camera bay inputs can be used per mission, more than five camera locations are available. If a mission requires a keel camera, only one camera can be positioned at one of four locations along the keel. If a camera is connected to the keel/EVA position, then TV camera views are not available while the EVA camera is in use. Each remote manipulator system can accommodate two cameras, mounted at the wrist and elbow locations; however, scenes are only available from either the wrist or the elbow, selectable by means of panel A7 switches.

    With the exception of the RMS wrist camera and keel camera, all exterior cameras are mounted on motor-driven pan-tilt units. Camera azimuth (left and right) position and elevation (up and down) from a reference can be controlled on board by the flight crew or by ground uplink. The front tilt unit can accommodate up to plus or minus 170 degrees of pan and plus or minus 170 degrees of tilt from center. In tilt, the cameras cannot look straight down; in pan, the cameras cannot look straight down or straight back, which accounts for the range of 340 degrees instead of 360 degrees. The pan and tilt units move at two speeds: the high rate allows 12-degree-per-second rotation, and the low rate allows 1.2-degree-per-second rotation. Each pan and tilt unit contains a thermostatically controlled heater to maintain proper temperature.

    Up to three cameras can originate from payload sources. These inputs can come from pallet-mounted cameras or from the Spacelab module during a Spacelab mission. If the payload cameras are part of the orbiter flight hardware, the video control unit can command the camera in focus and zoom. However, Spacelab provides the electrical interface camera outlets that are available inside the orbiter crew compartment for portable cameras. The flight crew would deploy, set up and stow the portable TV system.

    All interior cameras are equipped with a color lens. Monitors 1 and 2 and the TV viewfinder monitor provide a monochrome output regardless of the lens assembly utilized, allowing the crew to adjust and view the video scenes. The flicker that appears in the interior camera scenes is caused by the rotating color wheel used to generate the field-sequential color signal that allows color outputs on the ground. The color-conversion process on the ground eliminates the flicker before the color signal is distributed.

    The basic monochrome camera converts light images into a composite video signal (picture plus sync) and can produce either a black-and-white or color image, depending on the lens assembly used. The lens provided with the TV system is an f/1.4 color zoom lens equipped with a six-segment, three-color rotating filter wheel that produces sequential red, green and blue color fields. Motorized lens control functions that vary the zoom, iris and focus are controlled manually by lens switches remotely from panel A7 or by ground command. The small portable monochrome viewfinder monitor (8 by 4.25 by 3.6 inches, weighing 4 pounds) enables the crew to view camera video output for picture quality and scene verification when monitors 1 and 2 are not accessible or available. The TV cable connects the camera and the two TV system input stations located in the crew cabin at panels O19 and M058F. The cable provides the camera with 28-volt dc power; camera and lens commands along with the sync signal; and video, including camera and lens data, to the control unit for distribution. The viewfinder monitor cable, which is 9 feet long, provides an interface between the camera and the monitor.

    Interior camera mounts restrain the camera assembly where it is used. Fixed-location quick-shoe mounts interface with any smooth, flat crew compartment surface. Clamp adapters can be used to affix the camera to panel switch guards or handholds. In addition, a special baseplate can be attached to either of two mounting locations on the sills of overhead windows 9 and 10.

    The two black-and-white TV console monitors in the aft flight deck crew station on panel A3 are identical and are arranged one over the other. The top one is monitor 1 and the bottom is monitor 2. The monitors are used to view video on board. Camera scenes from any exterior, interior or payload camera can be viewed on either monitor. The video being downlinked can also be observed on either monitor. In addition, the monitors can be used to view the video being recorded or played back on the VTR. They can also display cabin TV output when it is more convenient than using the viewfinder monitor.

    Each monitor accommodates a split-screen image from any two cameras. This feature is generated by the control unit and selected by panel A7 commands. Alphanumeric displays of camera location, pan and tilt angles, and temperatures of any cameras in an overtemperature condition can be superimposed on the monitor scene. A cross hair is also available at the center of the screen as an aid in remote manipulator system and proximity operations. The monitors are equipped with displays and controls for direct control of most monitor operations; however, the assignment of an input source to the monitor is controlled by panel A7. Each monitor is 12.5 by 10 by 7 inches and weighs 21 pounds.

    The video tape recorder is a cassette recorder with the capability to record both video data and voice annotation. Video from any camera source can be recorded on board and in color if the selected camera has color capability using off-the-shelf cassettes of 20 and 30 minutes. Voice annotation can be input directly to the VTR through a headset interface unit. The HIU interfaces only with the VTR, not with the vehicle communications system. A tone may also be recorded to identify a particular location for subsequent playback operations. Voice annotation can be added during playback and records over any previously recorded audio.

    All actual VTR operations must be performed by the crew-everything from tape changeout to VTR activation. The VTR is configured to receive its video input from monitor 2 (to record video on the VTR, the desired camera output must be displayed on monitor 2, which is connected directly to the recorder). For onboard playback, the recorded video can be reviewed on monitor 2 by positioning its source switch to direct and initiating a playback mode. For downlink purposes, the VTR is connected to the payload 1 input, which is a panel A7 video input selection. The switch panel located above the VTR includes a circuit breaker for the recorder. The recorder is equipped with a no video light that is illuminated when no video source is present at the recorder. When recording, the operator should verify that the no video light is off. An end of tape light indicates when the cassette is out of tape, and the VTR automatically stops. Tape changeout is simply a matter of ejecting the cassette from the recorder.

    The extravehicular mobility unit TV is a fully portable remote television unit. It transmits black-and-white television pictures to the orbiter CCTV system from virtually any location outside the orbiter. Orbiter reception is through either of the S-band FM hemispherical antennas. The TV assembly fits over the EMU helmet and light assembly. It is battery powered (28 volts dc) and transmits video at 1,775.7 MHz to a video receiver/processing unit installed in the orbiter's middeck. When the receiver is connected to the TV input station at panel M058F by a standard 20-foot TV power cable, crew members can view real-time EVA video on either monitor by selecting the middeck camera input on panel A7. EVA video can also be selected for return link or taping.

    Auxiliary lighting to improve scene quality is available for both exterior and interior cameras. In the payload bay, floodlights are mounted on the forward bulkhead and at various locations along the lower payload bay interior. A spotlight is also connected to the RMS wrist. These lights are controlled from panel A7.

    Video signals can be downlinked by means of onboard or uplink commands. If a scene assignment for downlink is commanded through S-band or Ku-band uplink, the appropriate camera command is received through the network signal processor by the orbiter general-purpose computer. The GPC then transmits the command to the control unit on payload MDM PF2. The control unit selects the desired camera video and combines it with intercom A or B (if desired) and then sends a composite signal to the FM signal processor.

    The FM signal processor transfers the TV signal to the FM transmitter, which modulates the signal and transmits it to a ground station with TV capability. Video data consisting of one of four signals (including real-time TV, main engine data, recorder dumps and payload data) can be individually assigned for downlink to the single S-band FM wide-band channel. Thus, if one of the other sources is being downlinked, video cannot be transmitted to the ground simultaneously.

    The ground station receives, processes and records the downlinked TV signal. Real-time TV can be transmitted directly from the ground station to the Mission Control Center through a relay satellite. Sites with TV capability are Merritt Island, Goldstone and Hawaii.

    The orbiter Ku-band system can downlink TV signals through the TDRSS directly to the Houston Mission Control Center. Since the TDRSS is not restricted to specific tracking stations, more continuous TV capability is available. However, only one Ku-band FM channel is used to downlink real-time TV, main engine data, recorder dumps and payload data. Therefore, video cannot be transmitted if one of the other sources is being downlinked.


    The support system for the orbiter experiments was developed to record data obtained and to provide time correlation for the recorded data. The information obtained through the sensors of the OEX instruments must be recorded during the orbiter mission because there is no real-time or delayed downlink of OEX data. In addition, the analog data produced by certain instruments must be digitized for recording.

    The support system for OEX comprises three subsystems: the OEX recorder, the system control module and the pulse code modulation system. The SCM is the primary interface between the OEX recorder and the experiment instruments and between the recorder and the orbiter systems. It transmits operating commands to the experiments. After such commands are transmitted, it controls the operation of the recorder to correspond to the experiment operation. The SCM is a microprocessor-based, solid-state control unit that provides a flexible means of commanding the OEX tape recorder and the OEX and modular auxiliary data system.

    The PCM system accepts both digital and analog data from the experiments. It digitizes the analog data and molds it and the digital data received directly from the experiments into a single digital data stream that is recorded on the OEX recorder. The PCM also receives time information from the orbiter timing buffer and injects it into the digital data stream to provide the required time correlation for the OEX data.

    The SCM selects any of 32 inputs and routes them to any of 28 recorder tracks or four-line driver outputs to the T-0 umbilical; executes real-time commands; controls experiments and data system components; and provides manual, semiautomatic and automatic control.

    The recorder carries 9,400 feet of magnetic tape that permits up to two hours of recording time at a tape speed of 15 inches per second. After the return of the orbiter, the data tape is played back for recording on a ground system. The tape is not usually removed from the recorder.


    The SILTS experiment will obtain high-resolution infrared imagery of the upper (leeward) surface of the orbiter fuselage and left wing during atmospheric entry. This information will increase understanding of leeside aeroheating phenomena and will be used to design a less conservative thermal protection system. SILTS provides the opportunity to obtain data under flight conditions for comparison with data obtained in ground-based facilities.

    Six primary components make up the SILTS experiment system: (1) an infrared camera, (2) infrared-transparent windows, (3) a temper ature-reference surface, (4) a data and control electronics module, (5) a pressurized nitrogen module and (6) window protection plugs. These components are installed in a pod that is mounted atop the vertical stabilizer and capped at the leading edge by a hemispherical dome. (The SILTS pod replaces the top 24 inches of the vertical stabilizer.) Within this dome, the infrared camera system is mounted in such a way that it rotates to view the orbiter leeside surfaces through either of two windows-one offering a view of the orbiter fuselage and the other a view of the left wing. The camera is sensitive to heat sources from 200 to 1,000 F.

    The camera's indium-antimonide detector is cooled to cryogenic temperatures by a Joule-Thompson cryostat. The camera's field of view is 40 by 40 degrees. Its rotating prism system scans four 100-line fields each second, with a 4-1 interlace, resulting in a 400-line image.

    Each of the two infrared-transparent window assemblies consists of dual silicone windows constrained within a carbon-phenolic window mount. The windows and window mount assemblies are designed to withstand the entry thermal environment to which they would be subjected without active cooling. They are, however, transpiration cooled with gaseous nitrogen during experiment operation so that they do not reach temperatures at which they would become significant radiators in the infrared. A small thermostatically controlled surface between the two window assemblies provides an in-flight temperature reference source for the infrared camera.

    The pressurized nitrogen system comprises two 3,000-psi gaseous nitrogen bottles and all associated valves and plumbing. The pressure system supplies gaseous nitrogen to the cryostat for camera detector cooling, to the external window cavities for window transpiration cooling, and to pin pullers that initiate the ejection of the advanced flexible reusable surface insulation window protection plugs upon SILTS activation to expose the viewing ports and camera.

    The information obtained by the camera is recorded on the OEX tape recorder. The data, when reduced and analyzed, will produce a thermal map of the viewed areas.

    The SILTS experiment is initiated by the onboard computers approximately five minutes before entry interface, which occurs at an altitude of approximately 400,000 feet. The camera operates for approximately 18 minutes through the forward-facing window and left-facing window, alternating evenly between the two about every five seconds.

    After the six planned SILTS missions, an analysis of structural loads will determine whether the SILTS pod should be removed and replaced with the original structure or remain in position for other uses. The pod thermal protection system is high-temperature reusable surface insulation black tiles, whose density is 22 pounds per cubic foot.


    Accurate aerodynamic research requires precise knowledge of vehicle attitude and state. This information, commonly referred to as air data, includes vehicle angle of attack, angle of sideslip, free-stream dynamic pressure, Mach number and total pressure. An evaluation of the orbiter baseline air data system indicated that flight air data would not be available above approximately Mach 3.5 and that the accuracy of the air data would not satisfy aerodynamic research requirements. Therefore, SEADS was developed under the orbiter experiments program to take the measurements required for precise determination of air data across the orbiter's atmospheric flight-speed range (i.e., hypersonic, supersonic, transonic and subconic Mach numbers) or from lift-off to 280,000 feet during ascent and from 280,000 feet to touchdown during entry.

    The key to incorporating SEADS in the shuttle orbiter was the development of a technique for penetrating the orbiter's reinforced carbon-carbon nose cap to obtain the required pressure measurements. The SEADS nose cap penetration assembly evolved as a result of extensive design, fabrication and test programs that evaluated high-temperature (greater than 2,600 F) materials and configuration concepts. The coated columbium penetration assembly selected then was fabricated for installation in a specially modified baseline geometry nose cap. The SEADS nose cap contains an array of 14 penetration assemblies, associated coated columbium pressure tubing, support structure, pressure transducers and system-monitoring instrumentation. Data from the SEADS pressure transducers are transmitted to the OEX support system and stored on the OEX tape recorder for postflight data analysis.


    The SUMS experiment will obtain measurements of free-stream density during atmospheric entry in the hypersonic, rarefied flow regime. These measurements, combined with acceleration measurements from the companion high-resolution accelerometer package experiment, will allow calculation of orbiter aerodynamic coefficients in the flow regime previously inaccessible to experimental and analytic techniques. SUMS complements SEADS by providing data at higher altitudes. The resultant flight data base will aid in future development of analysis techniques and laboratory facilities for predicting winged-entry-vehicle performance in hypersonic rarefied flow. Furthermore, SUMS will measure equilibrium gas composition at the inlet port, making the experiment a pathfinder for future mass spectrometer application in the study of aerothermodynamic properties of the transition flow field.

    The SUMS experiment system consists of a sample orifice, an inlet system and a mass spectrometer. The sample orifice penetrates a thermal tile just aft of the fuselage stagnation point and just forward of the orbiter nose wheel well. The orifice is connected to the inlet system by a short tube through the forward nose wheel well bulkhead. The inlet system is connected through a longer tube to the mass spectrometer, which is mounted above the inlet system on the forward nose wheel well bulkhead. SUMS is designed for easy removal and reinstallation between flights to accommodate modification or repair.

    The mass spectrometer is a flight spare unit from the Viking project's upper atmosphere mass spectrometer system. The unit has been modified to be compatible with the orbiter's mechanical, electrical and data systems. The mass spectrometer measures gases from hydrogen through carbon dioxide at a five-second rate. The inlet system contains two switchable flow restrictors that expand the measurement range of the mass spectrometer and position its measurement interval over the desired altitude range. Data from SUMS are output to the OEX data system for recording during flight operation.

    SUMS is controlled by stored commands that are transmitted to the orbiter during flight and by internal software logic. Application of power for vacuum maintenance or for normal operation is controlled by stored commands; while internal control of system operation, such as opening and closing valves, is performed by preprogrammed logic. SUMS will be powered on shortly before deorbit burn initiation and will sample the inlet gases down to an altitude of 40 nautical miles. At an altitude of about 59 nautical miles, the range valve will close to switch between the two flow restrictors. At 59 nautical miles, the inlet valve and protection valve will close; but the mass spectrometer will continue to operate until landing, observing the pump-down and background signals after entry.

    Operation of SUMS on repeated shuttle flights will not only build a substantial body of aerothermodynamic data for future winged-entry-vehicle design applications, but also add to the knowledge of mass spectrometer applications in aerothermodynamic research. As a further benefit, data will be obtained on atmospheric properties in the altitude range where experimental data are, to date, extremely sparse.


    Although all of the generic data types required for aerodynamic parameter identification are available from the baseline orbiter systems, the data are not suitable for experimentation because of such factors as sample rate deficiencies, inadequate data resolution or computer cycle time and core size interactions. In addition, the baseline data are operational measurements that are not subject to the desired changes for conducting experiments. The ACIP is a group of sensors that will be placed on the orbiter to obtain experiment measurements unavailable through the baseline system.

    The primary ACIP objectives are as follows: (1) to collect aerodynamic data in the hypersonic, supersonic and transonic flight regimes, regions in which there has been little opportunity for gathering and accumulating practical data; (2) to establish an extensive aerodynamic data base for verifying and correlating ground-based test data, including assessments of the uncertainties in such data; and (3) to provide flight dynamics state-variable data in support of other technology areas, such as aerothermal and structural dynamics.

    Implementing the ACIP program will benefit the space shuttle because the more precise data obtained through the ACIP will enable earlier attainment of the spacecraft's full operational capability. Currently installed instrumentation provides sufficiently precise data for orbiter operations, but not for research. The result is that constraint removal would either be based on less substantive data or would require a long-term program of gathering the less accurate data.

    The ACIP incorporates three triads of instruments: one of linear accelerometers, one of angular accelerometers and one of rate gyros. Also included are the power conditioner for the gyros, the power control system and the housekeeping components for the instruments. The ACIP is aligned to the orbiter's axes with extreme accuracy. Its instruments continually sense the dynamic X, Y and Z attitudes and the performance characteristics of the orbiter during the launch, orbital, entry and descent phases of flight. In addition, the ACIP receives the indications of orbiter control surface positions and converts the information into higher orders of precision before recording it with the attitude data. The output signals are routed to the pulse code modulation system for formatting with orbiter time data and data from the orbiter experiments. The data are then stored in the OEX tape recorder.


    This experiment uses an orthogonal, triaxial set of sensitive linear accelerometers to take accurate measurements of low-level (down to micro-g's) aerodynamic accelerations along the orbiter's principal axes during initial re-entry into the atmosphere, i.e., in the rarefied flow regime.

    The aerodynamic acceleration data from the HiRAP experiment, output on existing ACIP channels, have been used to calculate rarefied aerodynamic performance parameters and/or atmospheric properties pertaining to several flights, beginning with the STS-6 mission. These flight data support advances in predicting the aerodynamic behavior of winged entry vehicles in the high-speed, low-density flight regime, including free molecular flow and the transition into the hypersonic continuum. Aerodynamic performance under these conditions cannot be simulated in ground facilities; consequently, current predictions rely solely on computational techniques and extrapolations of tunnel data. For improvement or advances, these techniques depend on actual flight data to serve as benchmarks, particularly in the transition regime between free molecular flow and continuum flow.

    Advancements in rarefied aerodynamics of winged entry vehicles may also prove useful in the design of future advanced orbital transfer vehicles. Such OTVs may use aerodynamic braking and maneuvering to dissipate excess orbital energy into the upper atmosphere upon return to lower orbits for rendezvous with an orbiter from the space station. A key aerodynamic parameter in the OTV design is the lift-to-drag ratio, which is measured directly in the HiRAP experiment. Furthermore, an OTV may require a flight-proven, sensitive onboard accelerometer system to overcome uncertainties in the upper atmosphere. The experience gained from the planned multiple HiRAP flights may provide valuable test data for the development of future navigation systems. In addition, the experiment provides data on key atmospheric properties (e.g., density) in a region of flight that is not readily accessible to orbital vehicles or regular meteorological soundings.


    This onboard instrumentation system measures and records selected pressure, temperature, strain, vibration and event data to support payloads and experiments and to determine orbiter environments during flight. It supplements existing orbiter operational instrumentation by conditioning, digitizing and storing data from selected sensors and experiments.

    The MADS collects detailed data during ascent, orbit and entry to define vehicle response to flight environments. It permits correlation of data from one flight to another and enables comparison of flight data from one orbiter to another orbiter.

    All MADS equipment installed in the orbiter is structurally mounted and environmentally compatible with the orbiter and mission requirements. Because of its location, the MADS does not intrude into the payload envelope. Equipment consists of a pulse code modulation multiplexer, a frequency division multiplexer, a power distribution assembly and appropriate signal conditioners mounted on shelf 8 beneath the payload bay liner of the midfuselage.

    In OV-102 (Columbia), MADS inputs its information to the system control module and records it on the OEX recorder located below the crew compartment middeck floor. In OV-103 (Discovery) and OV-104 (Atlantis), a MADS control module and recorder are mounted below the crew compartment middeck floor.

    MADS records approximately 246 measurements from the orbiter airframe, skin and orbital maneuvering system/reaction control system left-hand pod.

    The MADS interfaces with the orbiter through the orbiter's electrical distribution system and operational instrumentation inputs for status monitoring. Coaxial cables and wire harnesses from the sensors are routed through the orbiter payload bay harness bundles to the signal conditioners, PCM multiplexer and FDM, attached to the midfuselage shelf. After the signal conditioners and the multiplexers have processed the data, four outputs of the FDM and one output of the PCM multiplexer are routed forward to the SCM in OV-102 for recording on the OEX recorders. In OV-103 and OV-104, the four outputs of the FDM and one output of the PCM multiplexer are routed forward to the MCM for recording on five tracks of the MADS recorder. In addition, the MADS recorder is used during ascent to record additional space shuttle data consisting of solid rocket booster wide band and external tank signals.

    The MADS is not considered mandatory for launch, and its loss during flight does not cause a mission abort. It measures and records data for predetermined events established by test and mission requirements.

    For a typical mission, approximately five hours before launch, the MADS is powered on from the preset switch configuration to supply a prelaunch manual calibration. (Power is supplied from the orbiter's main buses A and B.) After calibration, all switches are returned to the preset configuration, leaving the MADS in the standby position and only the MCM receiving power. This mode continues until nine minutes before launch, at which time the MADS attains the full-system mode through uplink commands and all its components are powered on. In this mode, the MADS recorder is operating at a continuous tape speed of 15 inches per second, recording aerodynamic coefficient identification package, flight acceleration safety cutoff, ET, SRB, wide-band and PCM data. The MADS PCM bit rate is 64 kbps.

    The wide-band-only mode is used during the prelaunch automatic and manual calibrations. This mode records the ac and dc calibration levels provided by the FDM. Each manual calibration level is recorded for 10 seconds at a tape speed of 15 inches per second in the continuous mode.

    Approximately 12 minutes after launch, the MADS is commanded into the PCM-snapshot-with-strain-gauge-signal-conditioner mode. In this mode, the recorder is in the sample mode, conserving power and recorder tape by recording data for 10 seconds every 10 minutes at a PCM bit rate of 32 kbps and a tape speed of 3.75 inches per second. Two minutes before the second orbital maneuvering system thrusting period, the MADS is commanded into the full-system mode until the thrusting period is completed. Then it is commanded into the PCM-only mode, which continues during the orbit until a quiescent period is reached. In OV-102 only, one minute of ACIP calibration is required during this period, after which the MADS continues in the PCM-only mode. The system is switched to the full-system mode for the OMS separation thrusting periods and can be returned to the PCM-only mode for the majority of the on-orbit mission.

    The PCM-with-SGSC mode is similar to the PCM-only mode, but strain measurements are also recorded during this period. The SGSC operation is cycled along with the other MADS equipment and signal conditioners by uplink commands to maintain the required operational temperatures. This mode occurs between two full-system modes to minimize flight crew participation and conserve power and recorder tape. It can be initiated from the full-system mode or returned to the full-system mode by one uplink command. To shift this mode to the PCM-only mode, the SGSC must be commanded off manually by the flight crew. This mode is used on orbit.

    Two minutes before the deorbit thrusting period, the MADS is put into the full-system mode for one hour to record descent (entry) data. At the conclusion of the one-hour period, it is placed in the PCM-only mode for approximately four hours to measure postlanding thermal data and is then powered down for the entire postlanding period.

    With the use of the MADS switches located in the crew compartment, control can be initiated by the flight crew. To reduce the flight crew's participation, all commands are uplinked from the Mission Control Center in Houston and transmitted to the onboard payload forward 1 multiplexer/demultiplexer. The MDM then routes the commands to the SCM for processing in OV-102 and to the MCM in OV-103 and OV-104. Power for the MADS will be supplied by the orbiter's 28-volt dc main buses A and B.

    The flight acceleration safety cutoff system interfaces 12 orbiter main engine vibration measurements with the MADS. The variety of MADS measurements is collected by thermocouples, resistance thermometers, radiometers, vibration sensors, strain gauges or pressure transducers.

    The MADS shelf 8 components are protected from overheating by shelf temperature monitoring and control of MADS operation by ground commands. The MADS is thermally isolated from the orbiter structure by 0.049-inch thin-wall titanium struts. It is also protected from the orbiter environment by a 1.5-inch bulk-insulation enclosure.

    The MADS recorders in OV-103 and OV-104 are Data Tape/Kodak 28-track, wide-band, modular, airborne recording systems similar to the OV-102 orbiter experiments recorder. The recorders are capable of simultaneously recording, and subsequently reproducing, 28 tracks of digital biphase L data or any combination of wide-band analog and digital biphase L data up to 28 tracks.

    After OV-103 and OV-104 return from a mission, the recorder tape is played back to record the data on a ground recording system. The tape is not removed from the flight recorder. The total MADS weight is 641 pounds.

    The following are the contractors involved with the orbiter instrumentation and communication systems: Aydin Vector Division, Newton, Pa. (wide-band frequency division multiplexers); Communications Components, Costa Mesa, Calif. (UHF antenna); Conrac Corp., West Caldwell, N.J. (mission timer, event timer, ground command interface logic box, FM signal processor); Eldec Corp., Lynwood, Wash. (dedicated signal conditioner); Endevco, San Juan Capistrano, Calif. (piezoelectric accelerometer, piezoelectric pickup {acoustic and vibration}); Gulton Industries, Costa Mesa, Calif. (accelerometer linear flow frequency, vibration and acoustic); Harris Corp., Electronic Systems Division, Melbourne, Fla. (pulse code modulation master unit, orbiter/payload timing buffer and payload data interleaver); Hughes, El Segundo, Calif. (Ku-band radar, communication system deployable antenna and electrical assembly); K-West, Westminster, Calif. (wide-band signal conditioner, strain gauge signal conditioner); Magnavox, Ft. Wayne, Ind. (UHF receiver/transmitter mount, UHF receiver/transmitter); Micro Measurements, Romulus, Mich. (strain gauge); RDF Corp., Hudson, N.H. (sensors, transducers); Rosemount Inc., Eden Prairie, Minn. (transducers, sensors); Radio Corp. of America, Astro-Electronics Division, Princeton, N.J. (closed-circuit television); Spectran, La Habra, Calif. (sensors); Sperry Rand Corp., Flight Systems Division, Phoenix, Ariz. (multiplexer/demultiplexer); Stratham Instruments, Oxnard, Calif. (transducers); Systron-Donner, Concord, Calif. (accelerometer); Teledynamics Division of Ambac Industries, Fort Washington, Pa. (S-band transmitter, FM transmitter, S-band transceiver); Telephonics Division, Instruments System Corp., Huntington N.Y. (orbiter audio distribution system); TRW Systems, Electronic Systems Division, Redondo Beach, Calif. (S-band payload interrogator, S-band network equipment, network signal processor, payload signal processor); Transco Products, Venice, Calif. (S-band switch); Watkins Johnson, Palo Alto, Calif. (C-band radar altimeter antenna, L-band TACAN, UHF antenna, S-band power amplifier); Wavecom, Northridge, Calif. (S-band multiplexer); Westinghouse Electric Corp., Systems Development Division, Baltimore, Md. (master timing unit); AIL, Huntington, N.Y. (S-band preamplifier assembly); AVCO, Wilmington, Mass. (Ku-band MSBLS antenna, Ku-band waveguide); Teledyne-Microwave, Mountain View, Calif. (S-band switch assembly); Teledyne-Electronics, Newberry Park, Calif. (S-band FM transmitter); RCA, Government Communications Systems, Camden, N.J. (extravehicular activity communication system); Data Tape Inc., division of Kodak Corp., Pasadena, Calif. (MADS recorder); Gulton Data Systems, Gulton Industries Inc., Albuquerque, N.M. (pulse code modulation multiplexer); Odetics Space Born Corp., Anaheim, Calif. (operational and payload recorders); Rockwell, Anaheim, Calif. (S-band quad antenna, S-band hemi antenna, S-band payload antenna).

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Table of Contents

Information content from the NSTS Shuttle Reference Manual (1988)
Last Hypertexed Friday April 17 15:30:30 EDT 1998
Jim Dumoulin (

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