The guidance system includes the inertial guidance system (IGS), information interfaces with the digital command system (DCS) and the data acquisition system (DAS), with the time reference system (TRS) functioning as an integral part of the guidance system. This system performs navigation and steering logic computation for trajectory management. The guidance system output provides guidance information directly to control systems and/or to instrument displays which the crew interprets for manual control inputs. The Gemini guidance computer, guidance analysis, and guidance system integration subcontractor is IBM Federal Systems Division's Space Guidance Center.The control system provides commands to the thrusters, acting in response to outputs from the horizon sensors, the computer, attitude control and maneuver handles, and the attitude control and maneuver electronics (ACME). The propulsion system is an integral part of the control system. Both automatic and manual modes of operation provide angular control of the spacecraft about its three major axes and for maneuvering. Spacecraft translational maneuvers are manually controlled.
INERTIAL GUIDANCE SYSTEM
The inertial guidance system includes an inertial measuring unit (IMU), digital computer, manual data insertion unit (MDIU), incremental velocity indicator (IVI), and an auxiliary computer power unit (ACPU).
The inertial measuring unit, built by Honeywell, is a stabilized inertial platform including an electronic unit and a power supply. The inertial measurement unit provides a stable attitude reference and incremental velocity data.
The inertial platform a four-gimbal structure with an all-attitude capability. The gimbals. in sequence, starting with the innermost gimbal are pitch, inner-roll. yaw, and outer-roll. The stable element contains three single-degree-of- freedom rate-integrating gyros and three accelerometers of the force rebalance type, mounted orthogonally (mutually perpendicular). Each gimbal contains two analog pickoffs for angle measurement between body and platform axes. One set of pickoffs operates in conjunction with the computer to generate digital representation of the angles. The second set operates with the attitude display group to provide attitude information to the flight crew.
The IMU subsystem electronics provide circuitry and equipment for alignment, stabilization, and gimbal torquing, and accelerometer rebalancing. IMU malfunction detection circuitry is also included.
The IGS power supply operating from 28-volt DC spacecraft power provides the alternating current and direct current power necessary for operations of the IMU, the computer, the MDIU, the IVI, the ACME, and the horizon sensors. The attitude control and maneuver electronics have a separate power supply which can be operated when the inertial guidance system is not operating. The IGS power supply provides alternating current for components normally supplied by the attitude control and maneuver electronics unit, in the event of a ACME power supply failure.
DIGITAL COMPUTER
The onboard digital computer is a binary, fixed point, stored program, general purpose, digital computer which provides executive functions, computations, timing and signal processing for spacecraft guidance and control. The Gemini computer has a memory of 4,096 words (39 bits/ word). It is random access with a non-destructive readout. Division of the memory into three syllables provides flexibility in instruction and data storage location assignment. It can add, subtract, and conduct a transfer operation in 140 microseconds. Multiplication (full precision) is accomplished in 420 microseconds. Division (full precision) is achieved in 840 microseconds. Multiplication and division may be programmed concurrently with addition and subtraction or transfer operations. The clock rate of the Gemini digital computer is 500 KC arithmetic bit rate; 250 KC memory cycle rate.
Computer programs include an executor program, operational programs and standard computational sub-routines. The executor program performs diagnostic checks, determines elapsed time, selects the desired operational program, and performs all data input/output sub-routines.
The basic computer program was inserted at McDonnell prior to computer installation in the spacecraft. After installation, minor updating of constant words and time variable words is accomplished by the digital command system or the manual data insertion unit and is monitored by the data acquisition system. Operational program requirements are functions of the computer mode. These modes will be varied in accordance with the missions to be accomplished by the individual spacecraft. For Spacecraft 3, four modes are available: Prelaunch, Ascent, Catch-Up, Reentry.
In the prelaunch mode, the computer is programmed to perform diagnostic checkout routines and sub-routines. In this mode, sum checks are performed in the section of the computer memory which contains constants and fixed programs that cannot be varied during the mission.
Guidance in the ascent mode serves a backup function in the event of malfunction of the radio guidance system (RGS).
In the ascent mode, information is displayed for flight crew evaluation of mission status. The computer also provides navigational data for reentry guidance in the event an abort should become necessary during the launch phase.
While the RGS has primary responsibility for launch vehicle guidance from liftoff through launch vehicle/spacecraft separation, the inertial guidance system provides backup guidance by effectively duplicating RGS functions in an inertial frame, defined by the launch pad vertical and the desired azimuth at insertion. Stage 1 of standby guidance consists of a time-sequenced roll and pitch program based on the design constraints of the RGS. Stage 2 guidance also satisfies the guidance constraints defined for the RGS. The ascent program may be updated by the digital command system with ground-determined velocity data to reduce the effects of IGS uncertainties. In the event of a guidance switchover, the IGS provides faded (smoothed) attitude error signals to the launch vehicle secondary autopilot. Switchover can occur at anytime. Should switchover occur before liftoff, the switchover signal simultaneously directs engine shut-down, preventing liftoff on backup guidance.
The IGS provides orbit insertion guidance so that the effect of insertion errors on apogee and perigee altitude is nullified. The velocity increment to be added for proper insertion is computed from IGS navigational data and desired insertion condition data stored in the onboard computer. The velocity increment to be added is displayed on the incremental velocity indicator and the attitude error command necessary for proper vehicle orientation is displayed on the flight director indicator. The inertial guidance system then performs navigation computations as the insertion sequence proceeds, driving the incremental velocity indicator display in accordance with the velocity increment added. Orbit insertion guidance begins immediately after a normal second-stage engine cutoff (SSECO), or in the event of a premature engine shut-down, at least 300 seconds from liftoff.
After the incremental velocity has been applied at perigee, the flight crew can read out of the computer, via the MDIU, an incremental velocity to be added at apogee to correct the perigee altitude. The time to apply this velocity can also be read by the MDIU.
In the event of a launch abort after a velocity greater than 21,000 feet per second is achieved, reentry navigation equations will be initiated from the position and velocity information generated by the inertial guidance system. Thirty seconds after equipment section separation, an event initiated by the flight crew, the computer will generate transformation coefficients to relate the platform coordinate system to that required for reentry guidance equations. Following initiation of reentry navigation, the data flow will be identical to that which occurs in the reentry mode.
Between orbit insertion and reentry, the inertial guidance system is operated in the catch-up mode. Various maneuvers are performed in this mode on the GT-3 flight to study the spacecraft's capabilities relative to future missions. The catchup mode provides a means of displaying and applying a ground-command velocity change. The time-to-go to apply thrust and the velocity change to be applied will be computed at the integrated mission control center on the ground, based upon knowledge of the time of liftoff or upon ground tracking information or both. Time and velocity change components will be transmitted to remote tracking sites for transmission to the spacecraft by voice and digital communication. Velocity change is normally transmitted by the digital communication system, however, it may be transmitted by voice and inserted into the onboard computer by the astronauts via the manual data insertion unit. In addition to the MDIU and DCS capability, the velocity change to be applied ma also be displayed by the incremental velocity indicator. Spacecraft maneuver commands are manually applied using orbit attitude and maneuvering system thrusting.
REENTRY GUIDANCE APPLICATION
The Gemini Spacecraft is designed for guided reentries from orbit to permit spacecraft landing at pre-selected, prepared sites. This capability is provided through the use of the inertial guidance system to develop steering commands for proper orientation of the aerodynamic lift vector arising from an offset center of gravity position.
Preparation for reentry begins significantly ahead of retrograde to precisely establish the inertial platform reference and the initial conditions for navigation computation in the reentry mode. Following retrograde, the retrograde section is manually jettisoned and the spacecraft manually controlled to a flight crew "heads down" full-lift attitude. Reentry steering is initiated when the spacecraft enters the atmosphere and total acceleration achieves a prescribed predetermined value. Roll commands control horizontal and vertical components of lift as a function of downrange and crossrange errors. Roll errors are supplied to the attitude control and maneuver electronics for automatic control and also to the flight director indicator for manual control capability. When the density/altitude parameter reaches a pre-selected point within the atmosphere, a maximum lift attitude is commanded and reentry guidance is terminated.
The reentry configuration for Gemini has a center of gravity location resulting in angle-of-attack conditions producing an appreciable lift vector during the atmospheric regions of the reentry. The roll attitude of the spacecraft establishes the direction of lift application; zero bank angle produces range extension, banking to the right a left turn, banking to the left the opposite.
For standard reentry, ground computers determine the future Earth tracks of the spacecraft, determine approximately when the desired landing site will become available, and then perform iterative trajectory solutions to establish a retrograde time that would require about one-half the downrange extension provided by the reentry lift. This time, along with the position and velocity associated with the retrograde time and the target coordinates are transmitted to the spacecraft and inserted into the IGS.
The spacecraft computer, using the initial conditions transmitted from the ground, maintains a knowledge of the trajectory through conventional navigation equations. It then predicts a touchdown on the basis of "zero lift" and compares the predicted touchdown point with the desired target site. Downrange and crossrange errors are used to develop bank angle commands.
The system utilizes the ground complex to develop a very accurate orbit position and subsequent reentry guidance. Ground stations locations have been planned to maintain excellent coverage of the spacecraft during the majority of its flight.
CONTROL SYSTEM
The control system includes the attitude control and maneuver electronics and the horizon sensors. It is used in conjunction with the propulsion system and associated guidance systems to provide spacecraft orientation about its three major axes and for translational maneuvering. The orbit attitude and maneuver thrusters are employed to assist in spacecraft/launch vehicle separation and for attitude control and maneuvering prior to adapter section jettison. During retrograde and reentry, control thrusting is provided by the reentry control system thrusters.
Attitude control and maneuver electronics (ACME) include the attitude control electronics (ACE), orbit attitude and maneuver electronics (OAME), power inverter and two rate gyro packages. Input signals to the attitude control and maneuver electronics are supplied by the IGS, the horizon sensors, or by crew manipulation of the attitude control handle or the maneuver controller, depending on the operational mode the astronauts have selected.
The attitude control electronics accept input signals from the inertial guidance system, from the attitude control handle, from the rate gyros, and from horizon sensors. Input signals are converted into drive commands for the reentry control systems solenoids and to logic commands for the orbit attitude and maneuvering electronics subsystem. Honeywell, inc. is subcontractor for ACME, ACE, and OAME.
The orbit attitude and maneuver electronics accept input signals from the attitude control electronics and the maneuver controller for conversion to drive commands for the OAME solenoids.
Rate gyros sense angular rates about the pitch, roll and yaw axes of the spacecraft. Rate signals are supplied to the ACE in the rate command mode plus all other modes where rate damping is employed. The rate gyro packages also provide inputs to attitude displays and to the telemetry system.
Attitude control and maneuver electronics modes provide for maneuver control to effect spacecraft/launch vehicle separation, for translational maneuvering during the catch-up mode operation, and for maneuvering in all planes by attitude maneuvers and forward/aft thruster operations. (Up, down, left and right maneuvering thrusters are dummy installations in GT-3, therefore eliminating these maneuvers for the GT-3 mission.) Forward or aft displacement of the maneuver controller from the neutral position produces a direct command to the respective solenoid valve driver. The astronauts may select any one of six control modes for spacecraft orientation during orbit and reentry. These modes are rate command, direct, pulse, reentry rate command, horizon scan and reentry.
Rate command mode: Spacecraft angular rates are proportional to rate command signals initiated by flight crew displacement of the attitude control handle. The rate command signals are compared
Guidance System Functions
Mission Phase Applicable Guidance Mode Function Launch Ascent Provides backup guidance in event of a radio guidance system (RGS) malfunction.
Provides attitude error data for astronaut evaluation of mission status.
Provides navigation initial conditions for reentry guidance in event of an abort during boost.Post-SSECO Ascent Provides capability to correct apogee and perigee altitude deviations resulting from insertion velocity, flight path angle, and altitude errors. Orbit Catch-up Provides the capability for displaying and applying ground determined velocity changes. Orbit Prelaunch Provides a diagnostic sum check of computer syllable two instructions.
Provides standby guidance mode wherein digital acquisition is provided and MDIU/DCS locations are telemetered and provides different telemetry quantities.Retro Reentry Provides a real-time navigation during retrofire sequence.
Display of retrograde incremental velocity to astronauts.Reentry Reentry Performs real-time navigation.
Performs continuous prediction of the spacecraft zero lift impact point.
Provides control logic to deliver steering commands to control crossrange and downrange travel.
Provides an automatically controlled reentry in conjunction with reentry attitude control mode.in the attitude control electronics with rate gyro outputs and when the difference between the two signals exceeds the damping dead zone in the system, proper reaction control jets fire for attitude control. This mode is utilized during manually initiated attitude control operation and during retrofire and other velocity change maneuvers. In this mode it is possible to maintain attitude within +/- 1 degree using the flight director indicator for reference.
In the direct mode, switches on the attitude hand controller directly control the ACME solenoid drivers. Control switches are actuated when controller displacement is greater than approximately 1/4 of its full travel.
The pulse mode permits the astronauts to make fine attitude corrections of the spacecraft about its three major axes. In this mode, angular rates are incrementally changed by single thrust commands of fixed duration. Each displacement of the hand controller by the astronaut triggers a single pulse generator in the attitude control electronics and sends a single pulse command to the proper reentry control system or orbit attitude and maneuvering electronics solenoid valve drivers. The astronaut must return the hand controller to the neutral position before he can initiate another pulse to increase his angular rate or effect a braking action in the other direction.
Reentry rate command mode is utilized for manual attitude control during reentry. This mode provides similar operating characteristics to that of the rate command mode, except that damping dead bands are wider and roll rate crossfeed is included in the yaw damping loop, providing for conservation of fuel because the fine control provided by rate command is not necessary for manual performance of this phase of the mission.
The horizon scan mode provides for automatic control of the spacecraft about the pitch and roll axes during the orbit phase of the mission. ACME receives attitude information from the horizon sensor and generates an output to the proper thrusters to maintain the attitude within the damping dead band. When in this mode, the ACME supplies a nose-down pitch bias which enables the flight crew to view the horizon out of the window.
In the reentry mode spacecraft pitch and yaw angular rates are automatically- maintained within a damping dead band by the ACME. A roll attitude is determined by inputs from the computer to the ACME. The computer either provides a bank angle command or a fixed roll rate depending on the relationship between the predicted and the desired touchdown points. ACME will not accept rate commands from the attitude control handle when in the reentry mode.
Two horizon sensors, one primary and one secondary or standby unit, provide reference signals for alignment of the inertial platform and error signals to the ACME for controlling the spacecraft attitude about its pitch and roll axes. Horizon sensors operate by tracking the Earth's infrared horizon.
Two attitude displays, each incorporating a three-axis attitude reference ball with 360 degrees of rotation about each axis, are provided on the right and left instrument panels. These displays, built by the Lear-Siegler Corporation, are slaved to the positions of the normal inertial platform gimbals and provide a continuous all-attitude reference of roll, pitch, and yaw.
Integral rate and command flight director needles display control movements required to position the spacecraft in a commanded attitude or rate. When the commanded attitude or rate is achieved, the needles are centered.
An attitude hand controller is mounted on the console between the astronauts and provides a means of manually controlling the spacecraft attitude and rate in three axes. The controller can be operated by either astronaut, while either is in the restrained position, through simple wrist articulation and palm pivot motion. The controller is spring loaded to provide an increasing resistance as the handle is moved away from neutral. The total travel of the hand controller is +/-10 degrees from neutral in all three axes. Displacement or rotation of the controller causes the spacecraft to turn in the direction of displacement or rotation.
A maneuver hand controller initiates translation of the spacecraft. The controller contains centering springs and six switches, one for initiation of spacecraft displacements in each of six directions along with three major axes. Movement of the handle in any of these six directions initiates corresponding spacecraft translation. The handle may be removed and stored when not in use, providing clearance in the event of seat ejection. Only forward and aft maneuvering is available on the GT-3 spacecraft.
Attitude Control System Characteristics
Mission Phase Applicable Attitude Control Mode Function Launch Direct Rate information to flight crew. Post-SSECO Direct ON-OFF commands to attitude thrusters.
Rate Command Vehicle angular rates are proportional to attitude control handle displacements. Orbit Horizon Scan Automatic pitch and roll attitude control to a horizon sensor reference. Pulse mode capabilities retained for manual over-ride about all axes. Five-degree small-end down pitch bias included to maintain astronaut view of Earth's horizon. Pulse Angular rate can be changed in incremental steps by commanding thrust for a calibrated period of time. Effective about all three axes and can be used for fine attitude control. Rate Command Same as described for Post-SSECO. Direct Same as described for Post SSECO. Retro Rate Command Same as described for Post SSECO. Direct Same as described for Post-SSECO. Reentry Reentry Automatic rate damping about pitch and yaw axes. Roll attitude and rate controlled by computer and ACME. Roll rate cross-feed is included in yaw damping loop. Reentry Reentry-Rate Command Astronaut control of roll orientation. Rate damping provided about all axes. Roll rate cross-feed is included in yaw damping loop. Direct Astronaut control of roll orientation. Manual rate damping required in all axes. ACME Inputs and Outputs
Inputs
- Pitch and roll attitude error signals from the horizon sensors.
- Roll rate or roll attitude signals from the digital computer.
- Pitch, yaw, and roll on-off attitude acceleration commands from the attitude control handle.
- Pitch, yaw, and roll proportional rate commands from the attitude control handle.
- Pulse initiation signal from attitude control handle.
- Translation commands from the maneuver control handle.
- Mode of operation change commands as selected by the crew.
- Test inputs as necessary for satisfactory checkout during ground testing of the ACME in conjunction with associated equipment.
- Redundancy selection commands initiated by the crew.
Outputs
- Attitude maneuver control "on-off" command signals to the affected OAMS and RCS thrust chamber solenoids.
- Pitch, roll, and yaw rate signals.
- Signals to telemetry systems as required for monitoring ACME operation.
Copyright 1997, 1998, 1999 by John
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