The terminal count phase extends from T minus 20 minutes (T refers to lift-off time) through solid rocket booster ignition. Prior to T minus 20 minutes, prelaunch functions are controlled by a ground computer network at the launch site. This is called the launch processing system. The orbiter general-purpose computers interact with the processing system from T minus 20 minutes (when they are loaded with OPS 1 software and formed into the four GPCs referred to as the primary avionics software system) until lift-off. This interaction consists of commands being sent by the LPS to the computers, which carry out the commands and then respond to the LPS action until it is completed.
A planned 10-minute hold occurs at T minus nine minutes, when last-minute launch preparations are made.
Launch countdown is controlled by the LPS until 31 seconds before launch, when the onboard computers' automatic launch sequence software is enabled by launch processing system command. From this point on, the computers take control of the sequencing of events and perform functions by the onboard clock, but will honor hold, resume count and recycle commands from the LPS.
The computer launch sequence sets flags to command the arming of SRB ignition and hold-down release system pyro initiator controllers and the T-0 umbilical release PICs. After a delay, the SRB ignition PIC voltages are monitored for acceptable levels. The hold-down release system PICs and the T-0 umbilical release system PICs are monitored by the LPS. The computer launch sequence logic initiates a countdown hold if the SRB ignition PIC voltages fall below an acceptable level at any time before space shuttle main engine start commands are issued. If the SRB ignition PIC voltages are not acceptable after the start commands are issued, the engines are shut down.
The computer launch sequence also controls certain critical main propulsion system valves and monitors the engine-ready indications from the main engines. The MPS start commands are issued by the onboard GPCs at T minus 6.6 seconds (staggered start- engine 3, engine 2, engine 1-at 120-millisecond intervals), and the sequence monitors the thrust buildup of each engine. All three engines must reach the required 90 percent of thrust at three seconds; otherwise, an orderly shutdown is commanded and safing functions are initiated.
Normal thrust buildup to the required 90-percent thrust level results in the engines being commanded to the lift-off position at T minus three seconds as well as the fire 1 command being issued to arm the SRBs. At T minus three seconds, the vehicle base bending load modes are allowed to initialize (movement of approximately 25.5 inches, measured at the top of the external tank, with movement toward the external tank).
At T minus zero, the two SRBs are ignited, under command of the four onboard GPCs; the four explosive bolts on each SRB are initiated (each bolt is 28 inches long and 3.5 inches in diameter); the two T-0 umbilicals (one on each side of the spacecraft) are retracted; the onboard master timing unit, event timer and mission event timers are started; the three main engines are at 100 percent; and the ground launch sequence is terminated, initiating the onboard GPCs.
The ascent digital autopilot configuration logic, upon ascent initialization, places the main engine nozzles in the start position, places the SRB nozzles in the null position and places the orbiter aerosurfaces in the launch position. Upon receipt of the command to prepare main engines for lift-off, the engines are commanded to the launch position.
There are no guidance functions performed during prelaunch. Also, there are no active flight control functions, except for the commands to gimbal the engines for test and launch positions and slewing of the aerosurfaces.
Navigation is initialized during prelaunch. At T minus eight seconds, the launch pad location (latitude, longitude and attitude) of the navigation base is transformed to a position vector. Launch Complex 39-A at the Kennedy Space Center in Florida is at 28 degrees 36 minutes 29.7014 seconds north latitude and 80 degrees 36 minutes 15.4166 seconds west longitude. The space shuttle vehicle is oriented on the launch pad with its negative Z body axis (tail) pointed south. The inertial measurement units are set inertial at T minus 20 minutes but are biased for Earth rotation until T minus 12 seconds, when the bias is removed. Launch Complex 39-B at the Kennedy Space Center is at 28 degrees 37 minutes 26 seconds north latitude and 80 degrees 37 minutes 15.09 seconds west longitude. The shuttle vehicle is also oriented on this pad with the negative Z body axis (tail) pointed south.
First-stage ascent extends from SRB ignition through SRB separation, or SRB staging. The sequence of major guidance, navigation and control events proceeds as folllows: The vehicle lifts off the pad 0.3 second after SRB ignition, rising vertically in attitude hold until the SRBs' nozzles clear the lightning rod tower by approximately 41 feet. The vehicle begins a combined roll, pitch and yaw maneuver that positions the orbiter head down, with wings level and aligned with the launch pad. The orbiter flies upside down during the ascent phase. This orientation, together with trajectory shaping, establishes a trim angle of attack that is favorable for aerodynamic loads during the region of high dynamic pressure, resulting in a net positive load factor, as well as providing the flight crew with use of the ground as a visual reference. By about 20 seconds after lift-off, the vehicle is at 180 degrees roll and 78 degrees pitch.
During the first 90 seconds of flight, the flight control system provides load relief by making adjustments to reduce vehicle loads at the expense of maintaining a precise trajectory profile. A special schedule of elevon position with respect to velocity is followed to protect the wings from excessive loads and to hold the body flap and rudder/speed brake in place. The surface position indicator displays the position of the aerosurfaces. To keep the dynamic pressure on the vehicle below a specified level, on the order of 580 pounds per square foot (max q), the main engines are throttled down at approximately 26 seconds and throttled back up at approximately 60 seconds. This also reduces heating on the vehicle. Because of the throttling at this time, the term ''thrust bucket'' evolved. Maximum dynamic pressure occurs shortly after throttle up.
Thrust vector control is the hub of flight control. In the ascent phase, the four ascent thrust vector control drivers respond to commands from the guidance system. Thus the TVC commands from guidance are transmitted to the ATVC drivers, which transmit electrical signals proportional to the commands to the servoactuators on each main engine and solid rocket booster.
Thrust vector control closes the acceleration and rate loops within the outer attitude loops to generate body axis attitude error rates, which eventually are nulled out by the main engines and SRBs. The main propulsion system processor of the digital autopilot converts body axis and attitude error signals generated in TVC into pitch and yaw nozzle deflection commands for the main engines. The SRB processor of the DAP accomplishes the same functions as the MPS, except that it is referred to as rock and tilt instead of pitch and yaw.
The SRB pitch and yaw rate gyros are used exclusively during first-stage ascent, and control is switched to the orbiter rate gyros when the SRBs are commanded to null in preparation for separation. Pitch and yaw axes and a combination of rate, attitude and acceleration signals are blended to effect a common signal to the main engines and the TVC for both SRBs. In the roll axis, rate and attitude are summed to provide a common signal to both the main engines' and SRBs' TVC.
First-stage guidance is active from SRB ignition through SRB separation plus four seconds. In this stage, guidance uses a preflight-planned (canned) table of roll, pitch and yaw attitudes referenced to relative velocity. There are canned tables defined for the cases of three good main engines, as well as for left, center or right main engine failed cases. In addition to sending commands to flight control to obtain proper vehicle attitude, guidance also sends commands to the MPS throttle in accordance with a preflight-defined throttle schedule, which is a function of relative velocity unless the flight crew has taken manual control of the throttle.
Navigation during first stage propagates the vehicle state vector through use of inertial measurement unit data and a gravity model. This function can be used to aid in driving the cathode ray tube's predictor.
Boost guidance is responsible for performing the ''table lookup'' of the appropriate roll, pitch and yaw command, depending on relative velocity. The table used is determined by the number of main engines thrusting. If one engine fails, guidance automatically recognizes the failure and lofts the trajectory, commanding the remaining two engines to a higher thrust percent, in the case of a contingency abort, for the remainder of first-stage ascent. For intact aborts, the throttle remains at the nominal maximum setting. Each main engine controller is used during the ascent phase to monitor the operation of an engine, issue inhibit commands to prevent a second engine from automatically shutting down when one has shut down, monitor the state of flight deck crew displays and switches from the switch processor and issue appropriate commands. The controllers also monitor guidance, navigation and control software for the proper time to check the main propulsion system liquid oxygen and liquid hydrogen low-level sensors, monitor GN&C software for proper time of main engine cutoff, and issue the command to close the main propulsion system liquid oxygen and liquid hydrogen prevalves after engine shutdown.
No GN&C-related crew actions are planned for first-stage ascent unless a failure occurs. To ensure that the auto flight control system is maintaining the expected ascent profile, the flight crew can verify that the vehicle is at the correct pitch attitude (via the attitude director indicator) and altitude rate (via the altitude/vertical velocity indicator) at each of five designated times during first-stage ascent. The flight crew can monitor that the main engines correctly throttle down and up. They can also ensure that the Pc - 50 message (chamber pressure greater than 50 psi) correctly appears on the major mode 102 (first stage) ascent trajectory CRT display before SRB separation and that SRB separation occurs on time. Manual intervention by the crew is required if these events are not automatically accomplished. The crew is also responsible for monitoring main engine performance. During first-stage ascent, only limited information is available to the crew on the PASS and backup flight system major mode 102 displays.
In the automatic mode, flight control during first-stage ascent uses commands sent to it from guidance. If the flight crew has selected control stick steering, commands are input through the rotational hand controller. Feedback from external sensors (rate gyro assemblies and accelerometers) is used to generate commands to reposition the SRB nozzles and main engine gimbals. Flight control sends the current vehicle attitude rates and the errors between current and desired attitudes to the ADI for display to the crew. As mentioned earlier, flight control also performs load relief by orienting the vehicle to reduce normal and lateral accelerations and by commanding the elevons to alleviate hinge moment loads in accordance with the premission-defined elevon schedule. If the crew does select CSS, the resultant control mode is discrete rate/attitude hold. In this mode, when the RHC is moved, a specified attitude rate is commanded as long as the stick is out of its neutral, detent position. When the stick is in detent, attitude hold commands are generated by flight control. Moreover, load relief is not performed by the software when the crew selects CSS.
The next major event is SRB separation, which occurs six seconds after the SRB separation sequence software detects both SRB chamber pressures below 50 psi within 4.3 seconds of each other and detects vehicle rates within specified limits. These checks ensure that neither SRB is burning at separation and that the SRBs will not recontact the ET after separation. SRB separation occurs at about two minutes after launch. At separation, the first stage is complete, and the software automatically shifts to major mode 103 (second stage).
SRB separation is normally performed automatically by the onboard GPCs; however, the flight crew can command separation through use of the SRB separation switches on panel C3. The SRB separation auto/man (manual) switch is positioned to man and the SRB sep push button depressed.
This manual function for SRB separation provides a backup for the automatic function; however, the manual function uses the same separation logic as the automatic. The automatic sequence is initiated by the software in the GPCs when the SRB chamber pressure is below 50 psi.
At SRB separation command, a three-axis attitude hold is commanded by the reaction control system for four seconds. When the SRB separation command is received, the SRB nozzles are positioned to null, and the flight control system is switched to the orbiter rate gyro assemblies. Four seconds after SRB separation, second-stage main engine guidance takes over. In addition, if a main engine shutdown is detected, the failed engine is positioned to null, and trim changes are commanded on the remaining engines.
Second-stage ascent begins at SRB separation and extends through main engine cutoff and external tank separation. The GN&C software is in major mode 103 (second stage) at this time, calculating the required main engine steering commands to achieve preflight-defined MECO conditions. After SRB separation, attitude hold is commanded until guidance converges.
The orbiter rate gyro assemblies are used by flight control as feedbacks to find errors that are used for stability augmentation during ascent, entry and aborts and for display on the commander's and pilot's ADIs. In second-stage ascent, the body-mounted accelerometers are not used, and the elevons are held in position.
Second-stage navigation is the same as that of first stage. Second-stage flight control continues through MECO.
Second-stage guidance uses a cyclic, closed-loop scheme to calculate the necessary commands to take the vehicle to a specified set of target MECO conditions. These conditions include cutoff velocity, radius from Earth, flight path angle, orbital inclination and longitude of the ascending node. The targeting scheme is called powered explicit guidance 1. Guidance also governs the main engine throttle command so that acceleration does not exceed 3 g's. The predicted time of MECO (TMECO) is calculated and displayed to the crew on the ascent trajectory display. Following SRB separation, it may take the PEG 1 guidance algorithm several cycles to converge and for TMECO to become stable. Forty seconds before MECO, guidance begins targeting only for the desired cutoff velocity, ignoring position constraints.
The main engines are throttled down at approximately seven minutes 40 seconds into the mission to maintain 3 g's for physiological and structural constraints. Approximately 10 seconds before MECO, the MECO sequence begins; about three seconds later the main engines are commanded to begin throttling at 10-percent thrust per second to 65-percent thrust. This is held for approximately 6.7 seconds, and the engines are shut down.
At MECO, the vehicle attitude commands (roll, pitch and yaw) are frozen, and body rate damping is maintained during the coast period by the reaction control system, which is accomplished by the transition digital autopilot. During this period, an automatic sequence is initiated by GN&C moding, sequencing and control that confirms that all main engines have shut down, the MECO confirmed flag is set, the MPS prevalves are closed and a flag is set for the external tank separation sequence after the MPS prevalves for all main engines have been commanded closed. The ET separation software sends the necessary commands to close the 17-inch orbiter/external tank feed line liquid oxygen and hydrogen disconnect valves, dead-faces the orbiter/ET interface, tests for the 17-inch feed line disconnect valves' closure, dead-faces the orbiter/ET interface, unlatches and retracts the 17-inch disconnects within the orbiter aft fuselage, arms the two aft and one forward orbiter/ET structural separation pyro initiator controllers, and fires the ET liquid oxygen tank tumble vent valve.
ET separation is performed automatically by the onboard general-purpose computers. If automatic ET separation is inhibited due to the various orbiter/ET separation tests, separation can occur only if an out-of-tolerance condition comes back within tolerance or if the flight crew elects to continue the separation by overruling the inhibit. The manual separation would be accomplished by positioning the ET separation switch on panel C3 to man and depressing the ET separation push button. Once automatic or manual ET separation is initiated, the orbiter/ET structural separation PICs are fired, separating the ET from the orbiter.
In the automatic orbiter/ET separation sequence, the transition DAP commands separation 18 seconds after main engine cutoff.
After ET separation, the two umbilical doors (one each for the 17-inch liquid oxygen and liquid hydrogen umbilical disconnects) are closed automatically for the entry phase. If the automatic function fails to close the umbilical doors, the flight crew can manually close them by using the ET umbilical door switches on panel R2.
After initiation of the orbiter/ET separation sequence, there is approximately 11 seconds of mated coast before the orbiter and external tank separate. The ET tumble system produces a tumble rate of 10 to 50 degrees per second after separation. In Kennedy Space Center launches, the external tank is on a suborbital trajectory that normally results in an impact location in the Indian Ocean. Except for direct-insertion launches from Kennedy Space Center. the tank impacts in the Pacific Ocean. External tank breakup nominally occurs during entry into the Earth's atmosphere at an altitude of approximately 185,000 feet.
Just before orbiter/ET separation, the reaction control system is inhibited. It is re-enabled immediately after ET separation to an inertial attitude hold. The transition DAP then commands the RCS to thrust the four forward and six aft negative RCS jets for a minus Z translation to achieve a 4-foot-per-second separation vertically, ensuring orbiter clearance from the arc of the rotating tank. When the required separation is achieved, the thrusting commands to the negative RCS jets are removed. The orbiter continues to coast away from the tank in the inertial attitude hold mode, gaining additional vertical clearance.
When the orbiter has gained the necessary separation, the orbiter/ET separation is flagged complete, and Mission Control is responsible for issuing a go/no-go for the impending orbital maneuvering system thrusting sequence. The software makes an automatic transition to major mode 104 (OMS-1 insertion) when the negative Z translation is complete.
During second-stage ascent, the flight crew monitors the onboard systems to ensure that the major GN&C events occur correctly and on time. These events include closed-loop guidance convergence, 3-g throttling, MECO, ET separation and the negative Z translation following ET separation. To monitor these events, the flight crew uses the dedicated displays-the main engine status lights on panel F7 and the PASS ascent trajectory and the BFS ascent trajectory 2 displays.
The crew can monitor guidance convergence by noting if the guidance-computed time of MECO is stabilized on the ascent trajectory display. If not, the crew takes manual control of the vehicle. They can also ensure that acceleration does not exceed 3 g's via the BFS ascent trajectory 2 display as well as the accel tape on the alpha Mach indicator. The crew can monitor MECO velocity on the BFS ascent trajectory 2 display as well as on the M/vel tape on the AMI. MECO is detected by the illumination of three red main engine status lights and by the main propulsion system chamber pressure meters on panel F7 going to zero.
Depending on mission requirements, the crew may be required to translate in the plus X direction, using the translational hand controller for 11 seconds, to allow the external tank camera to photograph the tank.
At specified points during second-stage ascent, the Mission Control Center will make voice calls to the crew indicating their status with respect to aborts. For example, the ''negative return'' call indicates that it is too late to select a return-to-launch-site abort.
Selection of an ascent abort mode may become necessary if there is a failure that affects vehicle performance, such as the failure of a main engine or an orbital maneuvering system failure. Other failures dictating early termination of a flight, such as a cabin leak, might require the selection of an abort mode.
There are two basic types of ascent abort modes for space shuttle missions: intact aborts and contingency aborts. An intact abort would provide a safe return of the orbiter to a planned landing site, while a contingency abort is designed to permit the crew to survive following more severe failures when an intact abort is not possible. A contingency abort would generally result in a ditch operation.
The ATO mode is designed to allow the vehicle to achieve a temporary orbit that is lower than the nominal orbit. This mode requires less performance and permits time to evaluate problems and to choose either an early deorbit burn or an OMS maneuver to raise the orbit and continue the mission.
The AOA would permit the vehicle to fly once around the Earth and make a normal entry and landing. This mode generally involves two OMS burns, with the second burn being a deorbit maneuver. There are two types of AOA trajectories: a normal AOA and a shallow AOA (which is considered only for contingency aborts). The entry trajectory for the normal AOA is similar to the nominal entry trajectory. The shallow AOA results in a flatter entry trajectory, which is less desirable than that of the normal AOA but uses less propellant in the OMS maneuvers. The shallow entry trajectory is less desirable because it exposes the vehicle to a longer period of atmospheric heating and to less predictable aerodynamic drag forces.
There is a definite order of preference for the various abort modes. The type of failure and the time of the failure would determine which type of abort is selected. In cases where performance loss is the only factor, the preferred modes, in order, would be ATO, AOA, TAL and RTLS. The mode chosen would be the highest one that could be completed with the remaining vehicle performance. For certain support system failures, such as cabin leaks or vehicle cooling problems, the preferred mode might be the one that would end the mission most quickly. In these cases, TAL or RTLS might be preferable to AOA or ATO. A contingency abort would never be chosen if another abort option existed.
The Mission Control Center is primarily responsible for calling aborts, since the controllers have more precise knowledge of the orbiter's state vector (through the use of sophisticated tracking equipment and ground computer resources) than the crew can obtain from the onboard navigation system. Before MECO, Mission Control periodically calls the crew to tell them which abort mode is (or is not) available. If ground communications are lost, the flight crew has onboard methods, such as cue cards, dedicated displays and GN&C CRT display information, from which to determine the current abort region.
The abort mode selected would depend on the cause and timing of the failure causing the abort and on which mode is safest or improves the chances for mission success. If the emergency is a main engine failure, the flight crew and MCC would select the best option available at the time of the failure. If the problem is a system failure that jeopardizes the vehicle, the abort mode that would result in the earliest vehicle landing would be chosen. RTLS and TAL would be the quickest options (35 minutes), whereas an AOA would require approximately 90 minutes. Which of these is selected would depend on the time of the failure with three good main engines.
The flight crew would select the abort mode by positioning the abort mode switch on panel F6 and depressing the abort push button on the panel. This switch is a rotary model with off , RTLS , TAL-AOA-S (shallow) and ATO positions.
The RTLS abort mode is designed to allow the return of the orbiter, crew and payload to the Kennedy Space Center launch site, approximately 25 minutes after lift-off. The RTLS profile would accommodate the loss of thrust from one main engine between lift-off and approximately four minutes and 20 seconds, at which time not enough MPS propellant remains to return to the launch site. An RTLS would also be used if a time-critical failure occurred that required the earliest possible landing, such as a cabin leak or major cooling system failure. In the last two situations, all three engines would still be burning, which leads to the term ''three-engine RTLS.''
An RTLS is composed of three stages: a powered stage, during which the main engines would still be burning; an ET separation phase; and the glide phase, during which the orbiter would glide to a landing at the Kennedy Space Center. The powered RTLS phase would begin with crew selection of the RTLS abort, after SRB separation. This is done by positioning the abort rotary switch on panel F6 to RTLS and depressing the abort push button. The time at which the RTLS is selected would depend on the reason for the abort. For example, a three-engine RTLS is selected at the latest possible time, approximately three minutes 45 seconds into the mission, whereas an RTLS due to an engine-out condition at lift-off would be chosen at the earliest time, approximately two minutes 20 seconds into the mission (after SRB separation). Upon selection of the RTLS, the ascent trajectory display would change to RTLS trajectory, and several portions of the display related to RTLS would become active. Also, the software major mode would change from 103 to 601 (RTLS second stage).
The vehicle would continue downrange to dissipate excess MPS propellant. The goal is to retain only enough propellant to turn the vehicle around, fly back toward the Kennedy Space Center and achieve MECO conditions so that the vehicle can glide to the Kennedy Space Center after ET separation. During the downrange phase, a pitch-around maneuver would be initiated (the time depends in part on the time of a main engine failure) to orient the orbiter/ET configuration to a head-up attitude, pointing toward the launch site. At this time, the vehicle would still be moving away from the launch site, but the main engines would be thrusting to null the downrange velocity. In addition, excess OMS and RCS propellant would be dumped by continuous OMS and RCS jet firings to improve the orbiter weight and center of gravity for the glide phase and landing. The guidance software would compute the necessary MPS throttle commands to reach the desired MECO point with less than 2 percent excess propellant remaining in the external tank. At MECO minus 20 seconds, a pitch-down maneuver (called powered pitch-down) would take the mated vehicle to the required ET separation attitude and pitch rate. After MECO is commanded, the tank separation sequence would begin, including negative Z translation after tank separation, ensuring that the orbiter did not recontact the tank and that the orbiter had achieved the necessary pitch attitude to begin the glide phase of the abort.
After the negative Z translation maneuver is completed, the software would go to major mode 602 (glide RTLS-1), and the glide phase of the abort would begin. First, guidance would send commands to stabilize the trajectory after tank separation and to adjust the orbiter's angle of attack for range control. From then on, the abort would be handled like a normal entry.
The TAL abort mode was developed to improve the options available in cases where a main engine fails after the last RTLS opportunity but before the first time that an AOA can be accomplished with only two main engines or a major orbiter system failure occurs after the last RTLS opportunity, making it imperative to land as quickly as possible.
In a TAL abort, the vehicle would continue on a ballistic trajectory across the Atlantic Ocean, landing on a predetermined runway approximately 45 minutes after launch. The landing site, which is located near the nominal ascent ground track of the orbiter in order to make the most efficient use of main engine propellant, must have the necessary runway length and U.S. State Department approval. Weather conditions must also be nominal. The three landing sites that have been identified for a due-east launch are Moron, Spain; Dakar, Senegal; and Ben Guerir, Morocco (on the west coast of Africa).
The onboard software is designed to accomplish a TAL automatically under most circumstances. To select the TAL abort mode, the crew must place the abort rotary switch in the TAL/AOA position on panel F6 and depress the abort push button before MECO. (Depressing it after MECO selects the AOA abort mode.) The title on the ascent trajectory display then would read TAL TRAJ. The TAL software would begin sending commands to steer the vehicle toward the plane of the landing site. It would also roll the vehicle head-up before MECO and send commands to begin an OMS propellant dump (by burning the propellants through the OMS engines and the RCS jets). This dump is necessary to increase vehicle performance (by decreasing weight), to place the center of gravity in the proper place for vehicle control and to reduce the vehicle's landing weight. The software also would target for a specified MECO configuration to ensure that the vehicle had enough energy to glide to the runway. The crew must mode the software to OPS 3 (entry) after external tank separation and set up the entry display for a TAL landing site. From that point on, the TAL would be handled like a nominal entry.
The ATO would be used to boost the orbiter to a safe orbital altitude if performance has been lost and it is impossible to reach the planned altitude. If a main engine fails in a region that results in a MECO underspeed, the Mission Control Center would determine that an abort mode is necessary and would inform the crew. In addition, the crew can verify the nominal OMS burn solutions on the OMS-1 maneuver execute display and burn them instead of the nominal targets. Similarly, they would load the ATO OMS-2 burn targets and use them for OMS-2. This results in the orbiter being placed in a circular orbit.
Another reason for an ATO is loss of OMS performance due to various failures, such as loss of two OMS engines, loss of one OMS propellant tank or loss of main bus A and B electrical power. In these cases, the OMS-1 burn would be delayed about 10 minutes, and the ATO OMS-1 targets would be used. This would result in an approximately 105- by 80-nautical-mile (120- by 92-statute-mile) orbit, which is considered safe for 24 hours. Thus, an OMS-2 burn would not be necessary immediately. The delayed ATO would be accomplished by loading the delayed ATO targets in the OMS-1 maneuver execute display and performing only the OMS-1 burn. If an underspeed existed at MECO so that an ATO or worse abort would be required, the OMS-1 burn must not be delayed. An underspeed would result in the apogee moving close to the post-MECO orbiter position. Delaying OMS-1 would cause the burn to be performed after apogee, which requires increased delta velocity and, thus, more propellant.
The AOA abort mode would be used when vehicle performance has been lost to such an extent that either it is impossible to achieve a viable orbit or not enough OMS propellant is available to accomplish the OMS-1, OMS-2 and deorbit burns. AOA would also be used in cases in which a major system problem (cabin leak, loss of cooling) made it necessary to land quickly. In this abort, one OMS burn would be made to adjust the post-MECO orbit so that a second OMS burn would cause the vehicle to deorbit and land at the AOA landing site (Northrup, Edwards Air Force Base or Kennedy Space Center.. Thus, in an AOA, the orbiter would circle the Earth once and land approximately 90 minutes after lift-off,
Several options are available to perform an AOA. Selection of the OMS-1 targets would be based on whether the abort were caused by a MECO underspeed, a system problem or an OMS/RCS performance problem. Selection of the OMS-2 targets would depend on whether a MECO underspeed existed and its magnitude. (The AOA OMS-2 burn is really a deorbit burn.) One set of targets would result in a steeper trajectory than would the other as the vehicle enters the atmosphere (entry interface); thus, this trajectory is referred to as a steep AOA. This is a more normal trajectory and stays well within the vehicle's thermal limits after it penetrates the atmosphere. It would require more delta velocity and consequently more propellant for the deorbit burn. Thus, if the MECO underspeed were severe or if both OMS helium tanks had failed, the shallow AOA targets would be used, resulting in a more shallow trajectory at entry interface and placing the vehicle closer to the skip-out boundary and its thermal limits.
The flight crew would determine that an AOA is required by Mission Control Center call and by checking the OMS-1 target solution against the OMS targeting cue card. Depending on the kind of AOA required, the crew would load the required OMS-1 targets and execute the burn. They would then position the software mode to OPS 3 and load the appropriate OMS-2 (deorbit) targets. After the burn is executed, the flight crew would fly to a landing at the preplanned site, much as they would for a nominal entry.
Contingency aborts are caused by loss of more than one main engine or failures in other systems. Loss of one main engine while another is stuck at a low thrust setting may also necessitate a contingency abort. Such an abort would maintain orbiter integrity for in-flight crew escape if a landing cannot be made at a suitable landing site.
Contingency aborts due to system failures other than those involving the main engines would normally result in an intact recovery of vehicle and crew. Loss of more than one main engine may, depending on engine failure times, result in a safe runway landing. However, in most three-engine-out cases during ascent, the orbiter would have to be ditched. The in-flight crew escape system would be used before ditching the orbiter.
This capability is provided during first- and second-stage ascent by substituting the inputs from the rotational hand controller for the automatic commands from guidance to gimbal the main engines and SRBs during first stage or the main engines during second stage. The digital autopilot remains active to process the flight crew's input. This is referred to in this mission phase as manual thrust vector control. MTVC is available at lift-off (SRB ignition command plus 0.365 second). One of the control stick steering push buttons at the glareshield/eyebrow panel F2 or F4 must be depressed before MTVC is available.
Once MTVC is activated, the vehicle is in a rate command/attitude hold mode in all axes. When the RHC is in detent, with MTVC selected, the vehicle is in attitude hold. The DAP holds the vehicle in the attitude it had when the RHC was in detent. A rate command equal to zero replaces the rate command generated in guidance and control steering. The attitude error for the DAP is computed by integrating the measured rates of the rate gyro assemblies. However, there are limits on vehicle rates and attitude errors. If the limits are exceeded when attitude hold is requested by placing the RHC in detent, attitude hold will not be initiated until the rates and errors are within limits.
When the RHC is removed from detent, a rate command proportional to the amount of deflection replaces the rate command previously generated. The attitude error is zeroed. The larger the deflection of the RHC, the larger the command. The flight control system compares these commands with inputs from the rate gyro assemblies and accelerometer assemblies (what the vehicle is actually doing-motion sensors) and generates control signals to produce the desired rates. When the commander or pilot releases the RHC, it returns to center and the vehicle maintains its current attitude (zero rates).
Following MECO, the orbiter's altitude and velocity will vary depending on the mission requirements. For example, an 80-nautical-mile (92-statute-mile) altitude with an inertial velocity of approximately 25,660 feet per second would place the orbiter in a suborbital trajectory so that the ET would enter following separation. In order to boost the orbiter to a viable orbit that does not degrade appreciably during the mission and satisfies mission objectives, two propulsive thrusting periods are made with the OMS engines, except in the case of a direct insertion, when only one OMS thrusting period is required to circularize the orbit. The first thrusting period is referred to as OMS-1 and boosts the orbiter to the desired apogee; the second burn is called OMS-2 and typically circularizes the orbit.
The optimal orbital altitude (the altitude that satisfies mission and payload goals) is determined before launch. During flight, however, problems, such as main engine and SRB performance loss and OMS propellant leaks or certain electrical power system failures, may prevent the vehicle from achieving the optimal orbit. In these cases, the OMS-1 and OMS-2 burns would be changed to compensate for the failure by selecting a delayed OMS-1, AOA or ATO abort option.
The main events that occur during the orbit insertion phase include execution of the OMS-1 thrusting period, typically about two minutes after MECO; an MPS propellant dump, which begins during OMS-1; positioning of the main engine nozzles for entry; shutdown of the three auxiliary power units; MPS power-down; and MPS vacuum inerting to ensure that all traces of MPS propellants are vented to space.
The major GN&C-related orbital tasks include achieving the proper position, velocity and attitude necessary to accomplish the mission objectives. To do this, the GN&C computer maintains an accurate state vector, targets and initiates maneuvers to specified attitudes and positions, and points a specified orbiter body vector at a target. These activities are planned with several constraints in mind, including fuel consumption, vehicle thermal limits, payload requirements and rendezvous/proximity operations considerations.
The GN&C software for the majority of on-orbit operations is called OPS 2 (on orbit), which is further divided into major mode 201 (orbit coast, in which the majority of attitude and pointing operations occur) and major mode 202 (maneuver execute, in which OMS translations are targeted and executed). GN&C software is also used for the flight control system checkout before deorbiting (OPS 8). In this configuration, the crew checks out the navigation aid systems, the dedicated displays, the RCS jets, the aerosurfaces and the hand controllers.
The navigation software available in OPS 2 has several important features. As before, it propagates the orbiter state vector. During coasting flight, the software uses a model of atmospheric drag acceleration to propagate the state vector. If translational thrusts are anticipated, the flight crew can set a flag for navigation to use IMU-sensed acceleration, when above a noise threshold value. When this flag has been set via an item entry on an orbit display, the flight crew can monitor the thrust that is sensed.
Another navigation option that may be available on orbit is called rendezvous navigation. When this option is enabled by a flight crew input on the relative navigation display, the software maintains a target state vector and the orbiter state vector. In this mode, it is possible for navigation to use external sensor data from the star tracker, crewman optical alignment sight or rendezvous radar (based on reasonableness tests) to compute the orbiter target state vector. This assumes that the target vector is accurate.
On orbit, the accuracy of the orbiter state vector depends on the accuracy of the IMUs and the accuracy of the modeled drag acceleration. Since there is currently no method on board the orbiter to compute independently corrections to the state vector, periodic updates are sent from Mission Control to correct any errors that develop with the onboard state vector.
Another feature available in navigation in OPS 2 is the landing site update function, which allows the crew to select different runways to be used in the entry guidance computations. The crew interfaces with this capability through the universal pointing display only in major mode 201.
One of on-orbit software's several features, universal pointing is used to compute attitude changes required to point a specified orbiter body axis at a specified target, to maneuver to a predetermined attitude, to rotate the orbiter about a specified body axis or to maintain attitude hold. Although the complete capabilities of the universal pointing software are available only in major mode 201, a subset is available in major mode 202 and OPS 8 (on orbit).
Another guidance feature is PEG 7, or external delta-velocity, targeting for OMS or RCS burns. This targeting scheme is identical to that used in OPS 1 (ascent) and OPS 3 (entry). In this mode, guidance sends the commands to flight control to execute a burn specified by an ignition time and delta velocities in the local vertical coordinate system at ignition. Commands continue to be generated until the original delta-velocity requirement is met. This option is available in major mode 202 via the orbit maneuver execute display.
The third guidance capability is an on-orbit targeting scheme that is used to compute the parameters required to move the orbiter to a specified target offset position in a given amount of time. This feature, which is used to do onboard targeting of rendezvous maneuvers, is enabled via the orbit targeting CRT display. The actual thrusting period is still accomplished via the orbit maneuver execute display.
The orbit flight control software includes an RCS DAP, an OMS thrust vector control DAP, a module called an attitude processor to calculate vehicle attitude and logic to govern which DAP is selected. The attitudes calculated by the attitude processor are displayed on the ADI along with another crew display, universal pointing, which is available in major mode 201. The vehicle attitude is used by the DAP to determine attitude and rate errors.
The RCS DAP, used in OPS 2 at all times except when an OMS burn is in progress, controls vehicle attitudes and rates using 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. That selection depends on fuel-consumption considerations and how quickly the vehicle must be maneuvered to satisfy a mission objective.
The rotation rates and dead bands, translation rate and certain other options for the DAP may be changed by the crew during the orbit phase using the DAP configuration display. The crew can load the DAP with these options two ways at a time. One set is accessed by depressing the DAP A push button on the orbital DAP panel and the other by depressing the DAP B push button. For convenience, each planned DAP configuration is given a number and is referenced to that number and to the DAP used to access it. Typically, the DAP A configurations 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 more precision in executing maneuvers or in holding attitude.
The RCS DAP has both an automatic and a manual rotation mode. The one that is used depends on crew selection of the auto or man push buttons 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 DAP B and the information loaded in the DAP configuration display. Moreover, in automatic, the DAP determines the required attitude to be achieved from universal pointing. It 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 stated, in pulse, a single burst of jet fire is produced with each RHC deflection. The resultant rotational rate is specified on the DAP configuration display. In discrete rate, jet firings continue to be made as long as the RHC is out of detent to maintain the rotational rate specified on the DAP configuration display. In acceleration, 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 the firing of the RCS jets, depending on which transition DAP mode push button is selected on the orbital DAP panel. In pulse, a single burst of jet fire results; in normal, there are continuous jet firings with a specified subset of the available jets; in high, all upfiring jets fire continuously in a Z translation; and in low, a special technique is used to perform a Z translation with the forward- and aft-firing RCS jets in order not to fire directly toward a target (this avoids 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 via the orbit maneuver execute display. The TVC DAP uses the guidance-generated delta-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 because of OMS engine failure, RCS commands required 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.
During OPS 8 major mode 801, the orbital checkout of orbiter systems used during entry is performed, usually the day before deorbit. These activities take about 15 minutes. System checkout is performed in two parts. The first part, which requires the use of one auxiliary power unit/hydraulic system, involves the repositioning of the left and right main engine nozzles for entry and the cycling of aerosurfaces, hydraulic motors and hydraulic switching valves. After the checkout is completed, the auxiliary power unit is deactivated. The second part consists of a check of all the crew dedicated displays; self-test of the microwave scan beam landing system, TACAN, accelerometer assemblies, radar altimeter, rate gyro assemblies and air data transducer assemblies; and a check of the hand controllers, rudder pedal transducer assemblies, speed brake, panel trim switches, RHC trim switches, speed brake takeover push button and mode/sequence push button light indicators.
Deorbit guidance, navigation and flight control software operates through the transition DAP to provide maneuvering of the spacecraft to the OMS deorbit ignition attitude, OMS thrusting commands, OMS engine gimbaling for thrust vector control and RCS thrusting commands, in conjunction with use of the DAP similar to that for orbit insertion.
In returning home, the orbiter must be sufficiently decelerated by an OMS retrograde burn that when it enters the atmosphere, it maintains control and glides to the landing site. For the nominal end of mission, a retrofiring of approximately 2.5 minutes is performed at the appropriate point in the vehicle's trajectory. For this maneuver, the orbiter is positioned in a tail-first thrusting attitude. Deorbit thrusting is nominally accomplished with the two OMS engines and must establish the proper entry velocity and range conditions. It is possible to downmode to one OMS engine (with RCS roll control) or, in the event that both OMS engines malfunction, to plus X aft RCS jets.
Approximately four hours before deorbit, the environmental control and life support system radiator bypass/flash evaporator system is checked out, since the flash evaporator is used to cool the Freon-21 coolant loops when the ECLSS radiators are deactivated and the payload bay doors are closed. The high-load evaporator cools the coolant loops until the ECLSS ammonia boilers are activated by the GPCs at an altitude of some 140,000 feet. The orbiter IMUs are aligned, the star trackers are deactivated and the star tracker doors are closed.
About one hour before deorbit, the crew members take their seats. The spacecraft is then manually maneuvered using the RCS jets to the deorbit attitude (retrograde). About 30 minutes before deorbit, the OMS is prepared for deorbit thrusting. This consists of OMS thrust vector control gimbal checks, OMS data checks, orbiter vent door closure and single auxiliary power unit start. At the completion of the single OMS deorbit burn, the crew manually maneuvers the spacecraft to the required entry attitude (nose first) using the RCS jets. The propellants remaining in the forward RCS are dumped through the forward RCS engines, if required, and the two remaining APUs are started and remain operating through entry and landing rollout. Thermal conditioning of the spacecraft's hydraulic fluid system is also begun, if required.
The deorbit phase of the mission includes the deorbit burn preparations, including the loading of burn targets and maneuvering to burn attitude; the execution and monitoring of the burn; reconfiguration after the burn; and a coast mode until the atmosphere (and dynamic pressure buildup) is reached at approximately 400,000 feet. This is called the entry interface.
The deorbit and entry flight software is called OPS 3. Major mode 301 is a deorbit coast mode in which deorbit targets can be loaded, although the burn cannot be executed in this mode. This mode is necessary to execution of the burn. After the burn, a software transition is made to another coast mode, major mode 303, which is used to prepare for penetration into the atmosphere.
During the deorbit phase, navigation again propagates the orbiter state vector based upon a drag model or upon inertial measurement unit data if sensed vehicle accelerations are above a specified threshold. During OPS 3, navigation maintains and propagates three orbiter state vectors, each based on a different IMU. From these three state vectors, a single orbiter state vector is calculated using a mid value selection process and is passed on for use by guidance, flight control, dedicated display and CRT display software. Three separate state vectors are propagated to protect the onboard software from problems resulting from two IMU data failures. In such a case, once the bad IMU is detected and deselected, the state vector associated with the remaining good IMU will not have been polluted. This three-state vector system is used only during OPS 3 since this phase is most critical with respect to navigation errors and their effects on vehicle control and an accurate landing.
Another feature available during this phase is the software's computation of a statistical estimate of the error in the state vector propagation, which is used later in flight when external sensor data are available. Also, in this phase, it is possible for the crew or the Mission Control Center to input a delta state vector to correct navigation.
Guidance during deorbit is similar to that used in the orbit insertion phase. The PEG 4 scheme is used to target the deorbit burn and guide the vehicle during the burn, although the required conditions are different. The deorbit burn targets are for the proper conditions for entry interface, including altitude, position with respect to the Earth and thus the landing site, and satisfaction of certain velocity/flight path angle constraints. Together these ensure that the vehicle can glide to the landing site within thermal limits. Deorbit burn targets are specified before flight for a nominal mission, but it is possible for the ground to uplink changes or for the flight crew to recompute them using an onboard hand-held calculator program. It is also possible to specify that OMS fuel be wasted during the burn (burned out of plane) to establish an acceptable orbiter center of gravity for entry.
The crew is responsible for loading these targets on the deorbit maneuver execute display. Guidance then computes the necessary vehicle attitude to be established before the burn and displays it to the crew. As in OPS 1, it is possible to load an external delta-velocity (PEG 7) target, but this option is not normally used.
Flight control during the deorbit phase is similar to that used during orbit insertion-i.e., the transition DAP is once again in effect.
The flight crew interfaces with the guidance, navigation and flight control software during the deorbit phase via CRT display inputs, RHC/THC maneuvers and ADI monitoring. The major CRT display used is the deorbit maneuver execute. In major modes 301 and 303, the display is deorbit maneuver coast; whereas in major mode 302, it is deorbit maneuver execute. This display is identical in format to the OMS-1 maneuver execute and orbit maneuver execute displays, although there are some differences in its capabilities between OPS 1, 2 and 3. It is used to set up and target the OMS burn, to specify fuel to be wasted during the burn, to display the required burn attitude, to initiate an automatic maneuver to that attitude and to monitor the progress of the burn.
Another CRT display available during the deorbit phase is the horizontal situation. During deorbit preparation, the crew may verify that the display is ready for use during entry (correct runway selection, altimeter setting, etc.), but its other capabilities are not utilized until after entry interface.
The flight crew's task during this phase includes entering the correct burn targets in the deorbit maneuver execute display and maneuvering to burn attitude, either automatically or manually using the RHC. The burn itself is typically executed in auto, and the flight crew's task is to monitor the burn's progress in terms of velocities gained and OMS performance.
In cases of OMS failures (engine, propellant tank, data path), the flight crew must be prepared to reconfigure the system to ensure that the burn can safely continue to completion, that sufficient RCS propellant remains for entry and that the orbiter center of gravity stays within limits.
The entry phase of flight begins approximately five minutes before entry interface, which occurs at an altitude of 400,000 feet. At EI minus five minutes, the orbiter is at an altitude of about 557,000 feet, traveling at 25,400 feet per second, and is approximately 4,400 nautical miles (5,063 statute miles) from the landing site. The goal of guidance, navigation and flight control software is to guide and control the orbiter from this state (in which aerodynamic forces are not yet felt) through the atmosphere to a precise landing on the designated runway. All of this must be accomplished without exceeding the thermal or structural limits of the orbiter.
The entry phase is divided into three separate phases because of the unique software requirements. Entry extends from EI minus five minutes to terminal area energy management interface at an altitude of approximately 83,000 feet, at a velocity of 2,500 feet per second, 52 nautical miles (59 statute miles) from the runway and within a few degrees of tangency with the nearest heading alignment cylinder in major mode 304.
TAEM extends to the approach and landing capture zone, defined as the point when the orbiter is on glide slope, on airspeed, and on runway centerline, which occurs below 10,000 feet and is the first part of major mode 305. The orbiter attains subsonic velocity at an altitude of approximately 49,000 feet about 22 nautical miles (25 statute miles) from the runway.
Approach and landing begins at the approach and landing capture zone, an altitude of 10,000 feet and Mach 0.9 and extends through the receipt of the weight-on-nose-gear signal after touchdown, which completes major mode 305.
At 400,000 feet, a pre-entry phase begins in which the orbiter is maneuvered to zero degrees roll and yaw (wings level) and a predetermined angle of attack for entry. The flight control system issues the commands to the roll, yaw and pitch RCS jets for rate damping in attitude hold for entry into the Earth's atmosphere until 0.176 g is sensed, which corresponds to a dynamic pressure of 10 pounds per square foot, approximately the point at which the aerosurfaces become active.
When the orbiter is in atmospheric flight, it is flown by varying the forces it generates while moving through the atmosphere, like any other aerodynamic vehicle. The forces are determined primarily by the speed and direction of the relative wind (the airstream as seen from the vehicle). The direction of the airstream is described by the difference between the direction that the vehicle is pointing (attitude) and the direction that it is moving (velocity). It may be broken into two components: angle of attack (vertical component) and sideslip angle (horizontal component).
To rotate the orbiter in the atmosphere, aerodynamic control surfaces are deflected into the airstream. The orbiter has seven aerodynamic control surfaces. Four of these are on the trailing edge of the wing (two per wing). They are called elevons because they combine the effects of elevators and ailerons on ordinary airplanes. Deflecting the elevons up or down causes the vehicle to pitch up or down. If the right elevons are deflected up and the left elevons are deflected down, the orbiter will roll to the right-that is, the right wing falls and the left wing rises. The fifth control surface is the body flap, located on the rear lower portion of the aft fuselage. It provides thermal protection for the three main engines during entry, and during atmospheric flight it provides pitch trim to reduce elevon deflections. The sixth and seventh control surfaces are the rudder/speed brake panels, located on the aft portion of the vertical stabilizer. When both panels are deflected right or left, the spacecraft will yaw, moving the spacecraft's nose right or left, thus acting as a rudder. If the panels are opened at the trailing edge, aerodynamic drag force will increase, and the spacecraft will slow down. Thus, the open panels are called a speed brake.
On the flight deck display and control panel (panel F7 between the commander and pilot) are the surface position indicators, which display the position of each aerodynamic control surface.
The aft RCS jets maneuver the spacecraft until a dynamic pressure of 10 pounds per square foot is sensed; at this point, the orbiter's ailerons become effective, and the aft RCS roll jets are deactivated. At a dynamic pressure of 20 pounds per square foot, the orbiter's elevators become effective, and the aft RCS pitch jets are deactivated. The orbiter's speed brake is used below Mach 10 to induce a more positive downward elevator trim deflection. At Mach 3.5, the rudder become activated, and the aft RCS yaw jets are deactivated (approximately 45,000 feet).
Entry flight control is maintained with the aerojet DAP, which generates effector and RCS jet commands to control and stabilize the vehicle during its descent from orbit. The aerojet DAP is a three-axis rate command feedback control system that uses commands from guidance in automatic or from the flight crew's RHC in control stick steering. Depending on the type of command and the flight phase, these result in fire commands to the RCS or deflection commands to the aerosurfaces.
In the automatic mode, the orbiter is essentially a missile, and the flight crew monitors the instruments to verify that the vehicle is following the correct trajectory. The onboard computers execute the flight control laws (equations). If the vehicle diverges from the trajectory, the crew can take over at any time by switching to CSS. The orbiter can fly to a landing in the automatic mode (only landing gear extension and braking action on the runway are required by the flight crew). The autoland mode capability of the orbiter is used by the crew usually to a predetermined point in flying around the heading alignment cylinder. In flights to date, the crew has switched to CSS when the orbiter is subsonic. However, autoland provides information to the crew displays during the landing sequence.
The commander and pilot can select automatic or CSS flight control modes. The crew can select separate modes for pitch and roll and yaw (roll and yaw must be in the same mode). The body flap and speed brake have automatic and manual modes.
Automatic pitch provides automatic control in the pitch axis, and the automatic roll and yaw provides automatic control in the roll and yaw axes. During entry, the automatic mode uses the RCS jets until dynamic pressure permits the aerosurfaces to become effective; the aft RCS jets and spacecraft aerosurfaces are then used together until dynamic pressure becomes sufficient for aerosurface control only.
Control in the pitch axis is provided by the elevons, speed brake and body flap. The elevons provide control to guidance normal acceleration commands, control of pitch rate during slap-down (landing) for nose wheel load protection, and static load relief after slap-down for main landing gear wheel and tire load protection. The speed brake provides control to guidance surface deflection (open/close, increase/decrease velocity) command. The body flap provides control to null elevon deflection.
Control in the roll and yaw axes is provided by the elevons and rudder. The elevons provide control to guidance bank angle command during terminal area energy management and autoland and control to guidance wings-level command during flat turns, 5 feet above touchdown. The rudder provides yaw stabilization during TAEM and autoland and control to guidance yaw rate command during flat turn and subsequent phases.
When the orbiter is in the automatic pitch and roll and yaw modes, the crew's manual control stick steering commands are inhibited. In the CSS mode, the crew flies the orbiter by deflecting the RHC and rudder pedals. The flight control system interprets the RHC motions as rate commands in pitch, roll or yaw and controls the RCS jets and aerosurfaces. The larger the deflection, the larger the command. The flight control system compares these commands with inputs from rate gyros and accelerometers (what the vehicle is actually doing-motion sensors) and generates control signals to produce the desired rates. If the crew releases the RHC, it will return to center, and the orbiter will maintain its present attitude (zero rates). The rudder pedals position the rudder during atmospheric flight; however, in actual use, because flight control software performs automatic turn coordination, the rudder pedals are not used until the wings are leveled before touchdown.
The CSS mode is similar to the automatic mode except that the crew can issue three-axis commands, affecting the spacecraft's motion. These are augmented by the feedback from the same spacecraft motion sensors, except for the normal acceleration (velocity) accelerometer assemblies, to enhance control response and stability.
The commander's or pilot's RHC commands are processed by the GPCs in the CSS mode together with data from the motion sensors. The flight control module processes the flight control laws and provides commands to the flight control system, which positions the aerosurfaces in atmospheric flight.
Control in the roll and yaw axes is provided by the elevon and the rudder. The elevons augment the RHC control. The rudder interface between the roll and yaw channel automatically positions the rudder for coordinated turns. A rudder pedal transducer assembly is provided at the commander and pilot stations. The two rudder pedal assemblies are connected to their respective RPTAs. Because of the roll and yaw interface, rudder pedal use should not be required until just before touchdown. There is an artificial feel in the rudder pedal assemblies. The RPTA commands are processed by the GPCs, and the flight control module commands the flight control system to position the rudder.
In the CSS mode, the commander's and pilot's RHC trim switches, in conjunction with the trim enable/inhibit switch, activate or inhibit the RHC trim switch. When the RHC trim switch is positioned forward or aft, it adds a trim rate to the RHC pitch command; positioning it left or right adds a roll trim.
Manual control (CSS mode) in the pitch axis is provided by the elevons, speed brake and body flap. The elevons provide augmented control through the RHC pitch command. The speed brake can be switched to its manual mode at either the commander's or pilot's station by depressing a takeover switch on the speed brake/thrust controller handle. Manual speed brake control can be transferred from one station to the other by activating the takeover switch. When the SBTC is at its forward setting, the speed brake is closed. Rotating the handle aft, positions the speed brake at the desired position (open) and holds it. To regain automatic speed brake control, the push button must be depressed again. In the manual mode, speed brake commands are processed by the GPCs, and the flight control module commands the flight control system to position the speed brake and hold it at the desired position. The body flap can be switched to its manual mode at panel C3 by moving a toggle switch from auto/off to up or down for the desired body flap position. These are momentary switch positions; when released, the switch returns to off .
In the entry phase, the RCS commands roll, pitch and yaw. Lights on the commander's panel F6 are used to indicate the presence of an RCS command from the flight control system to the RCS jet selection logic; however, this does not indicate an actual RCS jet thrusting command. The minimum light-on duration is extended to allow the light to be seen even for minimum-impulse RCS jet thrusting commands. After the roll and pitch aft RCS jets are deactivated, the roll indicator lights are used to show that three or more yaw RCS jets have been requested. The pitch indicator lights are used to show elevon rate saturation.
At approximately 265,000 feet, the spacecraft enters a communications blackout, which lasts until the orbiter reaches an altitude of approximately 162,000 feet. Between these altitudes, heat is generated as the spacecraft enters the atmosphere, ionizing atoms of air that form a layer of ionized gas particles around the spacecraft. Radio signals between the spacecraft and the ground cannot penetrate this sheath of ionized particles, and radio communications are blocked for approximately 16 minutes.
During the entry subphase, the primary objective is to dissipate the tremendous amount of energy that the orbiter possesses when it enters the atmosphere so that it does not burn up (entry angle too steep) or skip out of the atmosphere (entry angle too shallow), stays within structural limits, and arrives at the TAEM interface with the altitude and range to the runway necessary for a landing. This is accomplished by adjusting the orbiter's drag acceleration on its surface using bank commands relative to vehicle velocity. During TAEM, as the name implies, the goal is to manage the orbiter's energy while the orbiter travels along the heading alignment cylinder, which lines up the vehicle on the runway centerline. A HAC is an imaginary cone that, when projected on the Earth, lies tangent to the extended runway centerline.
Guidance performs different tasks during the entry, TAEM and approach and landing subphases. During the entry subphase, guidance attempts to keep the orbiter on a trajectory that provides protection against overheating, overdynamic pressure and excessive normal acceleration limits. To do this, it sends commands to flight control to guide the orbiter through a tight corridor limited on one side by altitude and velocity requirements for ranging (in order to make the runway) and orbiter control and on the other side by thermal constraints. Ranging is accomplished by adjusting drag acceleration to velocity so that the orbiter stays in that corridor. Drag acceleration can be adjusted primarily in two ways: by modifying the angle of attack, which changes the orbiter's cross-sectional area with respect to the airstream, or by adjusting the orbiter's bank angle, which affects lift and thus the orbiter's sink rate into denser atmosphere, which in turn affects drag. Using angle of attack as the primary means of controlling drag results in faster energy dissipation with a steeper trajectory but violates the thermal constraint on the orbiter's surfaces. For this reason, the orbiter's bank angle (roll control) is used as the primary method of controlling drag, and thus ranging, during this phase. Increasing the roll angle decreases the vertical component of lift, causing a higher sink rate. Increasing the roll rate raises the surface temperature of the orbiter, but not nearly as drastically as does an equal angle of attack command. The orbiter's angle of attack is kept at a high value (40 degrees) during most of this phase to protect the upper surfaces from extreme heat. It is modulated at certain times to ''tweak'' the system and is ramped down to a new value at the end of this phase for orbiter controllability. Using bank angle to adjust drag acceleration causes the orbiter to turn off course. Therefore, at times, the orbiter must be rolled back toward the runway. This is called a roll reversal and is commanded as a function of azimuth error from the runway. The ground track during this phase, then, results in a series of S-turns.
If the orbiter is low on energy (the current range-to-go is much greater than nominal at current velocity), entry guidance will command lower-than-nominal drag levels. If the orbiter has too much energy (the current range-to-go is much less than nominal at current velocity), entry guidance will command higher-than-nominal drag levels to dissipate the extra energy.
Roll angle is used to control cross range. Azimuth error is the angle between the plane containing the orbiter's position vector and the heading alignment cylinder tangency point and the plane containing the orbiter's position vector and velocity vector. When the azimuth error exceeds an initialized-loaded number, the orbiter's roll angle is reversed.
Thus, descent rate and downranging are controlled by bank angles-the steeper the bank angle, the greater the descent rate and the greater the drag. Conversely, the minimum-drag altitude is wings level. Cross range is controlled by bank reversals.
The entry thermal control phase is designed to keep the thermal protection system's bond line within design limits. A constant heating rate is maintained until the velocity is below 19,000 feet per second.
In the equilibrium glide phase, the orbiter effects a transition from the rapidly increasing drag levels of the temperature control phase to the constant drag level of the constant drag phase. Equilibrium glide is defined as flight in which the flight path angle, the angle between the local horizontal and the local velocity vector, remains constant. This flight regime provides the maximum downrange capability. It lasts until drag acceleration reaches 33 feet per second squared.
The constant drag phase begins at 33 feet per second squared. Angle of attack is initially 40 degrees, but it begins to ramp down until it reaches approximately 36 degrees by the end of this phase.
The transition phase is entered as the angle of attack continues to ramp down, reaching about 14 degrees at TAEM interface, with the vehicle at an altitude of some 83,000 feet, traveling 2,500 feet per second (Mach 2.5), and 52 nautical miles (59 statute miles) from the runway. At this point, control is transferred to TAEM guidance.
TAEM guidance steers the orbiter to the nearest of two heading alignment cylinders, whose radii are approximately 18,000 feet and whose locations are tangent to and on either side of the runway centerline on the approach end. Normally, the software is set to fly the orbiter around the HAC on the opposite side of the extended runway centerline. This is called the overhead approach. If the orbiter is low on energy, it can be flagged to acquire the HAC on the same side of the runway. This is called the straight-in approach. In TAEM guidance, excess energy is dissipated by an S-turn, and the speed brake can be used to modify drag, lift-to-drag ratio and the flight path angle under high-energy conditions. This increases the ground track range as the orbiter turns away from the nearest HAC until sufficient energy is dissipated to allow a normal approach and landing guidance phase capture, which begins at 10,000 feet at the nominal entry point. The orbiter can also be flown near the velocity for maximum lift over drag or wings level for the range stretch case, which moves the approach and landing guidance phase to the minimum entry point.
At TAEM acquisition, the orbiter is turned until it is aimed at a point tangent to the nearest HAC and continues until it reaches way point 1. At way point 1, the TAEM heading alignment phase begins, in which the HAC is followed until landing runway alignment, plus or minus 20 degrees, is achieved. As the orbiter comes around the HAC, it should be lined up on the runway and at the proper flight path angle and airspeed to begin the steep glide slope to the runway.
In the TAEM prefinal phase, the orbiter leaves the HAC, pitches down to acquire the steep glide slope, increases airspeed and banks to acquire the runway centerline, continuing until it is on the runway centerline, on the outer glide slope and on airspeed.
The approach and landing guidance phase begins with the completion of the TAEM prefinal phase and ends when the orbiter comes to a complete stop on the runway. The approach and landing interface airspeed requirement at an altitude of 10,000 feet is approximately 290 knots, plus or minus 12 knots, equivalent airspeed, 6.9 nautical miles (7.9 statute miles) from touchdown.
Autoland guidance is initiated at this point to guide the orbiter to the minus 19- to 17-degree glide slope (which is more than seven times that of a commercial airliner's approach) aimed at a target approximately 0.86 nautical mile (1 statute mile) in front of the runway. The descent rate in the latter portion of TAEM and approach and landing is greater than 10,000 feet per minute (approximately 20 times higher than a commercial airliner's standard 3-degree instrument approach angle). The steep glide slope is tracked in azimuth and elevation, and the speed brake is positioned as required.
Approximately 1,750 feet above the ground, guidance sends commands to keep the orbiter tracking the runway centerline, and a preflare maneuver is started to position the orbiter on a shallow 1.5-degree glide slope in preparation for landing, with the speed brake positioned as required. At this point, the crew deploys the landing gear.
Final flare is begun at approximately 80 feet to reduce the sink rate of the vehicle to less than 9 feet per second. After the spacecraft crosses the runway threshold-way point 2 in the autoland mode-navigation uses the radar altimeter vertical component of position in the state vector for guidance and navigation computations from an altitude of 100 feet to touchdown. Touchdown occurs approximately 2,500 feet past the runway threshold at a speed of 184 to 196 knots (211 to 225 mph). As the airspeed drops below 165 knots (189 mph), the orbiter begins derotation in preparation for nose gear slap-down.
The navigation system used from entry to landing consists of the IMUs and navigation aids (TACAN, air data system, microwave scan beam landing system and radar altimeter). The three IMUs maintain an inertial reference and provide delta velocities until MSBLS is acquired.
Navigation-derived air data-obtained after deployment of the two air data probes at approximately Mach 3-is needed from entry through landing as inputs to the guidance, flight control and crew display. TACAN provides range and bearing measurements and is available at approximately 145,000 feet, nominally accepting the data into the state vector before 130,000 feet. It is used until MSBLS acquisition, which provides range, azimuth and elevation commencing at approximately 18,000 feet. Radar altimeter data are available at approximately 9,000 feet.
TACAN acquisition and operation are completely automatic, but the crew has the necessary controls and displays to evaluate TACAN system performance and to take over if required. When the distance to the landing site is approximately 120 nautical miles (138 statute miles), TACAN begins interrogating six navigation region stations. As the spacecraft proceeds, the distances to the remaining stations and to the next-nearest station are computed, and the next-nearest station is selected automatically if the spacecraft is closer to it than it is to the previous locked-on station. Only one station is interrogated if the distance to the landing site is less than approximately 20 nautical miles (23 statute miles). Again, TACAN automatically switches from the last locked-on navigation region station to begin searching for the landing site station. TACAN azimuth and range are provided on the CRT displaying the horizontal situation. TACAN range and bearing cannot be used to produce a good estimate of the altitude position component, so navigation uses barometric altitude derived from the air data system probes.
MSBLS acquisition and operation are also completely automatic, and the flight crew can evaluate system performance and take over if necessary. MSBLS acquisition occurs at approximately 18,000 feet and about 8 nautical miles (9.2 statute miles) from the runway. The range and azimuth measurements are provided by a ground antenna located at the end of the runway and to the left of the runway centerline. Elevation measurements are given by a ground antenna to the left of the runway centerline, about 2,624 feet from the runway threshold.
During entry, the commander's and pilot's altitude director indicators become two-axis balls displaying body roll and pitch attitudes with respect to local vertical/local horizontal. These are generated in the attitude processor from IMU data. The roll and pitch error needles each display the body roll and pitch attitude error with respect to entry guidance commands by using the bank guidance error and the angle of attack error generated from the accelerometer assemblies. In atmospheric flight, the roll attitude error and the normal acceleration error are displayed by the roll and pitch error needles, respectively. The sideslip angle is displayed on the yaw error needle. The roll and pitch rate needles display stability roll and body rates by using stability roll rate, rate gyro rate and pitch rate. The yaw rate needle displays stability yaw rate. After main landing gear touchdown, the yaw error with respect to runway centerline and nose gear slap-down pitch rate error are displayed on the roll and pitch error needles. During rollout, the pitch error indicator indicates pitch error rate.
During entry, the commander's and pilot's horizontal situation indicators display a pictorial view of the spacecraft's location with respect to various navigation points. The navigation attitude processor provides the inputs to the HSI until the communications blackout is passed, at approximately 145,000 feet. TACAN is then acquired and accepted for HSI inputs at about 130,000 feet until MSBLS acquisition at approximately 18,000 feet some 8 nautical miles (9.2 statute miles) from the runway.
When the approach mode and MSBLS source are selected for the commander's and pilot's HSI, data from the MSBLS replaces TACAN data. MSBLS azimuth, elevation and range are used from acquisition until the runway threshold is reached, and azimuth and range are used to control rollout.
At an altitude of 9,000 feet, radar altimeter 1 or 2 can be selected to measure the nearest terrain within the beamwidth of the altimeters. This indication is given to the altitude/vertical velocity indicator radar, altitude and meter display from 5,000 feet to landing.
The left and right air data system probes are deployed by the flight crew at about Mach 3. This system senses air pressures related to orbiter movement through the atmosphere for updating the navigation state vector in altitude, guidance in steering and speed brake command calculations, flight control for control law computations, and for display on the alpha Mach indicators and altitude/vertical velocity indicators.
The AMIs display essential flight parameters relative to the spacecraft's travel in the air mass, such as angle of attack, acceleration, velocity and knots of equivalent airspeed. The source of data for the AMIs is determined by the position of the air data select switch. Before the deployment of the air data system probe, the AMIs receive inputs from the navigation attitude processor. When the air data probes are deployed, the left or right air data system provides the inputs to all AMIs except the acceleration indicator, which remains on the navigation attitude processor, and the radar altitude. Neither is operational until the orbiter descends to 5,000 feet.
The three rate gyro assemblies of the flight control system measure and supply output data proportional to the orbiter's attitude rates about its three body axes, while the three accelerometer assemblies measure and supply output data proportional to the orbiter's normal (vertical) and lateral (right and left) accelerations. These assemblies are incorporated into the flight control system for augmenting stability because of the orbiter's marginal stability in its pitch and yaw axes at subsonic speeds.
The three IMUs constitute an all-attitude stabilized platform that also measures and supplies output data proportional to the spacecraft's attitude (rotation) and acceleration (velocity). They augment the rate gyro assemblies and accelerometer assemblies.
The rate gyro assembly pitch rate (rotation) and the accelerometer assembly normal acceleration (velocity) are used to generate elevon (elevator) deflection commands. The rate gyro assembly yaw rate (rotation) and the accelerometer assembly lateral acceleration generate the rudder deflection required for directional stability. The rate gyro assembly roll rate (rotation) generates the elevon (aileron) deflection command required for lateral (roll) stability. The speed brake and body flap positions generate the elevon deflection required for trim near neutral to maximize roll effectiveness of the elevons.
In the entry phase, navigation software functions as it did during the deorbit phase (three state vectors corresponding to each IMU) except that additional external sensor data are sequentially incorporated. These data provide the accuracy necessary to bring the orbiter to a pinpoint landing and, to some extent, to maintain vehicle control. The TACAN system, which becomes available at about 156,000 feet, provides slant range and magnetic bearing to various fixed stations around the landing site. It is used until the orbiter is approximately 1,500 feet above the ground, at which point it is rendered ineffective by ground reflection. The air data system, which includes two transducer assemblies attached to a probe on the left side of the vehicle and two on the right side, provides pressures from which angle of attack, Mach number, equivalent airspeed, true airspeed, dynamic pressure, barometric altitude and altitude rate are computed. Only barometric altitude is used by navigation. The other parameters are used by guidance and flight control as well as for display to the flight crew. The probes are normally deployed around Mach 3. The MSBLS precisely determines slant range, azimuth and elevation relative to the landing runway. For landing at runways with MSBLS ground stations, MSBLS data become available at 20,000 feet for processing by navigation.
One other tool used by navigation is a drag altitude software sensor, which uses a model of the atmosphere to correlate the drag acceleration measured by the IMUs to altitude. This measurement, then, is only as good as the atmospheric model on which it is based. The model is not perfect. However, it has been determined through testing and analysis that drag altitude data are important in keeping downrange and altitude errors bounded during the blackout portion of entry (from approximately 265,000 to 162,000 feet). During this time, the ground is unable to uplink state vector corrections to the orbiter, and TACAN data are not available because of the heat-generated ionization of the atmosphere around the vehicle.
Navigation also maintains a statistical estimate of the expected error in the state vector. This is called a covariance matrix and is propagated along with the state vector. When an external sensor, such as TACAN, becomes available to the navigation software, a check is made to see if the data lie within the current expected range of error. Flight crew controls are provided on an onboard CRT horizontal situation display to force the software to accept or inhibit the external sensor data whether or not the data lie within the expected range. Another control on the display may be selected to allow the software to use the external sensor data to update its state vector so long as the data lie within the expected range.
About five minutes before entry interface, the crew adjusts the software to major mode 304. During this mode, which lasts until TAEM interface, five CRTs become available sequentially and are used to monitor auto guidance and the orbiter trajectory compared to the planned entry profile. The five displays are identical except for the central plot, which shows the orbiter's velocity versus range or energy/weight versus range with a changing scale as the orbiter approaches the landing site. This plot also includes static background lines that allow the crew to monitor the orbiter's progression compared to planned entry profiles.
Once TAEM interface is reached, the software automatically makes a transition to major mode 305. The CRT vertical situation 1 display then becomes available. It includes a central plot of orbiter altitude with respect to range. This plot has three background lines that represent the nominal altitude versus range profile, a dynamic pressure limit in guidance profile and a maximum lift-over-drag profile. At 30,000 feet, the scale and title on the display change to vertical situation 2, and the display is used through landing. When the approach and landing interface conditions are met, a flashing A/L appears on the display.
Another prime CRT display used during entry is the horizontal situation. In addition to providing insight into and control over navigation parameters, this display gives the crew orbiter position and heading information once the orbiter is below 200,000 feet.
The entry trajectory, vertical situation and horizontal situation CRT displays, then, are used by the flight crew to monitor the GN&C software. They can also be used by the crew to determine whether a manual takeover is required.
The crew compartment of the orbiter contains the most complicated displays and controls ever developed for an aerodynamic vehicle. The displays and controls exist in a variety of configurations, with toggle, push button, thumbwheel and rotary switches. Meters are circular and rectangular dials and rectangular tapes. Switches and circuit breakers are positioned in groups corresponding to their functions.
All controls are protected against inadvertent activation. Toggle switches are protected by wicket guards, and lever lock switches are used wherever inadvertent action would be detrimental to flight operations or could damage equipment. Cover guards are used on switches where inadvertent actuation would be irreversible.
The displays and controls in the orbiter crew compartment enable the flight crew members to supervise, control and monitor the space shuttle mission and vehicle. They include controllers, cathode ray tube displays and keyboards, coding and conversion electronics for instruments and controllers, lighting, timing devices, and a caution and warning system.
The displays and controls are designed so that a crew of two can perform normal operations in all mission phases (except payload operations). They are designed to enable a safe return to Earth from either the commander's or pilot's seat; flight-critical displays and controls are accessible from the forward flight deck station from launch to orbital operations and from deorbit to landing rollout.
All displays and controls have dimmable floodlighting in addition to integral meter lighting.
There are more than 2,020 displays and controls in the forward and aft flight decks and middeck of the orbiter. This represents more than 100 times the number of controls and displays found in the average automobile.
Orbiter displays and controls consist of panel displays, mechanical controls and electrically operated controls. Generally, the displays and controls are grouped by function and arranged in operational sequence from left to right or top to bottom with the most critical and most frequently used devices located to maximize the crew's performance and efficiency.
The displays and controls are divided between the forward flight station and aft flight station. The forward station contains all the equipment necessary for the operation of the orbiter. The aft flight station contains displays and controls necessary for rendezvous and docking and for controlling the remote manipulator system and payloads.
The forward flight control area panels are labeled L for the left, or commander's position; R for the right, or pilot's position; F for the front section; O for the overhead position, and C for the lower center section.
The left panels contain circuit breakers, controls and instrumentation for the environmental control and life support system, communications equipment, heating controls, and the trim and body flap controls. The commander's speed brake and thrust controller is on the left panel. The right panels contain more circuit breakers; controls for the fuel cells, hydraulic system, auxiliary power units and engines; and the pilot's communication controls. Electrical power distribution controls are also located on the right-hand panels. The pilot's speed brake and thrust controller is to his left on the center console.
The overhead panel contains lighting controls, the computer voting panel and fuel cell purge controls. The center console contains the flight control system channel selector, air data equipment, and communication and navigation controls. It also contains fuel cell circuit breakers and the pilot's trim and body flap controls.
The center forward panel contains the three cathode ray tube display sets, the caution and warning system, aerosurface position indicators, backup flight control displays, and the fire protection system displays and controls. There are primary flight displays for both the commander and the pilot as well as auxiliary power unit and hydraulic displays and controls for the landing gear. The glareshield contains the head-up display.
The aft flight station contains left, right and center panels. The panels contain the power reactant storage and distribution cryo tank heater control, auxiliary power unit and hydraulic heater controls, reaction control system and orbital maneuvering system heater controls, Ku-band and remote manipulator system pyro jettison controls, communication and utility power plug-ins, translational and rotational hand controllers, an attitude director indicator, Ku-band and S-band communications controls, recorder controls, payload controls, remote manipulator system controls, closed-circuit television controls, a cathode ray tube and keyboard system.
The contractors involved are: Abbott Transistor, Los Angeles, Calif. (transformers); Aerospace Avionics Inc., Bohemia, N.Y. (propellant quantity indicator and annunciators); Aiken Industries, Mechanical Product Division, Jackson, Mich. (thermal circuit breakers); Applied Resources, Fairfield, N.J. (rotary switch); Bendix Corp., Teterboro, N.J. (surface position, alpha Mach, altitude/vertical velocity indicators); Bendix Corp., Davenport, Iowa (accelerometer indicator); SLI System, West Caldwell, N.J. (mission and event timer); Armtec Industries Inc., Manchester, N.H. (digital select thumbwheels, toggle switches); Eldec Corp., Lynwood, Wash. (tape meter); Honeywell Inc., Clearwater, Fla. (flight control system); IBM Corp., Federal Systems Division, Electronics Systems Center, Owego, N.Y. (cathode ray tube display unit, computer keyboard), ILC Technology, Sunnyvale, Calif. (cabin interior and exterior lighting); J.L. Products, Gardena, Calif. (push button switch); Lear Siegler, Grand Rapids, Mich. (attitude director indicator); Martin Marietta, Denver, Colo. (caution and warning status display, limit module); Weston Instruments, Newark, N.Y. (event indicator, electrical indicator meter); Collins-Rockwell, Cedar Rapids, Iowa (display driver unit, horizontal situation indicator); Aeropanel, Parisippany, N.J. (integrally illuminated panels); Betatronix, Hauppauge, N.Y. (potentiometers).
Information content from the NSTS Shuttle Reference Manual (1988)
Last Hypertexed Wednesday October 11 17:47:51 EDT 1995
Jim Dumoulin (firstname.lastname@example.org)
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