Control Requirements of the Shuttle Experiments

W. Haeussermann CONTROL REQUIREMENTS OF THE SHUTTLE EXPERIMENTS Waiter Haeussermann Associate Director for Science Science and Engineering Directorate NASA/Marshall Space Flight Center, AL 35812, USA SUMMARY The great variety of space missions which will be carried out with the Space Shuttle require more or less stringent attitude and pointing control as well as acceleration limitations. For high demands, which cannot be fulfilled by the Shuttle control system, either fine pointing control devices have to be provided for specific instruments or the payload has to become a free-flyer. The Space Shuttle control capabilities as well as fine pointing control systems for experiments operated from the Spacelab are described. Presently planned missions and payloads are reviewed with respect to their pointing control requirements and acceleration limitations.

SPACE SHUTTLE ORBITAL CONTROL CAPABILITIES III Beginning in 1979 the Space Shuttle will be the only large U. S. manned vehicle to transport payloads up to 27 000 kg (60000 lb) into orbit. Like a space station, it will also provide an operational base for scientific and technological experiments and observations. The arbiter's mission time is normally 7 days, but missions up to 30 days are not precluded. Furthermore, free-flying experiments that are designed for longer operation periods or that demand extreme pointing accuracies, such as the Space Telescope (ST), will be released from the Shuttle by a manipulator, and, if necessary, they will be retrieved for repair either in orbit or after being returned to Earth. The arbiter (Figure 1) has the capability to achieve and maintain, within certain thermal attitude constraints, any desired space, orbital object, or earth referenced attitude with respect to either the arbiter navigation base or a payload provided and mounted sensor for payload pointing purposes. The pointing accuracy of the desired reference attitude is a function of the error sources associated with the characteristics of a particular attitude sensor, the vehicle flexure between the attitude sensor and the pointing instrument, the type of control system, and the digital autopilot signal quantization errors. The arbiter is equipped with an inertial navigation, guidance and control system. An inertial measuring unit (IMU) , located in the forward avionics bay (Figure 2), provides the reference signals; it is initially aligned with an uncertainty of ±0.133° laxis, has a drift rate in orbit of ±0.105°;h, axis, and has a readout error of ±O. 073° lalds (all data are 3IT values). The additional error contributed by the flight control subsystem deadband is ±O.l ° laxis. 97

W. Haeussermann

Figure 1: Spacelab mission

STAR TRACKERS

Figure 2: Avionics installation configuration

98

W. Haeussennann Thus, the inertial attitude of the IMU's base and of the Shuttle, if it could be considered a rigid body, may be maintained to an accuracy of ±O. 5° for durations up to 1 h. To stay within this accuracy, the IMU requires realignment. Star trackers mounted next to the IMU can be used to supply a continuous attitude reference and thus avoid the drift of the IMU. Furthennore the orbital control system can accept attitude pointing signals directly from experiment senso rs and maintain the reference attitude within ±O. 44 ° with O. 01° Is pe rturbations if all the vernier reaction control system (RCS) jets are operational and to within 0.1° Is if any of the vernier jets do not function. The direct control by an experiment attitude reference signal avoids pointing errors up to ±2° which could accrue between the IMU and the experiment reference coordinate systems due to vehicle flexure, alignment, and calibration errors. Earth target and local vertical pointing is provided by using the IMU to supply the attitude reference signal. The onboard estimate of the vehicle position is used to compute the pointing angle with respect to the reference coordinate system. Navigation accuracy is a function of orbital geometry with respect to ground tracking stations or tracking and data relay satellites (TORS) and varies with orbital attitude and time. Based upon a 3u navigation uncertainty of 600 m utilizing TORS tracking with the Orbiter in a 180 km circular orbit, continuous Earth-surface-fixed target pointing can be maintained with an accuracy of ±O. 5° dUring 0.5 h after IMU realignment. Then, because of the IMU's drift rate, a realignment of the IMU becomes necessary. Obviously this method does not include attitude errors of the Earth pointing experiment such as from vehicle flexure; signals from a payload mounted star tracker or from a landmark tracker could correct such errors. Table 1 summarizes the various attitude pointing errors of the Orbiter for the different control modes. Additionally expected dynamic errors are so small that they are negligible (see the next topic) • Table 1. Attitude Pointing Accuracy of the Orbiter

Reference

Half-Cone Angle Pointing Accuracy (3 u)

Pointing Accuracy Degradation Rate (3 u)

Duration Between lMU Alignments

Inertial

±0.5° (±2°)*

0.1° Ih, axis

Augmented Inertial

±0.44°

( ±O. 01° Is stability)

Earth-SurfaceFixed Target

±O.5°

O. 1° Ih, axis

0.5 h

Local Vertical

±0.5°

0.1° Ill, axis

1.0h

Orbital Object

1.0h NA

Not Yet Defined

* Payload pointing accuracy without compensation such as for alignment uncertainty. vertical flexure, etc. 99

W. Haeussennann DISTURBANCE FORCES AND TORQUES AFFECTING ATTITUDE CONTROL OF THE ORBITER The Orbiter with its payload has a mass close to 90 000 kg. Thus, any perturbation in the order of 100 N causes an acceleration in the order of 10-4 g, a value which is just the limit for some experiments (Table 2). A perturbation of such a magnitude may be caused either by a single vernier control thruster firing (110 N with a minimum on-time of 40 ms), which occurs approximately once in 5 s, or by a crew motion (see Figure 3, which is derived from Skylab data and zero-g flights). Some attenuation of short duration perturbations, especially from crew motions, might occur because of the elasticity and the dampening effects of the vehicle structure; however, no data are yet available on this effect. Effects of aerodynamic drag on the Orbiter are below 3 x 10- 5 g for an altitude of 180 km and this drops rapidly with increasing altitude. In the gravity gradient stable mode (X-axis in the direction of the local vertical, z-axis perpendicular to the orbit plane), the drag effect is approximately 10- 5 g at 180 km altitude. Angular perturbations as a result of the expected disturbance forces have been evaluated from simulation tests. The angular motions do not exceed ±40"; thus, they are practically negligible in comparison to the Orbiter's pointing errors. The angular acceleration will cause translational motions of the experiments of 10- 4 g or less. For missions without attitude orientation requirements, the Orbiter can operate in either a free drift or a passive gravity gradient stabilized mode to satisfy acceleration levels below 10- 4 g. INSTRUMENT CONTROL REQUIREMENTS AND ACCELERATION LIMITATIONS /21 /3/ /4/ In the following the control and acceleration demands of specific classes of experiments, as they are presently considered, will be given; methods and means to fulfill them will be investigated in the next topic. The control requirements have to be divided into pointing accuracy and stability requirements (see Figure 4, which defines stability and pointing errors within the permissible error envelope). In general the pointing accuracy requirements of most of the instruments are modest and determined by the field of view. The stability requirements are more stringent since, over the exposure period, they determine the quality of the image. The most demanding instruments are the solar physics experiments, which have exposure times up to 20 mill, and the astronomical experiments with observation times up to 90 min.

100

W. Hauessennann Table 2. Orbiter RCS Maximum Acceleration Levels Translational Acceleration (m/s 2)

Direction +x

RCS System

-x

0.17 0.13 0.21

Vernier Thruster

0

0

-z

±q,

+ ~.

0.38 0.31

1.17

1. 32

±y

Primary Thruster

Rotational Acceleration (0 /S2)

+z

0.002 0

0.002

IMPULSE

~

a: o IL

0

~.

± ~.

1. 48

0.74

0.037 0.024 0.017 0.019

= 40 N 5

I

100

§

-

/

/

-100

o

I 0.8

1.6

2.4

3.2

TIME (5)

Figure 3: Crew motion design profile Some of the experiments can obtain their extreme stability requirements internally, such as by a gimbal arrangement peculiar to the instrument or by image motion compensation (IMC). This is comparatively easy for optical instrumentation, but usually it makes the instrument considerably more expensive. For extreme ultraviolet and X-ray telescopes, either the technological limitations are severe or IMC cannot be applied. Table 3 shows summarily how many experiments have the same pointing accuracy and stability requirements. Not listed are experiments for which the demands cannot be fulfilled with the Orbiter's onboard facilities, and which have to become free-flyers.

101

o

N

~ ERROR ENVELOPE

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(l)

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l'I1 l'I1

(l)

POINTING ERROR STABILITY ERROR LINE OF SIGHT tLOS)

LOS ANGLE

POINTING ERROR

-~---

7~?'"

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TIME

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Table 3. Number of Experiments Requiring Different Pointing Stability and Accuracy LOS Pointing Stability/Pointing Accuracy (3a) (Roll requirements are at least one order less) Class of Experiment

0.01" to ~ 0.1"

Accel e ration Limitation

0.1" to < 1"

1" to < 10"

10" to < 40"

2' to 6'

12' to 30'

~r

0/0/2

2

2

6

4

2

Astronomy

1/0/0

3/3/0

4/7/7

0/9/5

0/0/4

0/0/1

Solar Physics

2/0/0

2/8/0

0/1/2

0/3/10

0/1/1

0/4/4

0/3/0

0/0/1

0/6/6

0/4/5

0/2/3 0/0/1

High Energy Astrophysics

0/2/0

0/3/4

Earth Observations

0/5/0

0/9/9

0/1/6

0/2/1

0/2/4

0/2/3

0/3/0

0/8/7

0/0/4

0/4/2

0/5/6

0/0/1

0/2/0

Communication Navigation Earth and Ocean Physics Space Processing Applications

Not Applicable

Life Sciences

Not Applicable

10-3 g 13 10

Atmosphere and Space Physics

Space Technology

10-4g

16 4 Achievable with Orbiter control and instrument attitude signals

Definition of Requirement numbers:

xr/[Zpointing accuracy

[

... stability from EPM and Orbiter stability by IMC

..

IPS IMC ~--

...

,.

- - Achievable ----.. with Orbiter control

,. MPM and di'rect instrument control

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Haeussennann

EXPERIMENT POINTING MOUNTS

It has to be realized that experiments often complement each other and that

flight missions have to accommodate experiments which have to be operated simultaneously and aimed with high precision and/or stability at the same target, or target area. In contrast to this demand, sometimes simultaneous pointing to multiple targets has to be provided also. Thus, the Orbiter has to serve as a space platfonn, not only to align and point some instruments with modest control requirements directly but also to accommodate several subplatfonns or experiment pointing mounts (EPM) for which it reduces disturbing forces and torques to a minimum level. For this purpose the spacefixed orientation will be obtained by the Orbiter's inertial attitude control system, which is capable of accepting attitude pointing signals from a payload mounted sensor. It was intended first to develop a standardized experiment mount (SEM) /5/,

which would provide a better stabilized basis and could carry a gimbal mounted experiment package for highest demands. Figure 5 shows conceptually a solution to this problem - a suspended pallet which is stabilized by control moment gyros and which, indirectly through the pallet suspension, controls the attitude of the Orbiter. Simulation of the total system has shown that a pallet stabilization close to I" can be obtained and that the limiting factor for obtaining a tighter stabilization is the stiction of the pallet suspension and

THE SUSPENDED PALL.ET PROVIDES A HIGH lEVEL OF STABILITY FOR A W,DE VARIETY OF EXPERIMENTS

• OTHER FEATURES: • NO RCS REQUIRED • FULL SKY COVERAGE

• LOW WEIGHT • LARGE MOUNTING SURFACE

• MINIMUM PALLETI SHUTTLE INTERfACE

SHOCK MOUNTED CMG'S APPLY TORQUE DIRECTl Y TO THE PALLET AND SHUTTLE ORIENTATION IS MAINTAINED BV THE INTERCCNNECTING SUSPENSION DEVICES

SPRINGS OR OTHER NON-RIGID SUSPENSION DEVICES ISOLATE THE PALLET FROM SHUfTlE DISTURBANCES

SUCH AS MAN MOTION.

Figul'l' G: Conceptual drawing of thc sllspcndo_'d pallp! 104

W. Haeussermann connections between the pallet and the Orbiter. All other nonlinearities and disturbances have a lower order effect on stabilization. Because of cost and weight considerations, this promising experiment stabilization scheme had to be abandoned in favor of the standardized rigidly mounted pallet with individual experiment control and smaller instrument mounts. The present planning is to use an Instrument Pointing Subsystem for the larger payloads with high accuracy demands (Table 3)and a Miniaturized Pointing Mount for the smaller payloads with similarly high accuracy demands. The arbiter control system will be used to accommodate those systems that have lower accuracy requirements. The Instrument Pointing Subsystem (IPS) This experiment mount will accommodate the largest experiments or a group of experiments for simultaneous operation. Table 4 presents a composite of experiments, whose pointing and stability requirements demand an IPS mounting. The table summarizes the range of sizes and weights for the entire spectrum of fine pointing instruments with their acqUisition and control requirements. Table 4. IPS and MPM Requirements and Limitations

Experiment Characteristics Payload Mass Size Gimbal Range

LOS Roll

Slewing rate for target change

Requirements for IPS Capability

MPM Capability

30... 5000 kg max 3.7 m diameter max 9.5 m long

.•. 500 kg max 1 m diameter max 3 m long

0

0

±60 0 ±90

±70 :1:900 if needed

0

90 /min

Pointing Performance (3, ) LOS Roll

±1"••. 36" ±60"... 360"

Stability Performance (3, ) LOS Roll

±0.1"... 1. 0" ±2".•. 12"

Mass of Mount

IPS:

Approx: Volume of Mount

4.3 m 3

750 kg

0

90 /min

same or better obtainable MPM: 56 kg 0.37 m 3

The control engineer is initially attracted to an experiment pointing mount, which permits shifting the mass center of the pointing package to its center of suspension. This seems to be almost mandatory, because perturbation torques cannot result from forces acting on the payload's center of mass, such as from the arbiter's translational accelerations and also rotational accelerations, if the vehicle's center of mass is close to the IPS center of suspension. This ideal approach requires a voluminous and heavy outside gimbal design as was used in the Apollo Telescope Mount. It may still be awlied for some specific requirements. Simulation studies and test runs have been made on a 105

W. Haeussennann less expensive lightweight inside-out gimbal (lOG) arrangement similar to a universal joint, as it is now planned for the IPS (Figure 6). They have proved that the perturbations from the expected Space Shuttle motions are acceptable with a properly selected soft spring mounting of the IFS base. Bearing stiction and remaining torques from connecting wires are acceptable; thus, an earlier conceived combination of ball bearings with flex pivots was discarded in favor of ball bearings alone. Finally the lOG arrangement becomes preferable because of its smaller mass and less costly manufacturing in comparison to a standard gimbal mount. A simplified diagram of the IPS attitude control system is shown in Figure 7. Star trackers and rate gyros mounted on the IPS platfonn furnish the attitude and attitude rate signals for three-axis control (line of sight (LOS) and Roll). Attitude command signals pennit acquiring the target to within a few degrees; thus, the gimbal readout must have at least a resolution of O. 5° over the total provided gimbal range. Because of the gimbal sequence, control signals for roll and LOS pointing must be resolved into azimuth, elevation, and cross elevation commands. Residual bias and alignment errors can be corrected, and pre-flight and in-flight scale and alignment errors can be compensated. The Miniaturized Pointing Mount (MPM) /6/ To complement the IPS when small experiments have to be operated and a high flexibility in pointing is required, such as for aiming at a target on Earth, a "miniaturized" pointing mount is considered. It has approximately half the size, or it takes about one-tenth of the volume, of the IPS. Figure 8 shows the MPM conceptually with its gimbal mount. This gimbal mount is actually taken from a flight proven star tracker. It differs in gimbal sequence from the IPS arrangement. Although this change does not offer an advantage from a perfonnance viewpoint, there is an operational advantage because the roll axis can be driven directly. The MPM will use basically the same type pedestal mount as the IPS to attenuate external disturbances; the spring isolators will change the characteristics of the disturbance wavefonn like a low pass filter. Table 4 gives some characteristics and perfonnance data of the MPM; pointing and stability perfonnance can be expected to be at least as good as that of the IPS, because of the design which features disproportionally smaller bearings and the resulting lower friction level. As in the case of the IPS, the gimbal system will be disconnected during large acceleration periods (or during ascent and descent) from its payload which will be clamped down to the Spacelab or the pallet. Because of its smaller size, the MPM can also be deployed from an airlock of the Spacelab. SENSOR REQillREMENTS Some of the payloads will provide their own pointing infonnation; others will need additional attitude sensors. The derived signals can be used:

106

PAYLOAD INTEGRATION RING PAYLOAD

GYRO P"ac:AGE OI'TlCAL SENSOR PACKAGE ELEVATION DRI'VE

CROSS ELEVATION DRIVE PALLET

, AZIMUTH DRIVE

DATA ELECTRONICS

-l.-------~~~,~

PAYLOAD CLAMP SUI'l'ORT

:fj P"YLOofID CLAMP SUI'l'ORT

GIM8AL SUPPORT STRUCTURE

JETTISON DEVICE

ttl

~ I:::

-00 00

CD

Figure 6: Instrum ent Pointing Subsyste m o

'-J

a §

~

o

00

::c \l)

ctl ~

fIl fIl

(t)

a

RESIDUAL BIAS 110 ALIGNMENT ERRORS MUST BE < 1" DIFFERENCES DUE TO TRACKER NOISE 110 GYRO DRIFT

C..,...,OMP....,..-Iu~--E-....,I I

.H-20s ID.D5Hll

""'1

DIFFERENCE BETWEEN MEASURED 110 ESTIMATED IPS ATTITUDE

,.

KALMAN FILTER: COMPUTE UPDATES TO ESTIMATED ,PS ATTITUDE 110 ESTIMATED GYRO DRIFT RATES

UPDATE TO ESTIMATED IPS ATTITUDE I

ESTIMATED IPS ATTITUDE

DIGITAL COMPUTER COMMANDED IPS ATTITUDE

Figure 7: Simplified diagram of IPS attitude control system

~

W. Haeussermann

Figure 8: Pallet mounted miniaturized pointing mount 1.

For Shuttle attitude control if desirabl~ to compensate for drift of the Shuttle gyros, for vehicle flexure and alignment errors

2.

For EPM control (e. g., for payload LOS and roll control)

3.

For image motion compensation if the instrument permits it. This type of control is most desirable when utmost pointing and stability performances have to be achieved.

Stabilization information will generally be derived from rate gyros attached to the mounting base of the payload on the EPM; four of them will be used in a skewed arrangement to provide redundancy if one of them should fail. Stellar Instruments Stellar instruments demand the longest stability duration. After coarse acquisition through reference gyros, the star trackers are activated and search for their respective guide stars within their field of view (±2 0). After acquiring the guide stars, the star tracker outputs are used to control the EPM attitude with respect to LOS and roll. The star trackers must have a sensitivity for ninth-order magnitude stars to have a sufficient probability for a guide star available in the field of view. 109

W. Haeussermann Solar Instruments Solar observations usually require simultaneous us'e of Sun-centered instruments and instruments which can be offset to at least four times the solar radius or equivalently to ±1°. These requirements are very similar to the former Apollo Telescope Mount requirements. The LOS sensing of the Suncentered instruments will be obtained by controlling the IPS from signals generated internal to one of the instruments, whereas offset pointing will be controlled by signals from a fine Sun sensor mounted on the IPS and remotely offset as commanded. The offset pointing will either be obtained by image motion control within the instrument or through IPS gimbal control if no Sun centering is required simultaneously. Roll control, which is more modest than LOS control, will be derived from a star tracker which uses a guide star approximately 45 off LOS. 0

Most promising for obtaining the extreme pointing and stability requirements are an ultra fine sun sensor (UFSS) /7/ and a correlation technique using the pattern of solar granules. The UFSS is a further development of the ATM's fine sun sensor /8/. A fine deviation wedge has been added and only the solar limbs are sensed to provide differential detection. The sensor has demonstrated a null sensing resolution of O. 01" (lu) with an offset pointing capability of ±O. 6" resolution; this accuracy is independent of the solar flux and solar limb darkening effects. When it is required to track a certain phenomenon on the Sun over an extended period, the UFSS requires an open-loop control to compensp.te for the Sun's rotation (e. g., the center of the Sun's disk moves along the equator approximately 2" in 15 min). In contrast, the correlation sensor senses deviations directly from the phenomena and permits a closed-loop tracking. Since the sensor yields an accuracy comparable to the UFSS, it guarantees better tracking. Earth Instruments For Earth observation the arbiter will maintain an Earth-center-fixed or local vertical pointing mode, which will be obtained by the Shuttle's navigation and control system using ephemeris and Earth cartographic data. Star trackers will correct the IMU's drift as in the space-fixed mode of the arbiter. If the experiment pointing requirements are modest, instrument pointing can be achieved in an open-loop control mode by programming the pointing attitude with reference to the arbiter's IMU. For more stringent requirements closedloop control is needed and the pointing reference becomes payload peculiar. Pointing signals will be derived accordingly from reference beacons on the Earth's surface or from navigational satellites, from star trackers, or from manned control of a display.

Angular velocities and angular gimbal freedom of Earth experiment mounts are \\ithin the slewing requirements for other mount requirements, such as slewing of astronomical instrumentation for reacquisition. 110

W. H aeussennann SUMMARY AND OUTLOOK The control requirements for the arbiter as seen from the experiments point of view are, despite their stringent character, mainly within the state of the art. Further development is needed on sensors in general and on Earth sensors in particular, such as landmark trackers and applying autocorrelation techniques for pointing as well as angular rates for stabilization infonnation. From the experimenters' side the goal is to reduce image motion control, and simplify the experiments and thereby reduce their cost. However, the technological problems involved in developing a universal experiment mount with oneorder higher accuracy would be even more expensive. ACKNOWLEDGMENT Acknowledgment for contributions to this report is given to Messrs. W. B. Chubb, M. E. Nein, and P. D. Nicaise of the George C. Marshall Space Flight Center. REFERENCES

/1/

Gardiner, R.A. and Bradford, W.C.: Paper 6.1, IFAC Congress, Boston, Massachusetts, August 24-30, 1975.

/2/

Nein, M.E. and Nicaise, P.D.: Experiment Pointing Subsystem (EPS) Requirements for Spacelab Missions, NASA TM X-64978, December 1975. •

/3/

Summarized NASA Payload Descriptions - Sortie Payloads, Level A Data (Preliminary), NASA/MSFC (PD), July 1975.

/4/

Payload Descriptions - Sortie Payloads, Level B Data (Preliminary), NASA/MSFC (PD),July 1975.

/5/

Nicaise, P. D.: Simulation of an Experiment Pointing Subsystem for the Space Shuttle, NASA TM X-64779, September 1973.

/6/

Fritz, Carl G. et al.: A Miniaturized Pointing Mount for Spacelab Missions, NASA TM X-64972, November 25, 1975.

/7/

illtra Fine Sun Sensor, Phase B - Final Report, Honeywell Radiation Center, August I, 1975.

/8/

Johnston, J .D.: A Fine Sun Sensor for Skylab's Apollo Telescope Mount, NASA TM X-64599, March 31, 1971.

111