The IMUs consist
of an all-attitude, four-gimbal, inertially stabilized platform.
They provide inertial attitude and velocity data to the GN&C; software
functions. Navigation software uses the processed IMU velocity and
attitude data to propagate the orbiter state vector. Guidance uses
the attitude data, along with state vector from the navigation software,
to develop steering commands for flight control. Flight control
uses the IMU attitude data to convert the steering commands into
control surface, engine gimbal (thrust vector control) and reaction
control system thruster fire commands.
could be accomplished with only one, three IMUs are installed on
the orbiter for redundancy. The IMUs are mounted on the navigation
base, which is located inside the crew compartment flight deck forward
of the flight deck control and display panels. The navigation base
mounting platform is pitched down 10.6 degrees from the orbiter's
plus X body axis. The navigation base provides a platform for the
IMUs that can be repeatedly mounted with great accuracy, enabling
the definition of transformations that relate IMU reference frame
measurements to any other reference frame.
The IMU consists
of a platform isolated from vehicle rotations by four gimbals. Since
the platform does not rotate with the vehicle, its orientation remains
fixed, or inertial, in space. The gimbal order from outermost to
innermost is outer roll, pitch, inner roll and azimuth. The platform
is attached to the azimuth gimbal. The inner roll gimbal is a redundant
gimbal used to provide an all-attitude IMU while preventing the
possibility of gimbal-lock (a condition that can occur with a three-gimbal
system and cause the inertial platform to lose its reference). The
outer roll gimbal is driven from error signals generated from disturbances
to the inner roll gimbal. Thus, the inner roll gimbal will remain
at its null position, orthogonal to the pitch gimbal.
sensors consist of two gyros, each with two degrees of freedom,
that provide platform stabilization. The gyros are used to maintain
the platform's inertial orientation by sensing rotations of the
platform caused by vehicle-rotation-induced friction at the gimbal
pivot points. The gyros output a signal that is proportional to
the motion and is used by the gimbal electronics to drive the appropriate
gimbals to null the gyro outputs. Thus, the platform remains essentially
undisturbed, maintaining its inertial orientation while the gimbals
respond to vehicle motion. One gyro-called the vertical gyro-is
oriented so its input axes are aligned with the X and Y platform
axes; its input axes provide IMU platform roll and pitch stabilization.
The second gyro is oriented so that one input axis lies along the
platform's Z axis and the other lies in the X-Y plane. This gyro-the
azimuth gyro-provides platform yaw stabilization with the Z input
axis, while the second input axis is used as a platform rate detector
for built-in test equipment. Each gyro contains a two-axis pick-off
that senses deflection of the rotating wheel. The gyro also contains
a pair of two-axis torquers that provide compensation torquing for
gyro drift and a means to reposition the platform.
The spin axis
of a gyro is its axis of rotation. The inertial stability of the
spin axis is a basic property of gyroscopes and is used in stabilization
loops, which consist of the gyro pick-off, gimbals and gimbal torquers.
When the vehicle is rotated, the platform also tends to rotate due
to friction at the gimbal pivot points. Since the gyro casing is
rigidly mounted to the platform, it will also rotate. The gyro resists
this rotation tendency to remain inertial, but the resistance is
overcome by friction. This rotation is detected by the pick-offs
as a deflection of the rotating gyro wheel. A signal proportional
to this deflection is sent to the gimbal electronics, which routes
the signals to the appropriate torquers, which in turn rotate their
gimbals to null the pick-off point. When the output is nulled, the
loop is closed.
in an IMU are used to measure vehicle attitude. A resolver is located
at one of two pivot points between adjacent gimbals. The IMU resolvers
are electromechanical devices that apply the principle of magnetic
induction to electrically measure the angle between two adjacent
gimbals. This electrical signal is then transformed into a mechanical
angle by the IMU electronics. There are two resolvers on each gimbal:
one-speed (1X) and eight-speed (8X). The 1X electrical output represents
a coarse measurement of the true gimbal mechanical angle. For greater
resolution, the 8X electrical output represents a measurement eight
times that of the true angle. These outputs are converted to an
angle measurement in the IMU electronics and are sent to the GPCs,
where they are combined into a single gimbal angle measurement and
are used to determine vehicle attitude. Attitude information is
used by flight control for turn coordination and steering command
guidance. An attitude director indicator displays attitude and navigation
in each IMU measure linear vehicle accelerations in the IMU inertial
reference frame; one measures the acceleration along the platform's
X and Y axes, the other along the Z axis. The accelerometer is basically
a force rebalance-type instrument. When the accelerometer experiences
an acceleration along its input axes, it causes a pendulum mass
displacement. This displacement is measured by a pick-off device,
which generates an electrical signal that is proportional to the
sensed acceleration. This signal is amplified and returned to a
torquer within the accelerometer, which attempts to reposition the
proof mass to its null (no output) position.
data measured by the IMU are the primary sources that propagate
the orbiter state vector during ascent and entry. On orbit, a sophisticated
drag model is substituted for IMU velocity information, except during
large vehicle accelerations. During large on-orbit accelerations,
IMU velocity data are used in navigation calculations.
can be reoriented by two methods: slewing or pulse torquing. Slewing
rotates the platform at a high rate (72 degrees per minute), while
pulse torquing rotates it very slowly (0.417 degree per minute).
Platform reorientation relies on another property of gyroscopes:
precession. If a force is applied to a spinning gyroscope, the induced
motion is 90 degrees from the input force. In each IMU, a two-axis
torquer is located along the input axes of both gyros. Commands
are sent to the torquers from the GPC to apply a force along the
input axes. The result is a deflection of the gyro spin axis that
is detected and nulled by the stabilization loops. Since the gyro
spin axis is forced to point in a new direction, the platform has
to rotate to null the gyro outputs.
The three IMUs
have skewed orientations-their axes are not coaligned and are not
aligned with the vehicle axes. This is done for two reasons. First,
gimbaled platforms have problems at certain orientations. This skewing
ensures that no more than one IMU will have an orientation problem
for a given attitude. Skew allows resolution of a single-axis failure
on one IMU by multiple axes on another IMU since the possibility
of multiple-axis failure is more remote. Second, skewing is also
used by redundancy management to determine which IMUs have failures.
The IMU platform
is capable of remaining inertial for vehicle rotations of up to
35 degrees per second and angular accelerations of 35 degrees per
second squared. Each IMU interfaces with the five onboard GPCs through
a different flight forward multiplexer/demultiplexer of the data
bus network. Under GPC control, each IMU is capable of orienting
its platform to any attitude, determining platform alignment relative
to a reference and providing velocity and attitude data for flight
thermal control must be maintained in order to meet IMU performance
requirements. The IMU thermal control system consists of an internal
heater system and a forced-air cooling system. The internal heater
system is completely automatic and is powered on when power is initially
applied to the IMU. It continues to operate until the IMU is powered
down. The forced-air cooling consists of three fans that serve all
three IMUs. Only one fan is necessary to provide adequate air flow.
The IMU fan pulls cabin air through the casing of each IMU and cools
it in an IMU heat exchanger before returning it to the cabin. Each
IMU fan is controlled by an individual on/off switch located on
Each IMU is
supplied with redundant 28-volt dc power through separate remote
power controllers when control bus power is applied to the RPCs
by the IMU power switch. The IMU 1 , 2 and 3 on/off power switches
are located on panels O14, O15 and O16, respectively. Loss of one
control bus or one main bus will not cause the loss of an IMU.
Each IMU has
two modes of operation: a warm-up/standby mode and an operate mode.
When the respective IMU switch is positioned to on , that IMU is
powered and enters the warm-up/standby mode, which applies power
only to the heater circuits. It takes approximately 30 minutes for
the IMU to reach its operating range, at which time the IMU enters
a standby mode, when it can be moded to the operate mode by flight
crew command in GN&C; OPS 2, 3 or 9.
To mode the
IMU to operate, the controlling GPC sends the operate discrete to
the IMU through the forward flight multiplexer/demultiplexer. The
IMU, upon receiving this command, initiates its run-up sequence.
sequence first cages the IMU-a process of reorienting the IMU gimbals
and then mechanically locking them into place so that the gyros
may begin to spin. When the IMU is caged, its platform orientation
will be known when it becomes inertial. The caged orientation is
defined as the point at which all resolver outputs are zero. This
causes the IMU platform to lie parallel to the navigation base plane
with its coordinate axes lying parallel to the navigation base's
Once the IMU
gimbals are caged, the gyros begin to spin and power is applied
to the remaining IMU components. When the gyros have reached the
correct spin rate, the stabilization loops are powered, and the
IMU becomes inertial. At this time, the IMU returns an in operate
mode discrete to the GPC, indicating that the run-up sequence is
complete. This process requires approximately 38 seconds.
The IMUs are
in operate by the time the flight crew enters the vehicle before
launch and remain in that state for the duration of the flight unless
powered down to minimize power consumption. While in the operate
mode, the IMU maintains its inertial orientation and is used for
calibrations and preflight, flight and on-orbit alignments.
the IMUs are taken through three levels of calibration to correct
for hardware inaccuracies: factory calibration, hangar calibration
and preflight calibration. Sixty-one IMU parameters are developed
during this extensive calibration period. These parameters are stored
in the orbiter GPC mass memory units and are used in the software
to compensate for hardware inaccuracies.
At T minus
two hours during the launch countdown, the IMU calibration is complete
and the IMUs are ready for the preflight alignment. At T minus one
hour and one minute, theLaunch Control Center initiates this alignment
by a display electronics unit equivalent. (A DEU equivalent is simply
a ground command that looks to the GPCs like a crew keyboard input.)
requires 48 minutes to complete and consists of two different operations:
gyrocompass alignment and velocity/tilt initialization.
In the gyrocompass
alignment, each IMU is oriented so that the desired relative skew
is achieved when the platforms are at their alignment orientation.
During this phase, the IMUs are placed in two orientations relative
to the north-west-up coordinate system. These two orientations differ
only in a 90-degree rotation about the up axis. Data are collected
for 90 seconds by the accelerometers to remove any misalignment
resulting from the reorientation. The accelerometers are used here
because their accuracy is much better than that of the resolvers
and the acceleration due to Earth rotation is definitely known.
Therefore, any unexpected acceleration is due to IMU misalignment.
Once this misalignment is nulled, the platform is torqued about
the north axis to compensate for Earth rotation. Data are then collected
for 10 minutes to measure platform drifts. This sequence of data
collecting is repeated at the second orientation. The relative attitude
errors for each IMU pair are also computed, first with resolver
data and then with accelerometer data. The two values are subtracted
and transformed into body coordinates. A factory-calibrated relative
resolver error term is then subtracted, and a reasonableness test
is performed to check the relative alignment between each IMU pair
to assure a good preflight alignment. The velocity/tilt initialization
mode is then entered, during which the drifts experienced while
waiting for the OPS 1 transition are estimated. The compensation
developed by these drifts is applied to the gyros from the OPS 1
transition to T minus 12 minutes and is also used to compute the
current platform to the mean of 1950 reference stable member matrix
at the OPS 1 transition. In addition, a level-axis test is performed
on each platform three times a second; failure to pass this test
requires the alignment to be repeated.
At T minus
22 seconds, a one-shot data transfer from the primary avionics software
system to the backup flight system is commanded by display electronics
unit equivalent. IMU compensation data computed by the PASS GPCs
in OPS 9 are sent to the BFS GPC at this time so that it will have
the same data for controlling the IMUs if it is engaged.
At the OPS
1 transition, the IMUs enter the ''tuned inertial'' drift compensation
mode. It is tuned because a compensation factor computed in the
velocity/tilt is applied to the IMU gyros. At T minus 12 seconds,
this compensation is removed and the IMUs enter the ''free inertial''
mode. The IMUs are now flight ready, and all functions, both hardware
and software, remain the same throughout the flight.
the IMUs provide accelerometer and resolver data to the GN&C; software
to propagate the state vector, determine attitude and display flight
orbital flight phase, the IMUs provide GN&C; software with attitude
and accelerometer data.
are necessary to correct platform misalignment caused by uncompensated
IMU operation differs only in the manner in which accelerometer
data are used by navigation.
The IMU can
be safely powered off from either the warm-up/standby mode or the
operate mode. If an IMU is moded to standby, an internal timer inhibits
moding operation for three minutes to allow the gyros to spin to
a stop so that the proper sequencing to the operate mode can occur.
The IMU software
scheme is designed to select the best data for GPC use and to detect
system failures. This scheme is referred to as redundancy management.
In the event
of an IMU failure, the IMU red caution and warning light on panel
F7 will be illuminated. If temperatures are out of limits or if
built-in test equipment detects a failure, a fault message and SM
alert will be annunciated.
of the IMU deteriorates with time. If the errors are known, they
can be physically or mathematically corrected. Software based on
preflight calibrations is used to compensate for most of the inaccuracy.
The star trackers and crewman optical alignment sight are used to
determine additional inaccuracies.
The IMU subsystem
operating program processes the data from the IMUs and converts
it to a usable form for other users. The following computations
are performed in the IMU SOP: conversion of velocities to the mean
of 1950 coordinates; conversion of resolver outputs to gimbal angles;
computation of accelerations for displays; performance of additional
software built-in test equipment checks; support of selection, control
and monitoring of IMU submodes of the operate mode; and computation
of torquing commands based on the misalignment determined by the
star trackers, crewman optical alignment sight or another IMU. Misalignments
are due to gyro drifts.
IMU is 10.28 inches high, 11.5 inches wide and 22 inches long and
weighs 58 pounds.
A new high-accuracy
inertial navigation system will be phased in to augment the present
KT-70 IMU during 1988-89. The HAINS will provide spares support
for the inertial navigation system and will eventually phase out
the KT-70 IMU design. Benefits of the HAINS include lower program
costs over the next decade, ongoing production support, improved
performance, lower failure rates, and reduced size and weight. The
HAINS is 9.24 inches high, 8.49 inches wide and 22 inches long.
The unit weighs 43.5 pounds. The HAINS also contains an internal
dedicated microprocessor with memory for processing and storing
compensation and scale factor data from the vendor's calibrations,
thereby reducing the need for extensive initial load data for the
orbiter GPCs. The HAINS is both physically and functionally interchangeable
with the KT-70 IMU.
The IMU contractor
is Singer Electronics Systems Division, Little Falls, N.J.