There are three rotational hand controllers on the orbiter crew compartment flight deck: one at the commander's station, one at the pilot's station and one at the aft flight deck station. Each RHC controls vehicle rotation about three axes: roll, pitch and yaw. During ascent, the commander's and pilot's RHCs may be used to gimbal the SSMEs and SRBs. For insertion and deorbit, the commander's and pilot's RHCs may be used to gimbal the orbital maneuvering system engines and to command thrusting of the reaction control system engines. On orbit, the commander's, pilot's and aft flight station RHCs may be used to command RCS engine thrusting. During entry, the commander's and pilot's RHCs may be used to command RCS engine thrusting during the early portion of entry and may be used to position the orbiter elevons in roll and pitch axes in the latter portion of entry.
Human factors dictate that an RHC deflection produce a rotation in the same direction as the flight crew member's line of sight. The aft flight station RHC is used only on orbit. An aft sense -Z switch on panel A6 selects the line-of-sight reference about the minus Z axis (overhead windows), and the -X position selects the line-of-sight reference about the minus X axis (aft windows) in order for aft RHC commands to be correctly transformed to give the desired orbiter movement.
Several switches are located on the RHC. A backup flight system (BFS) mode button on the commander's and pilot's RHCs engage the BFS when depressed. The commander's, pilot's, and aft flight station RHCs have a two-contact trim switch that can be pushed forward or aft to add a trim rate to the RHC pitch command; pushing it left or right adds a roll trim rate. The aft RHC's trim switch is inactive. The communications switch on each RHC is a push-to-talk switch that enables voice transmission when the switch is depressed.
Each RHC contains nine transducers: three redundant transducers sense pitch deflection, three sense roll deflection and three sense yaw deflection. The transducers produce an electrical signal proportional to the deflection of the RHC. The three transducers are called channels 1, 2 and 3; the channel selected by redundancy management provides the command. Each channel is powered by a separate power supply in its associated display driver unit. Each controller is triply redundant; thus, it takes only one good signal from a controller for the controller to operate.
Each RHC has an initial dead band of 0.25 of a degree in all three axes. To move the RHC beyond the dead band, an additional force is required. When the amount of deflection reaches a certain level, called the softstop, a step increase in the force required for further deflection occurs. When a software detent position is exceeded, that RHC assumes control.
The softstop occurs at 19.5 degrees in the roll and pitch axes and at 9.5 to 10.5 degrees in the yaw axis. To reach the softstop in the roll axes, 40.95 inch-pounds of static torque deflection are required; 38.2 inch-pounds are needed in pitch and 7 inch-pounds in yaw.
The mechanical hardstop that can be obtained in an axis is 24.3 degrees in the roll and pitch axis and 14.3 degrees in the yaw axis.
Software normally flows from the RHCs to the flight control system through redundancy management and a SOP before it is passed to the aerojet digital autopilot.
In a nominal mission, the flight crew has manual control of the RHC during every major mode except terminal countdown. When an RHC deflection exceeds the detent in an axis, the RHC SOP generates a discrete signal that converts the RHC from the automatic mode to control stick steering, or hot stick. However, during ascent when the ascent digital autopilot is active, a CSS pitch and/or roll/yaw mode push button light indicator on panel F2 or F4 must be depressed in order for manual inputs to be implemented into the flight control system from the commander's or pilot's RHC. When a CSS pitch or roll/yaw push button light indicator is depressed on panels F2 or F4, the white light for the push button indicator will be illuminated and that axis will be downmoded from automatic to CSS.
When the flight crew commands three-axis motion using the RHC, the GPCs process the RHC and motion sensor commands; and the flight control system interprets the RHC motions (fore and aft, right and left, clockwise and counterclockwise) as rate commands in pitch, roll and yaw and then processes the flight control law (equations) to enhance control response and stability. If conflicting commands are given, no commands result.
During orbital flight, any one of the three stations can input three-axis control commands to the flight control system. During entry and landing, the commander and pilot have two-axis (roll and pitch only) capability. Roll, pitch and yaw aerosurface deflection trim is controlled by the panel trim switches, while roll and pitch vehicle rate trim is controlled with the trim switches on the RHC. For a return-to-launch-site abort, both the commander's and pilot's RHC have three-axis capability during major mode 601 and roll and pitch during major modes 602 and 603.
The commander's RHC is powered when the flt cntlr (controller) on/off switch on panel F7 is positioned to on . The pilot's RHC is powered when the flt cntlr on/off switch on panel F8 is positioned to on. The aft RHC is powered when the flt cntlr on/off switch on panel A6 is positioned to on .
If a malfunction occurs in the commander's or pilot's RHC, the red RHC caution and warning light on panel F7 is illuminated.
The RHC contractor is Honeywell Inc., Clearwater, Fla.
There are two translational hand controllers: one at the commander's station and one at the aft flight deck station. The commander's THC is active during orbit insertion, on orbit and during deorbit. The aft flight deck station THC is active only on orbit. The THCs are used for manual control of translation along the longitudinal (X), lateral (Y) and vertical (Z) vehicle axes using the RCS.
Each THC contains six three-contact switches, one in the plus and minus directions for each axis. Moving the THC to the right commands translation along the plus Y axis and closes three switch contacts (referred to as channels 1, 2 and 3). Redundancy management then selects the channel and provides the command.
An aft sense switch on panel A6 selects the line-of-sight reference along the minus X or minus Z axis of the orbiter for the aft THC. The aft sense switch must be in the -X position for aft windows and -Z for the overhead windows in order for the aft THC commands to be correctly transformed to give the desired orbiter movement.
The normal displacement of a THC is 0.5 of an inch from the center null position in both directions along each of the three THC axes. A force of 2 pounds is required to deflect either THC 0.5 of an inch in all axes.
The redundant signals from the forward and aft THC pass through a redundant management process and a SOP before being passed to the flight control system. If both the forward and aft THCs generate conflicting translation commands, the output translation command is given.
In what is referred to as the transition digital autopilot mode, the commander's THC is active and totally independent of the flight control orbital digital autopilot ( DAP ) push buttons on panel C3 or A6 or the position or status of the RHC. Whenever the commander's THC is out of detent plus or minus X, Y or Z, translation acceleration commands are sent directly to the RCS jet selection logic for continuous RCS thrusting periods. Rotational commands may be sent simultaneously with translation commands within the limits of the RCS jet selection logic; if both plus X and minus Z translations are commanded simultaneously, plus X translation is given priority.
The commander's THC is powered when the flt cntlr on/off switch on panel F7 is positioned to on . The aft THC is powered when the flt cntlr on/off switch on panel A6 is positioned to on .
The THC contractor is Honeywell Inc., Clearwater, Fla.
The pitch auto, CSS and roll/yaw auto , CSS push button light indicators are located on panel F2 for the commander and on panel F4 for the pilot. Each push button light indicator is triply redundant. During entry, depressing a CSS push button light indicator will mode flight control to augmented manual in the corresponding axis, illuminate both CSS lights and extinguish both auto lights for that axis. Depressing a CSS push button light indicator to auto will return flight control in that axis to auto, extinguish both CSS lights and illuminate both auto lights. During ascent, depressing any of the four CSS push button light indicators will mode flight control to augmented manual in all axes, illuminate all four CSS push button light indicators and extinguish all four auto lights. Depressing any of the four push button light indicators to auto will mode flight control to automatic, illuminate all four auto push button light indicators and extinguish the CSS push button light indicators.
There are two pairs of rudder pedals: one each for the commander and pilot. The commander's and pilot's rudder pedals are mechanically linked so that movement on one side moves the other side. When a pedal is depressed, it moves a mechanical input arm in a rudder pedal transducer assembly. Each RPTA contains three transducers-channels 1, 2 and 3-and generates an electrical signal proportional to the rudder pedal deflection. An artificial feel is provided in the rudder pedal assemblies.
The rudder pedals command orbiter rotation about the yaw axis by positioning the rudder during atmospheric flight. In atmospheric flight, flight control software performs automatic turn coordination; thus the rudder pedals are not used until the wings are level before touchdown.
The RPTA SOP converts the selected left and right commands from volts to degrees; selects the largest of the left and right commands for output to flight control software after applying a dead band; and if redundancy management declares an RPTA bad, sets that RPTA to zero.
The rudder pedals can be adjusted 3.25 inches forward or aft from the neutral position in 0.81-inch increments (nine positions). The breakout force is 10 pounds. A pedal force of 70 pounds is required to depress a pedal to its maximum forward or aft position.
The rudder pedals provide two additional functions unrelated to software after touchdown. Rudder pedal deflections provide nose wheel steering, and depressing the upper portion of the pedals by applying toe pressure provides braking. Differential braking may be used for nose wheel steering.
The commander's RPTA is powered when the flt cntlr on/off switch on panel F7 is positioned to on . The pilot's RPTA is powered when the flt cntlr on/off switch on panel F8 is positioned to on .
The RPTA contractor is Honeywell Inc., Clearwater, Fla.
There are two speed brake/thrust controllers: one on the left side of the flight deck forward on panel L2 for the commander and one on the right side of the center console on panel C3 for the pilot. The SBTCs serve two distinct functions: during ascent, the pilot's speed brake/thrust controller may be used to vary the thrust level of the three SSMEs. During entry, the commander's or pilot's speed brake/thrust controller may be used to control aerodynamic drag (hence airspeed) by opening or closing the speed brake.
Depressing a takeover switch (with three contacts) on each SBTC switches to manual control of the SSME thrust level setting (pilot's only) or speed brake position. Each SBTC contains three transducers-channels 1, 2 and 3, which produce a voltage proportional to the deflection. Redundancy management selects the output.
In the case of the SSME thrust-level setting, the top half of both spd bk/throt push button light indicators on panels F2 and F4 will be illuminated, indicating auto . Only the pilot's SBTC can be enabled for manual throttle control. The pilot depresses the takeover push button on the SBTC, causing the general-purpose computer throttle command to be frozen at its current value. While depressing the takeover button, the pilot moves the SBTC to match the frozen GPC command. Manual control is established when the match is within 4 percent. When the match is achieved, the spd bk/throt man push button light indicator on panel F4 will be illuminated and the auto light extinguished. The takeover push button is then released. If the takeover push button is released before a match is achieved, the system reverts to GPC auto commands. Under manual throttle command, depressing either or both push button light indicators on panel F2 or F4 will cause the system to revert to the GPC auto commands, extinguishing the pilot's man light and illuminating the auto lights on panels F2 and F4. Transferring back to auto during a return-to-launch-site abort leaves the throttle at the last-commanded manual setting until 3 g's, and the vehicle is held at 3 g's.
If the speed brake mode is in automatic and the commander or pilot wishes to control the speed brake manually, momentarily depressing the takeover push button grants control of the SBTC to the crew member who depressed the switch. The speed brake is driven to the position currently commanded by the SBTC. The spd bk/throt man push button light indicator on panel F2 will be illuminated if the commander takes control, extinguishing the auto light. If the pilot takes control, the spd bk/throt man push button light indicator on panel F4 will be illuminated, and the commander's light will be extinguished. To place the speed brake under software control, either or both spd bk/throt push button indicators on panel F2 or F4 can be depressed, and the auto lights on panels F2 and F4 will be illuminated.
The SBTC SOP converts the selected SSME throttle command to a setting in percent and the selected speed brake command from volts to degrees. In addition, the SBTC SOP selects the speed brake command from the SBTC whose takeover button was depressed last. If both takeover buttons are depressed simultaneously, the commander's SBTC is given control. If redundancy management declares an SBTC bad, the command is frozen.
The commander's SBTC is powered by the flt cntlr on/off switch on panel F7 when positioned to on . The pilot's SBTC is powered when the flt cntlr on/off switch on panel F8 is positioned to on .
The SBTC contractor is Honeywell Inc., Clearwater, Fla.
There are two body flap switches: one for the commander on panel L2 and one for the pilot on panel C3. Each switch is a lever-locked switch spring loaded to the center position. The body flap switches provide manual control for positioning the body flap for SSME thermal protection and for reducing elevon deflections during the entry phase.
The body flap can be switched from its automatic mode to its manual mode by moving either switch from the auto/off position to the up or down position. These are momentary switch positions; when released, the switch returns to auto/off . The white body flap man (lower half) of the push button light indicator on panel F2 or F4 is illuminated, indicating manual control of the body flap. To regain automatic control, the body flap push button light indicator on panel F2 or F4 is depressed, extinguishing the man white light and illuminating the auto white light. The push button indicator can also be depressed to man for manual body flap control. The push button light indicator is triply redundant.
The up and down positions of each switch have two power supplies from a control bus.
If the commander and pilot generate conflicting commands, a body-flap-up command will be output to flight control because up has priority.
The dual-redundant trim RHC/panel enable/inhibit switches on panel F3 provide signals to the GPCs, prohibiting software execution of the associated RHC and panel trim switch inputs while in the inhibit position. The enable position is not wired to the GPCs, permitting the RHC and panel trim switch inputs, which allows trimming.
When the trim RHC/panel switch on the left side of panel F3 is in enable, the commander's RHC trim switches command vehicle pitch and roll rates in major modes 304 and 305 (entry) and major modes 602 and 603 (return to launch site). The three trim switches on panel L2 for the commander are used to move the aerosurfaces in roll, pitch and yaw.
When the trim RHC/panel switch on the right side of panel F3 is in enable, the pilot's RHC trim switches command vehicle pitch and roll rates in major modes 304, 305, 602 and 603. The three trim switches on panel C3 for the pilot are used to move the aerosurfaces in roll, pitch and yaw.
Redundancy management processes the two sets of switches. If two switches generate opposing commands, the resultant trim command in that axis is zero.
The commander's trim roll l (left), r (right); the yaw trim l, r; and trim pitch up, down switches are located on panel L2. The pilot's switches are located on panel C3. The commander's trim switches on panel L2 are enabled when the trim panel on/off switch on the left side of panel F3 is positioned to on . The pilot's trim switches on panel C3 are enabled when the trim panel on/off switch on the right side of panel F3 is positioned to on. The corresponding trim RHC/panel enable/inhibit switch must be in enable in order for trimming to take place.
Each of the three trim switches on panels L2 and C3 are spring loaded to the center off position.
Redundancy management processes the two sets of switches. If two switches generate opposing commands, the resultant trim command in that axis is zero.
Aerosurface servoamplifiers are electronic devices that receive aerosurface commands during atmospheric flight from the flight control system software and electrically position hydraulic valves in aerosurface actuators, causing aerosurface deflections.
Each aerosurface is driven by a hydraulic actuator controlled by a redundant set of electrically driven valves (ports). There are four of these valves for each aerosurface actuator, except the body flap, which has only three. These valves are controlled by the selected ASAs.
In addition to the command channels from the ASAs to the control valves, there are data feedback channels to the ASAs from the aerosurface actuators. Each aerosurface has four associated position feedback transducers that are summed with the position command to provide a servoloop closure for one of the four independent servoloops associated with the elevons, rudder and speed brake. The body flap utilizes only three servoloops. The path from an ASA to its associated servovalve in the actuators and from the aerosurface feedback transducers to an ASA is called a flight control channel; there are, thus, four flight control channels, except for the body flap.
Each of the four elevons located on the trailing edges has an associated servoactuator that positions it. Each servoactuator is supplied with hydraulic pressure from the three orbiter hydraulic systems. A switching valve is used to control the hydraulic system that becomes the source of hydraulic pressure for that servoactuator. The valve allows a primary source of pressure (P1) to be supplied to that servoactuator. If the primary hydraulic pressure drops to around 1,200 to 1,500 psig, the switching valve allows the first standby hydraulic pressure (P2) to supply that servoactuator. If the first standby hydraulic pressure drops to around 1,200 to 1,500 psig, the secondary standby hydraulic source pressure (P3) is then supplied to that servoactuator. The yellow hyd press caution and warning light will be illuminated on panel F7 if the hydraulic pressure of system 1, 2 or 3 is below 2,400 psi and will also illuminate the red backup caution and warning alarm light on panel F7.
Each elevon servoactuator receives command signals from each of the four ASAs. Each actuator is composed of four two-stage servovalves that drive a modulating piston. Each piston is summed on a common mechanical shaft, creating a force to position a power spool that controls the flow of hydraulic fluid to the actuator power ram, controlling the direction of ram movement, thus driving the elevon to the desired position. When the desired position is reached, the power spool positions the mechanical shaft to block the hydraulic pressure to the hydraulically operated ram, locking the ram at that position. If a problem develops within a servovalve or it is commanded to a position different than the positions of the other three within an actuator, secondary delta pressure should begin to rise to 2,200 psi. Once the secondary delta pressure is at or above 2,200 psia for more than 120 seconds, the corresponding ASA sends an isolation command to the servovalve, opening the isolation valve, bypassing the hydraulic pressure to the servovalve, and causing its commanded pressure to the power spool to drop to zero, effectively removing it from operation. The pressure differential is sensed by a primary linear differential pressure transducer across the modulating piston when the respective FCS channel switch on panel C3 is in auto . This automatic function prevents excessive transient motion to that aerosurface, which could result in loss of the orbiter due to slow manual redundancy.
The FCS channel yellow caution and warning light on panel F7 will be illuminated to inform the flight crew of a failed channel. A red FCS saturation caution and warning light on panel F7 will be illuminated if one of the four elevons is set at more than plus 15 degrees or less than minus 20 degrees.
There are four FCS channel switches on panel C3- FCS channels 1, 2, 3 and 4; each has an override, auto and off position. The switch for a channel controls the channel for the elevons, rudder/speed brake and body flap, except channel 4, which has no body flap commands. When an FCS channel switch is in auto and that channel was bypassed, it can be reset by positioning the applicable switch to override . When an FCS channel switch is positioned to off , that channel is bypassed.
In each elevon servoactuator ram, there are four linear ram position transducers and four linear ram secondary differential pressure transducers. The ram linear transducers provide position feedback to the corresponding servoloop in the ASA, which is then summed with the position command to close the servoloop. This feedback is then summed with the elevon ram linear secondary differential pressures to develop an electrohydraulic valve drive current that is proportional to the error signal in order to position the ram. The maximum elevon deflection rate is 20 degrees per second.
During ascent, the elevons are deflected to reduce wing loads caused by rapid acceleration through the lower atmosphere. In this scheme, the inboard and outboard elevons are deflected together. By the time the vehicle reaches approximately Mach 2.5, the elevons have reached a null position, where they remain. This is accomplished by the initialized-loaded program.
The rudder/speed brake, which consists of upper and lower panels, is located on the trailing edge of the orbiter's vertical stabilizer. One servoactuator positions the panels together to act as a rudder; another opens the panels at the rudder's flared end so it functions as a speed brake.
The rudder and speed brake servoactuator receives four command signals from the four ASAs. Each servoactuator is composed of four two-stage servovalves that function like those of the elevons. The exception is that the rudder's power spool controls the flow of hydraulic fluid to the rudder's three reversible hydraulic motors and the power spool for the speed brake controls the flow of hydraulic fluid in the speed brake's three hydraulic reversible motors. Each rudder and speed brake hydraulic motor receives hydraulic pressure from only one of the orbiter's hydraulic systems. Each hydraulic motor has a hydraulic brake. When the motor is supplied with hydraulic pressure, the motor's brake is released. When the hydraulic pressure is blocked to that hydraulic motor, the hydraulic brake is applied, holding that motor and the corresponding aerosurface at that position.
The three hydraulic motors provide output to the rudder differential gearbox, which is connected to a mixer gearbox that drives rotary shafts. These rotary shafts drive four rotary actuators, which position the rudder panels.
The three speed brake hydraulic motors provide power output to the speed brake differential gearbox, which is connected to the same mixer gearbox as that of the rudder. This gearbox drives rotary shafts, which drive the same four rotary actuators involved with the rudder. Within each of the four rotary actuators, planetary gears blend the rudder positioning with the opening of the rudder flared ends.
There are four rotary position transducers on the rudder differential gearbox output and one differential linear position transducer in each rudder servoactuator. The rotary position transducers provide position feedback to the corresponding servoloop in the ASA. The feedback is summed with the linear differential pressures that develop the electrohydraulic valve drive current in proportion to the error signal in order to position the rudder.
There are also four rotary position transducers on the speed brake differential gearbox output and one differential linear pressure transducer in each speed brake servoactuator.
The rotary position transducers provide position feedback to the corresponding servoloop in the ASA, which is summed with the position command to close the servoloop. These are then summed with the linear differential pressures that develop the electrohydraulic valve drive current in proportion to the error signal to position the speed brake.
If a problem occurs in one of the four rudder or speed brake servoactuator channels, the corresponding linear differential pressure transducer will cause the corresponding ASA to signal a solenoid isolation valve to remove the pressure from the failed channel and bypass it if that FCS channel switch is in auto . The FCS channel switches' override and off positions and the FCS channel caution and warning light function the same as for the elevons. The hyd press light indicates a hydraulic failure. The rudder deflection rate is a maximum of 14 degrees per second. The speed brake deflection rate is approximately 10 degrees per second. If two of the three hydraulic motors fail in the rudder or speed brake, about half the design speed output will result from the corresponding gearbox due to its velocity summary nature.
Three servoactuators at the lower aft end of the fuselage are used to position the body flap; each is supplied with hydraulic pressure from an orbiter hydraulic system and has a solenoid-operated enable valve controlled by one of the three ASAs (the fourth ASA is not used for the body flap commands). Each solenoid-operated enable valve supplies hydraulic pressure from one orbiter hydraulic system to a corresponding solenoid-operated pilot valve, which is, in turn, controlled by one of the three ASAs. When the individual pilot valve receives a command signal from its corresponding ASA, it positions a common mechanical shaft in the control valve, allowing hydraulic pressure to be supplied to the hydraulic motors (normally one pilot valve is enabled and moves the other two). The hydraulic motors are reversible, allowing the body flap to be positioned up or down. The hydraulic brake associated with each hydraulic motor releases the hydraulic motor for rotation. When the desired body flap position is reached, the control valves block the hydraulic pressure to the hydraulic motor and apply the hydraulic brake, holding that hydraulic motor at that position. Each hydraulic motor provides the power output to a differential gearbox, which drives a rotary shaft, and four rotary actuators, which position the body flap. The rotary position transducer associated with each rotary actuator provides position feedback to the ASAs; the fourth ASA is used to provide position feedback to the flight control system software.
If the FCS channel switches are in auto , the ASAs will isolate a body flap channel through the solenoid-operated enable valve if the corresponding solenoid-operated pilot valve malfunctions or the control valve associated with the pilot valve does not provide the proper response and allows the hydraulic pressure fluid to recirculate. The FCS channel switches and FCS channel caution and warning light function the same as for the elevons. If the hydraulic system associated with the hydraulic motor fails, the remaining two hydraulic motors will position the body flap, and the hyd press caution and warning light will be illuminated. The body flap deflection rate is approximately 4.5 degrees per second.
Each ASA is hard-wired to a flight MDM. Flight control commands originate from guidance software or from controllers. These inputs go to the flight control software, where they are augmented and then routed to the ASAs.
There are several subsystem operating programs associated with the ASA commands and data. The SOPs convert elevon, rudder and speed brake commands from flight control software from degrees to millivolts; set commands to body flap valves based on an enable command from body flap redundancy management and up/down commands from flight control; convert position feedback to degrees for the elevons, rudder, speed brake and body flap; compute elevator position from elevon position feedbacks; calculate body flap and speed brake deflections as percentages; calculate elevon and rudder positions for display on the surface position indicator; monitor the FCS channel switches and if any are positioned to override, set the override command for that ASA; monitor hydraulic system pressures for failures; and rate-limit aileron and elevator commands according to the number of failures.
Each ASA is mounted on a cold plate and cooled by the Freon-21 coolant loops. Each is 20 inches long, 6.4 inches high and 9.12 inches wide and weighs 30.2 pounds.
The ASA contractor is Honeywell Inc., Clearwater, Fla.
The digital autopilot is the heart of flight control software. It is composed of several software modules that interpret maneuver requests; compare them to what the vehicle is doing; and generate commands for the appropriate effectors, as needed, to satisfy the requests. There are different DAPS for different flight phases and various modes and submodes within each.
At main engine cutoff, the transition digital autopilot becomes active, sending attitude hold commands to the reaction control system. External tank separation is automatically commanded 18 seconds after main engine cutoff, and the transition DAP immediately sets commands to fire the orbiter's minus Z RCS jets, causing the orbiter to translate in the minus Z direction. When a rate of negative 4 feet per second is reached, RCS fire commands are removed. The transition DAP is used from MECO until transition to OPS 2 (on orbit).
In the transition DAP mode, the external tank separation module initialized by the external tank separation sequencer compares Z-delta velocities from the DAP attitude processor with an initialized-loaded desired Z-delta velocity. Before this value is reached, the transition DAP and steering processor send commands to the RCS jet selection logic, which uses a table lookup technique for the primary RCS jets and commands 10 RCS jets to fire in the plus Z direction. When the desired Z-delta velocity is reached, the translation command is set to zero. Rotation commands are permitted during the external tank separation sequence.
The RCS reaction jet driver forward and aft assemblies provide the turn-on/turn-off jet selection logic signals to the RCS jets. There is also a driver redundancy management program that permits only ''good'' RCS jets to be turned on. The RCS jet yellow caution and warning light indicates a failed-on, failed-off or leaking RCS jet.
The 19 RCS jets thrusting in the plus or minus Z direction provide Z translation and roll or pitch rotation control and are considered independent of other axes. The 12 RCS jets thrusting in the plus or minus Y direction provide Y translation and yaw rotation only. The seven RCS jets thrusting in the plus or minus X direction provide X translation only.
Insertion flight control is accomplished using the transition DAP. The transition DAP uses commands from guidance for automatic maneuvers to orbital maneuvering system burn attitude using RCS jets. During OMS-1 and OMS-2 (or only OMS-1 in a direct insertion), the transition DAP uses the OMS engines and RCS jets, as required. The transition DAP also receives commands from the flight crew through the commander's THC and the commander's or pilot's RHC. The transition DAP then takes these commands and converts them into appropriate RCS commands. The transition DAP monitors the resultant attitude and attitude rates and sends the necessary commands to achieve the targeted attitude and attitude rate within premission-specified dead bands.
The transition DAP reconfiguration logic controls the moding, sequencing and initialization of the control law modules and sets gains, dead bands and rate limits. The steering processor is the interface between the guidance or manual steering commands and the transition DAP. The steering processor generates commands to the RCS processor, which generates the RCS jet command required to produce the commanded spacecraft translation and rotation using attitude and rotational rate signals or translation or rotation acceleration commands.
The flight crew interfaces with the transition DAP through the forward RHCs and THC and indirectly through entries to the OMS mnv exec CRT displays and the DAP panel push button light indicators on panel C3.
In the transition DAP, the commander's THC is active and totally independent of the DAP push button light indicators or RHC position or status. Whenever the commander's THC is out of detent plus or minus X, Y or Z, translation acceleration commands are sent directly to the RCS jet selection logic for continuous RCS jet firing. Rotational commands may be sent simultaneously with translation commands within the limits of the RCS jet selection logic; if both plus X and minus Z translations are commanded simultaneously, plus X translation receives priority.
For rotations, the flight crew can select either automatic or manual control through the use of the DAP panel push button light indicators or by moving the RHC. In manual, the capability exists to rotate in any axis in a pulse mode, in which each RHC deflection results in a single burst of jet fire, or in a discrete rate mode, in which RHC deflection results in a specified rate being commanded in that axis for the entire time the RHC is deflected. It is also possible to go to a free drift mode, in which no RCS jets are fired, or to an attitude hold mode, in which the DAP sends commands to maintain the current attitude with null rates within premission-specified dead bands. Also, if the RHC is deflected beyond a certain point, continuous RCS jet firings will result. In translation, movement of the THC results in continuous jet firings.
For the OMS thrusting period, the orbital state (position and vector) is produced by navigation incorporating inertial measurement unit delta velocities during powered and coasting flight. This state is sent to guidance, which uses target inputs through the CRT to compute thrust direction commands and commanded attitude for flight control and thrusting parameters for CRT display. Flight control converts the commands into OMS engine gimbal angles (thrust vector control) for an automatic thrusting period. OMS thrust vector control for normal two-engine thrusting is entered by depressing the orbital DAP auto push button light indicator with both RHCs within software detents. OMS manual thrust vector control for both OMS engines is entered by depressing the orbital man DAP push button light indicator or by moving the commander's or pilot's RHC out of detent; the flight crew supplies the rate commands to the TVC system instead of guidance. The manual RHC rotation requests are proportional to RHC deflections and are converted into gimbal angles. OMS thrust in either case is applied through the spacecraft's center of gravity.
The DAP controls the orbiter in response to automatic or manual commands during insertion and on orbit. The effectors used to produce control forces and moments on the orbiter are the two OMS engines and the 38 primary RCS engines. The forward and aft RCS engines also provide attitude control and three-axis translation during external tank separation, insertion and on-orbit maneuvers and roll control for a single-OMS-engine operation. The OMS provides propulsive and three-axis control for orbit insertion, orbit circularization, orbit transfer and rendezvous. Failure of a single OMS engine will not preclude a nominal orbit insertion.
Before OMS ignition, the spacecraft is maneuvered to the OMS ignition attitude by using the RHC and RCS jets, reducing transient fuel losses. Normally, the first OMS thrusting period raises the orbiter's low elliptical orbit after the external tank is jettisoned, and the second OMS thrusting places the spacecraft into the circular orbit designated for that mission. For orbital maneuvers that use the OMS, any delta velocities greater than 6 feet per second use the two OMS engines. Some missions use a direct insertion, requiring only one OMS thrusting period.
Automatic thrust vector control for one OMS engine is identical to that for two, except that the RCS processor is responsible for roll control. Single-OMS-engine thrust is also through the spacecraft's center of gravity, except when pitch or yaw rate commands are non-zero. If the left or right OMS engine fails, an OMS TVC red light on panel F7 will be illuminated.
The orbital flight control software includes an RCS DAP, an OMS TVC DAP and an attitude processor module to calculate vehicle attitude as well as logic to govern the selection of a DAP. The attitudes calculated by the attitude processor are displayed on the attitude display indicator along with another crew display, universal pointing, which is available in major mode 201 (orbit coast). The vehicle attitude is used by the DAP to determine attitude and rate errors.
The only time the RCS DAP is not used in OPS 2 is during an OMS burn. This DAP controls vehicle attitudes and rates through the use of RCS jet fire commands. Either the larger primary jets or the less powerful vernier jets are used for rotational maneuvers, depending on whether norm or vern is selected on the panel C3 orbital DAP panel. The choice of primary or vernier thrusters depends on fuel consumption considerations and how quickly the vehicle needs to be maneuvered to satisfy a mission objective.
The rotation rates and dead bands, translation rate and certain other DAP options can be changed by the flight crew during the orbit phase using the DAP CRT display. The flight crew can load the DAP with these options in two ways: one option set may be accessed by depressing the DAP A push button on the orbital DAP panel, the other by depressing the DAP B push button. For convenience, each planned DAP configuration is given a number and is referred to by that number and the DAP used to access it. Typically, the DAP A configurations will have larger dead bands and higher rates than the DAP B configurations. The wide dead bands are used to minimize fuel usage, while the tight dead bands allow greater precision in executing maneuvers or in holding attitude.
The RCS DAP can operate in both an automatic and a manual rotation mode, depending on whether the flight crew selects the auto or man push button light indicators on the orbital DAP panel. The manual mode is also accessed when the RHC is moved out of its detent (neutral) position. In both the automatic and manual modes, the rotation rate is controlled by the selection of DAP A or B and the information loaded in the DAP config display. In addition, in automatic, the DAP determines the required attitude to be achieved from universal pointing and then computes the RCS jet fire commands necessary to achieve these requirements within the current set of dead bands. In the manual rotation mode, the RCS DAP converts flight crew inputs with any of the three RHCs to RCS jet fire commands, depending on whether pulse, disc rate or accel is selected on the orbital DAP panel. Simply, when pulse is selected, a single burst of jet fire is produced with each RHC deflection. The resultant rotational rate is specified on the DAP config display. When disc rate is selected, jet firings continue to be made as long as the RHC is out of detent in order to maintain the rotational rate specified on the DAP config display. When accel is selected, continuous jet firings are made as long as the RHC is out of detent.
Another manual RCS DAP mode, local vertical/local horizontal, is used to maintain the current attitude with respect to the rotating LVLH reference frame. It is selected through the LVLH push button on the orbital DAP panel.
The RCS DAP has only a manual translation capability, which is executed through the forward or aft THC. Only the primary RCS jets are used. Deflections of the THC result in RCS jet firings based on the transition DAP mode push button light indicator selected on the orbital DAP panel. Pulse results in a single burst of jet fire. Norm results in continuous jet firings with a specified subset of the available jets. High results in all up-firing jets firing continuously in a Z translation. And low enables a special technique that accomplishes a Z translation using the forward- and aft-firing RCS jets in order to not fire directly toward a target (avoiding plume impingement and contamination of a target payload).
The OMS thrust vector control DAP is available when an OMS burn is executed in major mode 202 (maneuver execute) through the orbit mnvr exec display. The TVC DAP uses the guidance-generated velocity requirements and converts these into the appropriate OMS gimbal commands to achieve this target, assuming auto is selected on the orbital DAP panel. It generates the OMS fire commands; the OMS shutdown commands; and, if necessary due to OMS engine failure, required RCS commands to maintain attitude control. If manual is selected, the TVC DAP uses inputs from the RHC to control attitude during the burn.
As with the transition DAP, there are many subtleties in the operation of the orbital DAP.
There are 24 orbital DAP push button light indicators on panels C3 and A6. Assuming no electrical, computer bus (MDM) or hardware failures that could affect the operation of the push button light indicators, inputs made to one panel will be reflected in the configuration of the other panel. All of the push button light indicators are active in the orbital DAP, but only a subset of these are operational in the transition DAP or when the backup flight system is engaged. None of the push button light indicators are operational during ascent or entry. As with other aft flight deck controls, aft panel A6 push button light indicators are only operational while the vehicle is on orbit.
The orbital DAP select, control, RCS jets, and manual mode translation and rotation push button light indicators are illuminated by flight control when that mode is implemented in the flight control system in the transition or orbital DAP.
The orbital DAP select A or B push button light indicator selects the values the DAP will use from the DAP configuration parameter limits CRT display software loads. The values of attitude dead band, rate dead band and vehicle change in rotation rate are a function of the DAP selection (A or B) and RCS jet selection ( norm or vern ). The select A push button light indicator is illuminated when depressed and select B is extinguished. If select B is depressed, select B is illuminated and select A is extinguished.
When the automatic mode is selected by depressing the auto push button light indicator, the indicator is illuminated and the man push button light indicator is extinguished.
Automatic rotation commands are supplied by the universal pointing processor. The universal pointing processor, through the operational sequence display, provides three-axis automatic maneuver, tracking local vertical/local horizontal about any body vector, rotating about any body vector at the DAP discrete rate, and stopping any of these options and commanding attitude hold. The parameters of these maneuvers are displayed in current attitude, required attitude, attitude error and body rates. Either total or DAP attitude errors may be selected for display on the attitude display indicator error needles.
The automatic maneuver option is used to calculate a commanded vehicle attitude and angular rate or to hold a vehicle attitude. The desired inertial commanded and rotation attitude is input into the operational sequence display in pitch, yaw and roll. When the maneuver option is selected, universal pointing sends the required attitude increment and body rate to flight control, and flight control performs the maneuver when the DAP is in automatic.
The automatic rotation option calculates a rotation about a desired body axis. This option is used for passive thermal control, also known as barbecue. Pitch and yaw body components of the desired rotational axis are first input. The orbiter is maneuvered automatically or manually so that the rotational axis is oriented properly in inertial space. When the rotational option is selected, universal pointing will calculate the required body attitude and send it to flight control. Flight control performs the maneuver if the DAP is in automatic.
The LVLH automatic option calculates the attitude necessary to maintain LVLH with a desired body orientation. LVLH, which is available only on orbit in automatic or manual, calculates the attitude necessary to track the center of the Earth with a given body vector. First, pitch and yaw body components of the desired pointing vector are input. Then omicron roll angle about the body vector is input. When the LVLH option is selected, the required LVLH attitude and its associated maneuver are calculated and sent to flight control. When the LVLH attitude is reached, universal pointing will calculate an attitude and send it to flight control. Flight control performs the maneuver if the DAP is in automatic.
The automatic stop/attitude hold option cancels universal pointing processing of the automatic maneuver, rotation or LVLH options. When the stop/attitude hold option is selected, universal pointing will cancel the processing of the maneuver options and send the current attitude to flight control. When flight control is in automatic, attitude hold will be initiated about the current attitude.
The nine rotation push button light indicators on panel C3 or A6 are meaningless when auto is selected. The transition or orbital DAP can be switched to manual by depressing the man push button light indicator or by positioning the RHC out of detent while operating in the automatic mode. In the manual mode, the man light is illuminated and the auto light is extinguished. When the manual mode is selected, the nine rotation roll, pitch and yaw push button light indicators determine the kind of control the RHC will provide.
There are three RCS rotation submodes available in the orbital DAP: automatic, manual or LVLH. LVLH is not available in the transition DAP; thus the choice in transition DAP is automatic or manual. Within each of these submodes are submodes that depend on conditions, such as the DAP push button light indicator configuration and the RHC state. Manual RCS translation modes are independent.
Depressing the RCS jets norm push button enables the primary RCS jets for rotational and translational firings and disables the vernier RCS jets. The RCS jets norm light is illuminated and the RCS jets vern light is extinguished.
Depressing the RCS jets vern push button enables the vernier RCS jets for rotational (not translational) RCS firings and disables the primary RCS jets. The RCS jets vern light is illuminated and the RCS jets norm light is extinguished.
The manual mode rotation, roll, pitch and yaw push button light indicators are used on orbit and during transition DAP operations. The RCS rotation is selected on an axis-by-axis basis. For example, pitch could be in discrete rate, yaw in acceleration and roll in pulse.
Depressing the rotation disc rate push button for an axis causes the appropriate primary or vernier RCS jets to fire to attain a predetermined rotational rate in that axis while the RHC is out of detent. When the RHC is returned to detent, the rate is nulled and attitude hold is re-established. When depressed, the disc rate push button light indicator is illuminated and the accel or pulse indicators are extinguished for that axis. Rotation units are in degrees per second.
Depressing the rotation accel push button for an axis causes the primary or vernier jets to fire when the RHC is out of detent, producing a moment with the same sense as the RHC deflection in that axis. The jets remain on as long as the RHC is out of detent and shut off when the RHC is returned to detent, allowing attitude to drift freely. When depressed, the accel push button light indicator is illuminated and the disc rate or pulse push button light indicators are extinguished for that axis. The accel push button light indicator is not functional in the transition DAP, but the same effect can be achieved by taking the RHC beyond the softstop in any mode.
Depressing the rotation pulse push button for an axis causes the primary or vernier jets to fire for preset increments in response to each deflection of the RHC in that axis. No further firing occurs until the RHC is returned to detent and is again deflected, allowing attitude to drift. When depressed, the pulse push button light indicator is illuminated and the disc rate or accel push button light indicators are extinguished for that axis. The units for rotation pulse are in degrees per second, and the resulting vehicle rate will be a product of the pulse size and number of pulses.
Depressing the LVLH push button light indicator places the DAP in a manual mode. With rotation in discrete rate for all three axes and the RHC in detent, LVLH hold will be maintained and the LVLH push button light indicator will remain on.
The manual mode translation push button light indicators are used on orbit, and mixed modes are permitted on an axis-by-axis basis. Except for the case of the automatic minus Z external tank separation firing, all RCS translations must be performed manually. In transition DAP, none of the translation push button light indicators will be illuminated. Depressing the translation high push button light indicator causes all nine up-firing primary plus Z jets to be fired as long as the THC is held out of detent in that axis. For a minus Z translation, the high Z mode functions identically to the normal Z and low Z. The Z high push button light indicators are illuminated and the low Z , Z norm and Z pulse push button light indicators are extinguished.
Depressing the low Z push button inhibits all up-firing jets in order to prevent plume damage to payloads or injury to extra-vehicular activity crew members. If a plus Z translation is requested, plus X and minus X jets are fired simultaneously, producing a downward translation because the X jets are oriented in such a way that they have small plus Z thrust components. Minus Z uses the six down-firing jets with the selected high or norm Z. The jets continue to fire as long as the THC is out of detent. The low Z push button light indicators are illuminated and the Z high push button light indicator is extinguished.
Depressing the norm push button light indicator for an axis causes the appropriate primary jets in that axis to be fired for as long as the THC is out of detent. This mode is used for most RCS translations and propellant dumps. The norm push button light indicators are illuminated and the pulse and Z high indicators are extinguished.
Depressing the pulse push button light indicator for an axis causes the appropriate primary jets to fire for a preset increment in response to each deflection of the THC. The firing duration is a function of pulse size. No further firing occurs until the THC is returned to detent and is again deflected. The pulse push button light indicators are illuminated and the norm or Z high indicators are extinguished.
Following insertion, a spacecraft's orbit is essentially fixed, although effects, such as venting and atmospheric drag, can cause orbital perturbations. OMS or RCS thrusting periods can be used to correct or modify the orbit, as required, for mission operations. The direction and magnitude of the thrusting period, as well as the time of application, determine the resulting shape of the orbit.
A posigrade thrusting period increases the speed at the point of application and will raise every point of the orbit except the thrusting point. A retrograde thrusting period decreases the speed at the point of application and will lower every point of the orbit except the thrusting point.
An out-of-plane thrusting period alters the inclination of the spacecraft's orbital plane. It does not change the vehicle's period of orbit or height above the Earth.
A radial thrusting period is one in which the thrust is applied in a direction perpendicular to the spacecraft's velocity vector and in the vehicle's orbital plane. With the vehicle in a circular orbit, a radial thrusting period would be applied along the radius vector either toward or away from the center of the Earth.
The various black boxes of the avionics systems are located in the crew compartment flight deck, middeck avionics bays and the aft avionics bays.
The dedicated displays provide the flight crew with information required to fly the vehicle manually or to monitor automatic flight control system performance. The data on the dedicated displays may be generated by the navigation or the flight control system software or more directly by one of the navigation sensors. The dedicated displays are located in front of the commander's and pilot's seats and on the aft flight deck panel by the aft-facing windows.
The dedicated displays are the attitude director indicators on panels F6, F8 and A1; horizontal situation indicators on panels F6 and F8; alpha Mach indicators on panels F6 and F8; altitude/vertical velocity indicators on panels F6 and F8; surface position indicator on panel F7; reaction control system activity lights on panel F6; g-meter on panel F7; and head-up display on the glare-shield in front of the commander's and pilot's seats.
Not all of the dedicated displays are available in every operational sequence or major mode. Their availability is related to the requirements of each flight phase.
The display driver unit is an electronic mechanism that connects the general-purpose computers and the primary flight displays. The DDU receives data signals from the computers and decodes them to drive the dedicated displays. The unit also provides dc and ac power for the ADIs and the rotational and translational hand controllers. It contains logic for setting flags on the dedicated instruments for such items as data dropouts and failure to synchronize. The orbiter contains three DDUs: one at the commander's station, one at the pilot's station and one at the aft station.
All display parameters, regardless of their origin, are ultimately processed through the dedicated display processor software (except for the g-meter, which is totally self-contained). The display parameters are then routed to the respective displays through either a DDU or multiplexer/demultiplexer; DDUs send data to the ADI, HSI, AMI and AVVI displays, while MDMs provide data for the SPI and RCS activity lights.
There are three display driver units. One interfaces with the ADI, HSI, AVVI and AMI displays on panel F6 at the commander's station, and the second interfaces with the same instruments on panel F8 at the pilot's station. The third unit interfaces with the ADI at the aft flight station.
Associated with each DDU is a data bus select switch. The commander's switch is on panel F6, and the pilot's is on panel F8. The select switch for the aft flight station is on panel A6. Positions 1, 2, 3 and 4 allow the flight crew to select any one of four forward flight-critical data buses (FC1 through 4) as the data source for that DDU and its dedicated displays. Because the flight-critical data buses are dedicated to specific orbiter general-purpose computers, the data bus select switch also provides a means of assessing the health of individual computers, if they are assigned to FC1, 2, 3 or 4.
The commander's attitude director indicator is powered from the main bus A and B DDU circuit breakers on panels O14 and O15 through DDU 1 power supply D, which provides ac and dc power. The pilot's ADI is powered from the main B and C DDU circuit breakers on panels O15 and O16 through DDU 2 power supply D, which also provides ac and dc power. The aft flight station ADI is powered from the main A and C DDU circuit breakers on panels O14 and O16 through DDU 3 power supply D, which provides ac and dc power.
The instrument power flt MPS/off/flt switch on panel F6 supplies main bus A power to the commander's HSI, AMI and AVVI displays; the single SPI; and the main propulsion instruments when positioned to flt MPS . The instrument power on/off switch on panel F8 supplies main bus B power to the pilot's HSI, AMI and AVVI displays and the hydraulic and auxiliary power unit displays.
The RCS activity lights receive power from annunciator control assemblies.
The DDU contractor is Rockwell International, Collins Radio Group, Cedar Rapids, Iowa.
The commander's and pilot's ADIs are supported throughout the mission, while the aft ADI is active only during orbital operations. They give the crew attitude information as well as attitude rate and attitude errors, which can be read from the position of the pointers and needles. Each ADI has a set of switches by which the crew can select the mode or scale of the readout. The commander's switches are located on panel F6, the pilot's on panel F8 and the aft switches on panel A6.
The orbiter's attitude is displayed to the flight crew by an enclosed ball (sometimes called the eight ball) that is gimbaled to represent three degrees of freedom. The ball, covered with numbers indicating angle measurements (a zero is added as the last digit of each), moves in response to software-generated commands to depict the current orbiter attitude in terms of pitch, yaw and roll.
The ADI attitude select switch determines the unit's frame of reference: inrtl (inertial), LVLH (local vertical/local horizontal), and ref (reference). The inrtl position allows the flight crew to view the orbiter's attitude with respect to the inertial reference frame, useful in locating stars. The LVLH position shows the orbiter's attitude from an orbiter-centered rotating reference frame with respect to Earth. The ref position is primarily used to see the orbiter's attitude with respect to an inertial reference frame defined when the flight crew last depressed the att ref push button. It is useful when the crew flies back to a previous attitude or monitors a maneuvering system thrusting period for attitude excursions. The two forward switches are active during ascent, orbital and transition flight phases but have no effect during entry, the latter part of a return to launch site or phases when the backup flight system is driving the ADIs. The aft switch, like the aft ADI, is operational only in orbit.
Each attitude director indicator has a set of three rate pointers that provide a continuous readout of vehicle body rotational rates. Roll, pitch and yaw rates are displayed on the top, right and bottom pointers, respectively. The center mark on the graduated scale next to the pointers shows zero rates, while the rest of the marks indicate positive or negative rates. The adi rate switch for each indicator unit determines the magnitude of full-scale deflection. When this switch is positioned to high (the coarsest setting), the pointer at the end of the scale represents a rotation rate of 10 degrees per second. When the switch is positioned to med, a full-range deflection represents 5 degrees per second. In the low position (the finest setting), a pointer at either end of the scale is read at a rate of 1 degree per second. These pointers are ''fly to'' in the sense that the rotational hand controller must be moved in the same direction as the pointer to null a rate.
ADI rate readings are independent of the selected attitude reference. During ascent, the selected rates come directly from the solid rocket booster or orbiter rate gyros to the ADI processor for display on the rate pointers. During entry, only the pitch rate follows the direct route to the ADI display. The selected roll and yaw rates first flow through flight control, where they are processed and output to the ADI as stability roll and yaw rates. (This transformation is necessary because, in aerodynamic flight, control is achieved about stability axes, which in the cases of roll and yaw differ from body axes.)
Three needles on each attitude director indicator display vehicle attitude errors. These needles extend in front of the ADI ball, with roll, pitch and yaw arranged just as the rate pointers are. Like the rate indicators, each error needle has a background scale with graduation marks that allow the flight crew to read the magnitude of the attitude error. The errors are displayed with respect to the body-axis coordinate system and, thus, are independent of the selected reference frame of the attitude display.
The ADI error needles are driven by flight control outputs that show the difference between the required and current vehicle attitude. These needles are also ''fly to,'' meaning that the flight crew must maneuver in the direction of the needle to null the needle. For example, if the pitch error needle points down, the flight crew must manually pitch down to null the pitch attitude error. The amount of needle deflection indicating the number of degrees of attitude error depends upon the adi error switch for each ADI. In the high position, the error needles represent 10 degrees, med represents 5 degrees and low represents 1 degree.
At the aft flight station on panel A6, the aft sense switch allows the flight crew to use the aft ADI, RHC and translational hand controller in a minus X or minus Z control axis sense. These two options of the aft ADI and hand controllers correspond to the visual data out of the aft viewing (negative X) or overhead viewing (negative Z) windows.
Each ADI has a single flag labeled off on the left side of the display whenever any attitude drive signal is invalid. There are no flags for the rate and error needles; these indicators are driven out of view when they are invalid.
The ADI contractor is Lear Siegler, Grand Rapids, Mich.
The horizontal situation indicator for the commander and pilot displays a pictorial view of the vehicle's position with respect to various navigation points and shows a visual perspective of certain guidance, navigation and control parameters, such as directions, distances and course/glide path deviation. The flight crew uses this information to control or monitor vehicle performance. The HSIs are active during the entry and landing and ascent/RTLS phases.
Each HSI provides an independent source to compare with ascent and entry guidance, a means of assessing the health of individual navigation aids during entry and information needed by the flight crew to fly manual ascent, RTLS and entry.
Three switches are associated with each horizontal situation indicator. The commander's select switches are on panel F6 and the pilot's are on panel F8. The HSI select mode switch selects the mode-entry, TACAN or approach. The HSI select source switch selects TACAN, navigation or microwave scan beam landing system; its 1, 2, 3 switch selects the data source. When positioned to nav, the HSI is supplied with data from the navigation attitude processor and the 1, 2, 3 switch is not used. In TACAN, the HSI is supplied with data derived from the 1, 2, 3 switch, thus TACAN 1, 2 or 3. In MLS , the HSI is supplied with data derived from the 1, 2, 3 switch, thus MLS 1, 2 or 3.
Each HSI displays magnetic heading (compass card), selected course, runway magnetic course, course deviation, glide slope deviation, primary and secondary bearing, primary and secondary range, and flags to indicate validity.
Each HSI consists of a case-enclosed compass card measuring zero to 360 degrees. At the center of the compass card is an aircraft symbol, fixed with respect to the case and about which the compass card rotates.
The magnetic heading (the angle between magnetic north and vehicle direction measured clockwise from magnetic north) is displayed by the compass card and read under the lubber line located at the top of the indicator dial. (A lubber line is a fixed line on a compass aligned to the longitudinal axis of the craft.) The compass card is positioned at zero degrees (north) when the heading input is zero. When the heading point is increased, the compass card rotates counterclockwise.
The course pointer is driven with respect to the HSI case rather than the compass card. Therefore, a course input (from the DDU) of zero positions the pointer at the top lubber line, regardless of compass card position. To position the course pointer correctly with respect to the compass card scale, the software must subtract the vehicle magnetic heading from the runway azimuth angle (corrected to magnetic north). As this subtraction is done continuously, the course pointer appears to rotate with the compass card, remaining at the same scale position. An increase in the angle defining runway course results in a clockwise rotation of the course pointer.
Course deviation is an angular measurement of vehicle displacement from the extended runway centerline. On the HSI, course deviation is represented by the deflection of the deviation bar from the course pointer line. Full scale on the course deviation scale is plus or minus 10 degrees in terminal area energy management and plus or minus 2.5 degrees during approach and landing. The course deviation indicator is driven to zero during entry. When the course deviation input is zero, the deviation bar is aligned with the end of the course pointer. With the pointer in the top half of the compass card, an increase in course deviation to the left (right) causes the bar to deflect the right (left). Therefore, the course deviation indicator is a fly-to indicator for flying the vehicle to the extended runway centerline. Software processing also ensures that the CDI remains fly to, even when the orbiter is heading away from the runway.
In the TAEM example, at a range of 9 nautical miles (10 statute miles), the CDI would read about 7.5 degrees, with the extended runway centerline to the right of the orbiter. In course deviation geometry, if the orbiter is to the left of the runway, it must fly right (or if the orbiter is to the right of the runway, it must fly left) to reach the extended runway centerline. The corresponding course deviation bar would deflect to the right (or to the left in the latter case). The reference point at the end of the runway is the microwave landing system station. The sense of the CDI deflection is a function of vehicle position rather than vehicle heading.
Glide slope deviation, the distance of the vehicle above or below the desired glide slope, is indicated by the deflection of the glide slope pointer on the right side of the HSI. An increase in glide slope deviation above (below) the desired slope deflects the pointer downward (upward); the pointer is a fly-to indicator. In the HSI example, the pointer shows the vehicle to be below the desired glide slope by about 4,000 feet (in TAEM, each dot represents 2,500 feet).
The ''desired glide slope'' is actually only a conceptual term in HSI processing. At any instant, glide slope deviation is really the difference between the orbiter altitude and a reference altitude computed in the same fashion as the guidance reference altitude. Also included in the reference altitude equation are factors for a ''heavy orbiter'' and for high winds.
The GSI computation is not made during entry or below 1,500 feet during approach and landing; therefore, the pointer is stowed and the GSI flag is displayed during those intervals.
The primary and secondary bearing pointers display bearings relative to the compass card. These bearings are angles between the direction to true or magnetic north and to various reference points as viewed from the orbiter. For the bearing pointers to be valid, the compass card must be positioned in accordance with vehicle heading input data.
When the bearing inputs are zero, the pointers are at the top lubber line, regardless of compass card position. Like the course pointer, the bearing pointer drive commands are developed by subtracting the vehicle heading from the calculated bearing values. This allows the pointers to be driven with respect to the HSI case but still be at the correct index point on the compass card scale. When the bearing inputs are increased, the pointers rotate clockwise about the compass card. The pointer does not reverse when it passes through 360 degrees in either direction.
For example, if the primary bearing is 190 degrees and the secondary bearing is 245 degrees, the bearing reciprocals are always 180 degrees from (opposite) the pointers. The definition of primary and secondary bearing varies with the flight regime.
The HSI is capable of displaying two four-digit values in the upper left and right side of its face. These numbers are called primary and secondary range, respectively. Each display ranges from zero to 3,999 nautical miles (4,602 statute miles). While their meaning depends on the flight regime, both numbers represent range in nautical miles from the vehicle to various points relative to the primary and secondary runways. In the HSI example, the primary range is 9 nautical miles (10 statute miles); the barberpole in the secondary range slot is an invalid data indication.
The HSI has four flags- off, brg (bearing), GS (glide slope) and CDI-and two barberpole indications that can respond to separate DDU commands, identifying invalid data. Off indicates that the entire HSI display is invalid because of insufficient power. Brg indicates invalid course, primary bearing, and/or secondary bearing data. GS indicates invalid glide slope deviation. CDI indicates invalid course deviation data. Barberpole in the range slots indicates invalid primary or secondary range data.
When the HSI source switch is in nav , the entire HSI display is driven by navigation-derived data from the orbiter state vector. This makes the HSI display dependent on the same sources as the navigation software (IMU, selected air data, selected navigational aids), but the display is independent of guidance targeting parameters. As stated previously, when the TACAN/nav/MLS switch is in the nav position, the source 1, 2, 3 switch is not processed.
The TACAN or MLS position of the source switch should be used only when TACAN or MLS data are available. TACAN data can be acquired in Earth orbit but would be unavailable during blackout; therefore, TACAN is generally not selected until acquisition after blackout. MLS has a range of 20 nautical miles (23 statute miles) and is normally selected after the orbiter is on the heading alignment cylinder.
The glide slope deviation pointer is stowed when the entry mode is selected and the flag is displayed. The GSI in TAEM indicates deviation from guidance reference attitude in plus or minus 5,000 feet. The GSI in approach indicates guidance reference altitude for approach and landing in plus or minus 1,000 feet; it is not computed below 1,500 feet and the flag deploys.
In the entry mode, the compass card heading indicates the magnetic heading of the vehicle's relative velocity vector. In TAEM and approach, the compass card indicates magnetic heading of the body X axis.
In the entry mode, the course deviation indicator is a valid software zero with no flag. In TAEM, the CDI indicates the deviation from the extended runway centerline, plus or minus 10 degrees. In approach, the CDI indicates the deviation from the extended runway centerline, plus or minus 2.5 degrees.
In the entry mode, the primary bearing indicates the spherical bearing to way point 1 for the nominal entry point at the primary landing runway. The secondary bearing indicates the spherical bearing to WP-1 for the NEP to the secondary landing runway. In TAEM, the primary bearing indicates the bearing to WP-1 on selected HAC for the primary runway. The secondary bearing indicates the bearing to the center of the selected HAC for the primary runway. In approach, the primary and secondary bearings indicate the bearing to WP-2 at the primary runway.
In the entry mode, the primary range indicates the spherical surface range to WP-2 on the primary runway via WP-1 for NEP. The secondary range indicates the spherical surface range to WP-2 on the secondary runway via WP-1 for NEP. In TAEM, the primary range indicates the horizontal distance to WP-2 on the primary runway via WP-1. The secondary range indicates the horizontal distance to the center of the selected HAC for the primary runway. In approach, the primary and secondary ranges indicate the horizontal distance to WP-2 on the primary runway.
During ascent major modes 102 and 103 (first and second stage) and RTLS, the horizontal situation indicator provides information about the target insertion orbit. The compass card displays heading with respect to TIO, and north on the compass card points along the TIO plane. The heading of the body plus X axis with respect to the target insertion orbit is read at the lubber line.
The course pointer provides the heading of the Earth-relative velocity vector with respect to the TIO plane. The CDI deflection indicates the estimated sideslip angle, the angle between the body X axis and the relative velocity vector.
The primary bearing pointer during major modes 102 and 103 is fixed on the compass card at a predetermined value to provide a turnaround heading in the event of an RTLS abort. During RTLS major mode 601, the pointer indicates the heading to the landing site runway. The secondary bearing provides the heading of the inertial velocity vector with respect to the TIO plane.
The horizontal situation CRT display allows the flight crew to configure the software for nominal winds or high head winds. The software item entry determines the distance from the runway threshold to the intersection of the glide slope with the runway centerline. The high-wind entry pushes the intercept point close to the threshold. The distance selected is factored into the computation of reference altitude from which the GSI is derived.
The HSI contractor is Rockwell International, Collins Radio Group, Cedar Rapids, Iowa.
The two alpha Mach indicators are located next to the attitude director indicators on panels F6 and F8. The AMIs consists of four tape meters displaying angle of attack ( alpha ), vehicle acceleration (accel), vehicle velocity ( M/vel ) and equivalent airspeed ( EAS ). The two units are driven independently but can have the same data source.
Alpha displays vehicle angle of attack, defined as the angle between the vehicle plus X axis and the wind-relative velocity vector (negative wind vector). Alpha is displayed by a combination moving scale and moving pointer. For angles between minus 4 degrees and plus 28 degrees, the scale remains stationary and the pointer moves to the correct reading. For angles less than minus 4 degrees or greater than plus 28 degrees, the pointer stops (at minus 4 or plus 28 degrees) and the scale moves so that the correct reading is adjacent to the pointer. The alpha tape ranges from minus 18 to plus 60 degrees with no scale changes. The negative scale numbers (below zero) have no minus signs; the actual tape has black markings on a white background on the negative side and white markings on a black background on the positive side.
The accel scale displays vehicle drag acceleration, which is the deceleration along the flight path. This is a moving tape upon which acceleration is read at the fixed lubber line. The tape range is minus 50 to plus 100 with a scale change at zero feet per second squared. Minus signs are assumed on the accel scale also; the negative region has a black background and the positive side has a white background.
The M/vel scale displays Mach number or relative velocity. Mach number is the ratio of vehicle airspeed to the speed of sound in the same medium. Relative velocity in this case is the vehicle airspeed. The actual parameter displayed is always Mach number; the tape is simply rescaled above Mach 4 to read relative velocity in thousands of feet per second (above 2,000 feet per second, Mach number = V REL /1,000). The M/vel scale is a moving tape from which Mach/velocity is read at the fixed lubber line. The scale ranges from zero to 27 with a scale change at Mach 4.
The EAS scale is used to display equivalent airspeed. On the moving-tape scale, equivalent airspeed is read at the fixed lubber line. The tape range is zero to 500 knots, and scaling is 1 inch per 10 knots.
Each scale on the AMI displays an off flag if the indicator malfunctions, invalid data are received at the DDU or a power failure occurs (all flags appear).
The air data source select switch on panel F6 for the commander and panel F8 for the pilot determines the source of data for the AMI and altitude/vertical velocity indicator. The nav position of the air data switch ensures that the alpha , Mach and EAS on the AMI are the same parameters sent to guidance, flight control, navigation and other software users; accel comes from navigation software.
The left, right position of the air data switch selects predetermined data from the left or right air data probe assembly after deployment of the left and right air data probes at Mach 3 for alpha, M/vel and EAS display. Accel is always derived from navigation software during entry. It is driven to zero during terminal area energy management and approach and landing.
The altitude/vertical velocity indicators are located on panel F6 for the commander and panel F8 for the pilot. These indicators display vertical acceleration ( alt accel ), vertical velocity ( alt rate ), barometric altitude ( alt ) and radar altitude ( rdr alt ).
The alt accel indicator, which displays altitude acceleration of the vehicle, is read at the intersection of the moving pointer and the fixed scale. The scale range is minus 13.3 to 13.3 feet per second squared, and the scaling is 6.67 feet per second squared per inch. Software limits acceleration values to plus or minus 12.75 feet per second squared.
The alt rate scale displays vehicle altitude rate, which is read at the intersection of the moving tape and the fixed lubber line. The scale range is minus 2,940 to plus 2,940 feet per second with scale changes at minus 740 feet per second and plus 740 feet per second. The negative and positive regions are color-reversed: negative numbers are white on a black background and positive numbers are black on white.
The alt scale, a moving tape read against a fixed lubber line, displays the altitude of the vehicle above the runway (barometric altitude). The scale range is minus 1,100 feet to plus 165 nautical miles (189 statute miles), with scale changes at minus 100, zero, 500 feet and plus 100,000 feet. The scale is in feet from minus 1,100 to plus 400,000 and in nautical miles from plus 40 to plus 165 (46 to 189 statute miles). Feet and nautical miles overlap from plus 40 to plus 61 nautical miles (46 to 70 statute miles).
The rdr alt scale is a moving tape read against a fixed lubber line. It displays radar altitude (corrected to wheels) during major mode 305, below 9,000 feet (normally not locked in until below 5,000 feet; prior to radar altimeter lock-on, the meter is ''parked'' at 5,000 feet). The scale ranges from zero to 9,000 feet with a scale change at 1,500 feet. Each scale on the AVVI displays an off flag in the event of indicator malfunction, invalid data received at the DDU or power failure (all flags appear).
With the air data source switch in the nav position, the alt accel, alt rate, and alt scales are navigation-derived. The rdr alt indicator is controlled by the radar altm switch on panel F6 for the commander and panel F8 for the pilot. Radar altm positioned to 1 selects radar altimeter 1; 2 selects radar altimeter 2.
The air data switch is positioned to left or right to select the right or left air data probe, respectively, after air data probe deployment at Mach 3. The alt and alt rate scales receive information from the selected air data probe. Alt accel receives navigation data. The rdr alt scale receives data from the radar alt select switch.
The surface position indicator is a single display on panel F7 that is active during entry and during the entry portion of RTLS. The SPI displays the actual and commanded positions of the elevons, body flap, rudder, aileron and speed brake.
The four elevon position indicators show the elevon positions in the order of appearance as viewed from behind the vehicle (from left to right: left outboard, left inboard, right inboard, right outboard). The scales all range from plus 20 to minus 35 degrees, which are also the software limits to the elevon commands. The pointers are driven by four separate signals and can read different values, but normally the left pair is identical and the right pair is identical. Positive elevon is below the null line and negative is above.
The body flap scale reads body flap positions from zero to 100 percent of software-allowed travel. Zero percent corresponds to full up (minus 11.7 degrees); 100 percent corresponds to full down (plus 22.5 degrees). The small pointer at 35 percent is fixed and shows the trail position.
Rudder position is displayed as if viewed from the rear of the vehicle. Deflection to the left of center represents left rudder. The scale is plus 30 degrees (left) to minus 30 degrees (right), but software limits the rudder command to plus or minus 27.1 degrees.
The aileron display measures the effective aileron function of the elevons in combination. Aileron position equals the average of the left and right elevon divided by two. Deflection of the pointer to the right of center indicates a roll-right configuration (left elevons down, right elevons up) and vice versa. The scale is minus 5 to plus 5 degrees, with minus 5 at the left side. The aileron command can exceed plus or minus 5 degrees (maximum plus or minus 10 degrees), in which case the meter saturates at plus or minus 5 degrees.
The speed brake position indicator indicates the actual position on the upper scale and commanded position on the lower scale. The position ranges zero to 100 percent; zero percent is fully closed and 100 percent is fully open, which corresponds to 98 degrees with respect to the hinge lines. The small point at 25 percent is fixed and represents the point at which the speed brake surfaces and the remainder of the tail form a smooth wedge.
The speed brake command is scaled identically to position and has the same travel limits. It always represents the speed brake auto guidance command. The off flag is set only for internal meter problems or during OPS 8 display checkout.
These indicators are located on panel F2 for the commander and panel F4 for the pilot. The flight control system's push button light indicators transmit flight crew moding requests to the digital autopilot in the flight control software and reflect selection by illuminating the effective DAP state.
The push button light indicators are used to command and reflect the status of the pitch control mode. The pitch and roll/yaw indicators transmit moding requests to the digital autopilot and indicate the effective state of the pitch, roll and yaw DAP channels by lighting.
Auto indicates that control is automatic and no crew inputs are required. CSS is control stick steering; crew inputs are required but are smoothed by the DAP (stability augmentation, turn coordination).
The spd brk/throt (speed brake/throttle) push button light indicator has two separate lights, auto and man (manual), to indicate that the DAP speed brake channel is in the automatic or manual mode. The push button light indicator transmits only the auto request.
The body flap push button light indicator also has separate auto and man lights, indicating the state of the body flap channel. Like the spd brk/throt push button light indicator, the body flap indicator transmits only the auto request.
The reaction control system command lights on panel F6 are active during the entry and RTLS flight phases. Their primary function is to indicate RCS jet commands by axis and direction; secondary functions are to indicate when more than two yaw jets are commanded and when the elevon drive rate is saturated.
During major modes 301 through 304, up until the roll jets are no longer commanded (dynamic pressure exceeds 10 pounds per square foot), the roll l and r lights indicate that left or right roll commands have been issued by the DAP. The minimum light-on duration is extended so that the light can be seen even during minimum-impulse firings. When dynamic pressure is greater than or equal to 10 pounds per square foot, the roll lights are quiescent until 50 pounds per square foot, after which time both lights are illuminated whenever more than two yaw jets are commanded on.
The pitch u and d lights indicate up and down pitch jet commands until dynamic pressure equals 20 pounds per square foot, after which the pitch jets are no longer used. When dynamic pressure is 50 pounds per square foot or more, the pitch lights, like the roll lights, assume a new function: both light whenever the elevon surface drive rate exceeds 20 degrees per second (10 degrees per second if only one hydraulic system is left).
The g-meter is a self-contained accelerometer and display unit mounted on panel F7. It senses linear acceleration along the Z axis (normal) of the vehicle. A mass weight in the unit is supported vertically by two guide rods and is constrained by a constant-rate helical spring. The inertial force of the mass is proportional to the inertial force of the vehicle and, hence, to the input acceleration, under conditions of constant acceleration. Displacement of the mass is translated to pointer displacement through a rack-and-pinion gear train whose output is linear with input acceleration. The display indicates acceleration from minus 2 g's to plus 4 g's. The g-meter requires no power and has no software interface. Like all the dedicated displays, it has an external variable incandescent lamp.
The head-up display is an optical miniprocessor that cues the commander and/or pilot during the final phase of entry and particularly in the final approach to the runway. With minimal movement of their eyes from the forward windows (head up) to the dedicated display instruments (head down), the commander and pilot can read data from HUDs located in front of them on their respective glareshields. The HUD displays the same data presented on several other instruments, including the ADI, SPI, AMI and AVVI.
The HUD allows out-of-the-window viewing by superimposing flight commands and information on a transparent combiner in the window's field of view. The baseline orbiter, like most commercial aircraft, presents conventional electromechanical display on a panel beneath the glareshield, which necessitates that the flight crew look down for information and then up to see out the window. During critical flight phases, particularly approach and landing, this is not an easy task. In the orbiter, with its unique vehicle dynamics and approach trajectories, this situation is even more difficult.
Since the orbiter is intended to be in service for several years, the addition of a HUD was considered appropriate. Most recent military aircraft include HUD systems, as do several European airliners. Additionally, since the display portion of some existing HUD systems could be easily installed in the orbiter, the HUD system requirements for the orbiter were patterned after existing hardware to minimize development costs.
While the display portion of the orbiter system could be similar to existing HUD systems, the drive electronics could not. Since the orbiter avionics systems are digital and minimal impact on the orbiter was paramount, the HUD drive electronics were designed to receive data from the orbiter data buses. Most existing HUD drive electronics use analog data or a combination analog/digital interface. In the orbiter system, the HUD drive electronics utilize, to the maximum extent possible, the same data that drive the existing electromechanical display devices.
The orbiter display device, designed by Kaiser Electronics of San Jose, Calif., uses a CRT to create the image, which is then projected through a series of lenses onto a combining glass (a system very similar to one they developed and produce for the Cobra jet aircraft). Certain orbiter design requirements, such as vertical viewing angles, brightness and unique mounting, dictated some changes from the Cobra configuration.
A HUD power on/off switch located on the left side of panel F3 provides and terminates electrical power to the commander's HUD. The same switch is also located on the right side of panel F3 for the pilot's HUD.
Each HUD is a single-string system but connected to two data buses for redundancy. It is an electronic/optical device with two sets of combiner glasses located above the glareshield in the direct line of sight of the commander and the pilot. Essential flight information for vehicle guidance and control during approach and landing is projected on the combiner glasses and collimated at infinity.
For example, looking through the HUD and out the window in the final phase of the preflare maneuver, the commander might see EAS = 280 knots (left scale), altitude = 500 feet (right scale), and orbiter heading ( + ) slightly to the left of runway centerline, which indicates a light crosswind from the left. The velocity vector symbol is just crossing the runway overrun. The guidance diamond is centered inside the velocity vector symbol. The flare triangles on the wing tips indicate that the pilot is following the flare command precisely. The lighted outline of the start of the runway zone appears at the top of the combiner. The HUD can display speed brake command and position; discrete messages, such as gear; and, during rollout, deceleration and wing-leveling parameters.
The images, generated by a small CRT and passed through a series of lenses, are displayed to the flight crew on the combiners as lighted symbology. The transmissiveness of the combiner allows the crew to look through it and see actual targets like the runway.
For instance, if the crew is conducting an instrument approach at 7,000 feet on the final approach course in a solid overcast, the base of which is at 5,000 feet, the lighted outline of the runway would be displayed on the combiner. However, when the orbiter exits the overcast at 5,000 feet, the lighted outline of the runway would be superimposed on the real runway. As the orbiter proceeds down the steep glide slope, the velocity vector is superimposed over the glide slope aim point. At preflare altitude, flare triangles move up to command the pullout. The pilot maintains the velocity vector symbol between the triangles. After a short period of stabilized flight on the shallow glide slope, the guidance diamond commands a pitch-up until the nose is about 8 degrees above the horizon, which is essentially the touchdown attitude. After touchdown, during the rollout phase, the crew maintains the approximate touchdown attitude, plus 6 degrees theta (nose above the horizon), until 180 knots equivalent airspeed and then commands a derotation maneuver.
The HUD has proved to be a valuable landing aid and is considered the primary pilot display during this critical flight phase.
Information content from the NSTS Shuttle Reference Manual (1988)
Last Hypertexed Friday April 17 00:05:39 EDT 1998
Jim Dumoulin (firstname.lastname@example.org)
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