Closed Loop Dynamic Testing of The Intelsat V Attitude Determination and Control Subsystem Hardware

CLOSED LOOP DYNAMIC TESTING OF THE INTELSAT V ATTITUDE DETERMINATION AND CONTROL SUBSYSTEM HARDWARE H. Bittner, E. Briiderle, A. Brauch, H. Pfefferl, A. Scheit, and CH. Roche Space Division of Messerschmitt-Bolkow-Blohm GmbH, Ottobrunn bei Miinchen, Federal Republic of Germany Abstract: Each set of hardware equipment of the INTELSAT V Attitude Determination and Control Subsystem (ADCS) is subjected to closed loop Dynamic Bench Tests (DBT). The test set-up comprises a three-axes dynamic motion simulator carrying sensor optical heads and associated signal processing electronics, earth and sun simulators for sensor stimulation, control electronics, TM/TC simulator and a hybrid computer for the simulation of spacecraft dynamics, disturbance torque profiles and digital test data recording. The discussions to follow cover closed loop hardware test philosophy as compared to other test concepts, test procedure, dynamic bench test equipment and specifications, simulation software, test data recording and evaluation and examples of test results as obtained during flight model No. 1 ADCS testing. Keywords. INTELSAT V, control system, hardware equipment, dynamic testing, closed loop, subsystem performance.

1.

INTRODUCTION the sic manoeuvre sequence and the related control modes will be given and one particular manoeuvre will then be used as an example to demonstrate in detail the capability of the test facility, the benefits of the test philosophy, the implementation of the test procedure and the validity of FMl test results with respect to subsystem performance requirements.

At present the hardware equipment of flight model number 2 (FM2) of the INTELSAT V ADCS, designed and developed at MBB is undergoing acceptance tests. In order to demonstrate proper subsystem performance under most realistic operational and extremum environmental conditions these tests are performed by executing all in-orbit sic manoeuvres in closed loop, operated by a simulated ground station and subjecting the ADCS hardware equipment to temperature cycling between 0 -20 and +45 centigrade. For the implementation of such test methods a special Dynamic Bench Test Facility (DBTF) has been established at the Control and Simulation Department of the MBB Spaceflight Division, which consists of a high precision three-axes motion simulator carrying sensor heads and signal conditioning electronics in temperature-controlled housings, temperature chambers for the control electronics and flywheel, earth- and sun simulators for sensor stimulation and an EAI hybrid computer for on-line simulation of sic dynamics, kinematic relations, coordinate conversions, data recording and off-line test data evaluation. This test equipment has also been used to perform dynamic tests of sensor equipment and closed loop dynamic testing of development model, engineering model and prototype ADCS hardware according to the objectives and requirements of the respective hardware development phases. In order to facilitate the understanding of the test efforts associated with a complete ADCS subsystem hardware test a short description of

2. THE ATTITUDE DETERMINATION AND CONTROL SUBSYSTEM OF INTELSAT V In transfer orbit the sic is spin stabilized. Spin axis measurement from earth and sun crossing pulses and reorientation by activation of offset thrusters by ground command is performed in the classical manner. The transfer orbit (TO) equipment, consisting of TO sun sensor and infrared sensor and the associated control electronics are not subjected to closed loop testing and will not be dealt with in detail here. The following discussions will be concentrated on the in-orbit operational modes starting with the acquisition manoeuvres. 2.1

Manoeuvre Sequence

The mission sequence of the IV spacecraft is shown in Fig. 2.1. If T3 denotes the time when drift orbit velocity corrections have been compl~ted, the acquisition sequence starts 5 min. later with the despin manoeuvre:

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H. Bittner et al. Acquisition manoeuvres: o Despin to SO/sec ~ Wx all axes o Rate damping to Wx Wz = 0

=

Wy

=

0.5 o/sec;

o Deployment of solar arrays & antennas T3 + 15 min. o Sun acquisition T 3 + lh and earth search Wx = 0.5/0.25 o/s o Earth acquisition roll/pitch: T3 + 4h to T3 + 8h o Wheel run-up (station keeping control loops activated) o Yaw acquisition with normal mode roll/yaw control system (WHECON principle) on-orbit manoeuvres: o Pitch FMW control with automatic unloading o Whecon - roll/yaw regulator operation o Station keeping N/S & E/W Special manoeuvres o Small angle offset pointing o Large angle offset pointing and antenna mapping. The rate damping mode ensures that residual sic body rates of up to 300 /s after de5pin 0 are reduced to commanded values of 0,5 in roll and pitch or zero in yaw respectively to an accuracy of at least 0,13 o/s. The objective of the sun acquisition manoeuvre is to point the sic x-axis to the sun from arbitrary initial attitude and to initiate earth search. The earth acquisition terminates with the z-axis pointing to the earth centre. Yaw acquisition erects the angular momentum vector to a direction perpendicular to the orbit plane. The normal orbit mode of oPeration ensures sic three-axes attitude stabilization in orbit, which is also maintained during corrections of satellite position (station keeping). The special manoeuvres to be performed by the sic serve the purposes of pointing the payload antenna beams to different locations on the earth, offset by 0,2 deg. in roll and 2 deg. in pitch respectively from the nominal pointing direction (small angle offset pointing) or to measure the antenna pattern over radiation areas of ~ 12,5 deg. in roll and pitch (large angle offset pointing) 2.2

Attitude Control Concepts

A review of the essential features of the different sic operational modes and the associated attitude control concepts is given in table 2.1. Furthermore for the discussion of aspects related to ADCS testing in subsequent chapters, the sensor equipment implemented in respective control loops

has been listed. In general it is expected that the information contained in table 2.1 is sufficiently detailed except for some particular aspects of the acquisition modes, which bear consequence on the complexity of related tests and should be more emphasized. The sun acquisition manoeuvre represents a fully automatic mode, initiated on ground command and executed by onboard logic. Necessary switchings of control loops and rate commands are controlled by logic sun presence signals generated by the coarse analog sun sensors. Sun acquisition auto0 matically terminates with 0,5 /s sic roll rate for earth search. Earth acquisition in roll and pitch is performed in separate steps on successive ground commands. For end of life thrust level conditions the commanded roll earth 0 search rate must be reduced to 0,25 /s after the first earth crossings to ensure roll earth capture for the given earth sensor field of view. Pitch acquisition from large 0 pitch offsets (up to 12.5 ) requires switching of the yaw attitude reference from coarse analog to fine digital sun sensor which provides larger off-axis field of view. A part from steady state system operation around terminal state, acquisition manoeuvres from various initial sic attitudes to initiate all subsystem internal switchings and to verify respective transients have to be covered within the tests. In particular motions over large attitude angles use to conflict with the limited diameter of the sun simulator beam and loss of sun reference, causing erroneous control logic states during the test manoeuvre, has to be prevented by respective motions of the reference source. 2.3

Attitude Control During Orbit Corrections. (Station keeping)

As has already been mentioned one particular I.V attitude control mode will be dealt with in detail throughout the paper as a specific example, by means of which the subsystem and test aspects can be sufficiently illustrated. A station keeping manoeuvre has been selected for this purpose. Therefore the control loops used for attitude stabilization during orbit corrections will subsequently be described. A functional block diagram of the IV station keeping attitude control loops is shown in Fig. 2.2. Attitude reference is derived from a geostationary IR-sensor in pitch and roll and from a wide angle fine digital sun sensor (FDSS) in yaw. Lead/lag compensation networks, dead bands and PWPF modulation of reaction jets is employed in all axes. Three values of controller gains can be selected on ground command to match attitude accuracy and loop dynamics to disturbance torque conditions for respective orbit correction manoeuvres and to control torque variations over sic mission life time. The minimum pulse bit size of the modulator can be set to nominal values of 125 or 25 msec respectively from ground to limit the maximum number of thruster firings for large disturbance

Closed Loop Dynamic Testing torques on the one hand and to ensure low body rates before transition to normal mode on the other. Furthermore adjustment of the FDSS output sensitivity to gain variations originating from sun incidence angles according to sic sun constellation can be performed from ground. In Fig. 2.2 test points have been entered, which identify the signals recorded on strip chart or respectively digital magnetic tape for quick look assessment and/or detailed subsystem performance evaluation. 3.

TEST PHILOSOPHY

3.1

Hardware Test Philosophy for IV

The necessity for most affirmative test methods for space system hardware equipment, expected to operate for seven years under severe environmental temperature conditions in orbit, after having passed heavy vibrational loads during launch is evident. Concentrating on the functional tests only for the moment, in fact different types of tests have to be performed, both, in open loop and closed loop, as well as static and dynamic manner. Open loop tests on the one hand offer the advantages of: - high sensitivity and accuracy - easy localization of (primarily static) errors Interface incompatibilities with other equipment in the control loop and unforeseen dynamic phenomena, however, are difficult to de tect. Closed loop tests on the other hand have the following advantages: - Component operation takes place under realistic signal conditions Signal sign and proper arrangement is easily verified - Testing of interface compatibility between subsystem components is covered - Operational subsystem performance evaluation as compared to requirements is easily possible Test results represent reference recordings for inflight TM measurements - Comparison with theoretical analysis and simulation results is at hand - Verification of mathematical and simulation models is achieved - Validity of sensitivity analyses is checked - Testing of closed loop system with open loop sensitivity is possible (model reference technique) Disadvantages of closed loop test methods are: - Higher complexity is inevitable - Localization of error sources is difficult due to error propagation along the loop

A.CI.S.

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Before going into a detailed discussion of the test methods finally adopted for IV ADCS hardware testing, the test requirements and characteristic properties of test equipment capable of meeting these requirements have to be considered. 3.2

ADCS Hardware Test Requirements

When specifying ADCS hardware test requirements and necessary performance of test facilities the resolution accuracy of sensor equipment and dynamic range of the system to be tested have to be considered. In table 3.1 the main characteristics of the sensors used for attitude determination are summarized. The operational range on the one hand gives an indication of the maximum attitude angles, which are encountered during the tests and sensor resolution a measure of the subsystem accuracy, which must be resolved. The test facility must be able to at least operate over the whole range (except for maximum rates of the ARA) and provide a resolution, which is an order of magnitude better than that of the sensors. Furthermore a number of additional requirements has been imposed on the IV. ADCS testing: - TLM and test data recording, which enables complete system performance evaluation - Temperature cycling of all hardware components within a maximum range of -25 0 and 0 +50 centigrade. - Switching between redundant components (sensors and control electronics) during manoeuvre operation - Testing of acquisition manoeuvres from initial attitudes corresponding to sunand earth reference out of sun- and earth sensor fields of view (see table 3.1) - Subsystem operation under test conditions for at least o a sufficient period to establish feasibility and performance of all control modes using development model hardware (without temperature cycling) o 200 h subsystem operation using engineering model hardware o 300 h (400 h if no ABT-tests are performed) subsystem test operation covering all manoeuvre modes and transitions with prototype ADCS hardware components including special tests o 96 h for each set of flight model equipment (F1 to F7) covering all sic operational modes and transitions according to a redefined test procedure and temgerature cycling between -20 o and +40 . In view of the test requirements listed above implementation of a closed loop test principle was obligatory for subsystem level tests. Closed loop dynamic testing can in principle be performed using Air Bearing Table (ABT) or Dynamic Bench Test (DBT) equipment. In fact the customer originally

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requested both ABT and DBT. The final decision on the type of tests to be performed was made on the basis of the feasibility of ABT and DBT equipment as compared to the test requirements. 3.3

Feasibility of Test Requirements with ABT Test Facility

Testing of hardware equipment in closed loop requires 1:1 time and signal amplitude scale to ensure proper system dynamics and signal to noise ratio. When comparing the physical characteristics of the ABT equipment available, with the test requirements, the following situation was encountered: 2 - If sic moments of inertia of 2400 Nms about roll and yaw axis were to be realized the air bearing was not able to carry the weight of the mechanical simulation model - Typical disturbance torque levels of the air bearing and test table environment are about 3.10- 3 Nm whereas the sic disturbance torques are of the order 3-10- 5 Nm under normal in-orbit operation conditions. In order to establish some similarity between system in-orbit operation and ABT-testing attitude and dynamic rescaling would have been required e.g. according to the model condition that the relation between the three main characteristics o~ sic attitude motion - nutation - attitude drift through controller deadband due to (artificial) disturbance torques and - earth rotation rate remain to be the same, Gain factors time constants and deadbands in the ADCS hardware then have to be changed accordingly and reestablished after testing, which is an intolerable situation and actually no hardware test. The ABT test idea also conflicted with TLM data recording, temperature cycling and large attitude angle requirements, the latter two causing additional table imbalances. The situation encountered when analyzing the possibility of testing the station keeping attitude control system on the ABT was not quite so desasterous but still subjected to severe restrictions. For such reasons the ABT tests have been dropped completely in agreement with the customer and have been replaced by additional 100 h of operational test time on the DBTF. No major difficulty has been encountered to meet the test requirements of item 3.2 using a DBTF, which will be described in more detail in section 4.

3.4

Test procedure

The test procedure for the Intelsat V ADCS DB-tests has been established in parallel to the development- and engineering model tests. In general, the procedure has to reflect the requirements of the qualification and acceptance test plan and to define the details of the test performance. In particular the following items have to be covered by such a procedure: - test philosophy, with an explanation of extent, validity and limitations of the subsystem tests; description of the facility, comprising principle test configuration, flight motion simulator, optical simulators, servo- and simulation computers; - description of the hardware test set up containing details of component location, mounting and adjustment procedures; - description of simulation software as spacecraft dynamics, mathematical models of simulated components, disturbance torque profiles, servo control loops, data handling and other computer interface; - description of the test control unit,which in case of Intelsat DB-tests consisted of a power supply for the different ADCS components, a telemetry display and a manual telecommand unit, which permits access to the control system similar to later flight operations via ground station; - detailed description of manoeuvre simulation comprising hardware identification, parameter recording requirements, acceptance criteria and comparative computer simulation runs, table of significant parameters and initial conditions for each test to be performed and a command sequence and action list for ADCE conditioning, manoeuvre preparation, execution and termination; - annex with significant performance data of test equipment used in the simulation. In the following an extract from the test procedure for a station keeping manoeuvre is presented. Table 3.4.1 contains the initial conditions and parameters for one selected North/South correction manoeuvre (N/S-4/5-4 ). The command sequence and action for the test performance of the manoeuvre N/S-4/5-4 is given in table 3.4.2. This command list, verified during the tests is used as a basis for the design of the operational flight plan.

4.

DYNAMIC BENCH TEST FACILITY (DBTF)

Subsequently a description of the DBTF, especially established to perform closed loop dynamic hardware testing of the INTELSAT V attitude determination and control subsystem will be given.

Closed Loop Dynamic Testing 4.1

Accuracy Requirements and Servo Table Specifications

The requested resolution accuracy of the test facility is dictated by the resolution of the sensor equipment under test. For the numeral values of sensor characteristics reference is made to table 3.1. As a rule of thumb the resolution accuracy of the test equipment must be one order of magnitude higher than the sensor resolution. The maximum range of attitude angles reflected in the gimbal angular freedom of the servo table must exeed the sensor field of view as a consequence of the test requirements listed in section 3.2. Furthermore constraints imposed on the sic manoeuvres due to shadowing of the sun simulator beam as seen from the sensor optical heads by the gimbal arrangement have to be avoided as far as possible. Finally the dynamic response of the test equipment must be at least an order of magnitude better than the sic dynamics under extremum operational rate and acceleration conditions in order to prevent impact of the test facility on the subsystem closed loop dynamic behaviour. Although sic motions use to be fairly slow, care must be taken that structura"l flexibility oscillations do not experience noticeable phase shift, which would cause instability of structural modes under test conditions. For INTELSAT V the first bending mode frequencies are about 0,5 Hz for the first normal bending, 1 Hz for first torsional and 1,5 Hz for the first in-plane bending mode, sic solar array flexibility and stiffness of the attachment walls included. The performance specifications of the dynamic motion simulator are summarized in table 4.1. Girnbal angular freedom, minimum and maximum rates in respective axes, angular rates, attitude accuracy and resolution as well as payload capacity are given. 4.2

Mechanical Arrangement of Dynamic Motion Simulator

The mechanical arrangement of the dynamic motion simulator is shown in fig. 4.1. It consists of a gal low of welded steel plates mounted on a steel construction base. The top to bottom dimensions are 2,6 m. The 0shaped center gimbal arranged within the gallow can be rotated about tne vert~cal axis by + 120 deg, this axis of rotation representing the sic yaw (z-)axis. It is driven by a hydraulic motor. The U-shaped middle gimbal providing free sensor field of view in almost a hemispherical region rotates about a horizontal axis with + 60 deg. of angular freedom. This axis of-rotation corresponds to the sic roll (x- axis). Again a hydraulic drive motor is used. The inner diameter of this frame is 0,8 m allowing for quite a bulky test payload to be fixed on the inner gimbal. The inner gimbal consists of a mounting platP., driven by a DC servomotor which is moun-

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ted in the middle of the U-shaped middle gimbale This axis of rotation corresponds to the sic pitch(y- axi~ and is unlimited in angular degree of freedom (continuous pitch) . The DC-motor provides a maximum torque of 45 Nm. All rotary gimbals are made of a special aluminium alloy which ensures high torsional stiffness and light weight for rapid acceleration. For signal- and electrical energy transmission from and to the test payload 96 slipringscapable of standing different levels of transmitted power are available. 4.3

Servo Table Control Loops and Performance.

Attitude measurement of the gimbal motion is performed by electro-optical pick-ups in all axes, providing an angular resolution of two arc seconds. Additionally analog tacho generators are connected to the gimbal axes to der~ve damping signals for the servo loops. A photograph of the servo table with the sensor equipment mounted on the inner gimbal is shown in fig. 4.1 and the complete test room with strip chart recording equipment in the foreground TCU and temperaturechamber to the left, servo table in the middle and table control equipment at the right hand side is shown in fig. 4.2. The outer and U-shaped gimbal are furthermore equipped with accelerometers to control gimbal vibration initiated by rapid motions of the high power hydraulic actuators. Servo table attitude is contrd If·d by a Honeywell 316 R digital process control computer having 16 bit word length (single precision) and a memory cycle time of 1,6 ~sec. The servo table control programm, written partly in FORTRAN IV and partly in assembler language to save computation time, occupies about 4,5 k words of memory. The sampling rate of gimbal attitude measurement and processing is 200 Hz. Tachogenerator and accelerometer signals are processed in analog phase shaping networks, bypassing the digital control computer, and directly added to the power amplifier inputs, which drive the MOOG-valves of the hydraulic actuators or the DC-motor respectively. In fig. 4.3 a typical example of the servo table dynamic performance in roll (U-shaped frame) as obtained during acceptance tests of the DBTF is shown. For input signal amplitudes of 0,25 deg. commanded and actual girnbal attitude angle as well as attitude control error for 0,1 Hz sine wave and triangular wave shape of the test signal are recorded. As can be seen the average attitude error signal is smaller than 0,0027 deg. under such dynamic conditions (bit noise ignored). 4.4

Closed loop Dynamic Bench Test Set-Up

In fig. 4.4 a block diagram of the I.V closed loop dynamic bench test set-up is shown. The right hand part of the diagram

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refers to the dynamic motion simulator (3-axes table) with the associated girnbal control loops. Actual servo table position is fed to the digital control computer and compared to the commanded attitude, which arrives through the input buffer. The attitude errors are converted to analog and used to drive the servo table actuators via the analog control equipment, comprising power amplifiers and phase shaping of tacho and accelerometer signals. Additional DA-converters for data registration purposes by analog or digital means are also indicated. Output signals of sensors and gyros under test, mounted in temperature controlled housings on the servo table, are fed into the Attitude Determination and Control Electronics (ADCE), which is located in a temperature chamber. The logic state of the ADCE is controlled and the signals within the control electronics are monitored by the external Telecommand Unit (TCU), which simulates the ground station (left hand bottom part of fig. 4.5). The information on subsystem hardware available in the TCU, which contains actual ADCE logic status, sensor outputs, wheel control torque, thruster activation pulses and so on is transmitted to the sic simulation computer and used to drive the sic dynamic simulation model or is stored on magnetic tape for manoeuvre reconstruction and data evaluation respectively. The control loop is closed by transmitting simulated sic attitude to the input buffer of the servo table control computer as girnbal positioning commands. In order to include the real flywheel and the associated angular momentum dumping logic in the ADCS tests a model following loop has been implemented. Wheel torque commands from the ADCE are fed into the real flywheel and a flywheel simulation model in parallel, measured actual and simulated wheel speed are compared, and the difference is used to correct deviations of the simulation model originating from the unkrown part of flywheel friction. Temperature control of the sensor equipment on the servo table is performed by Peltierelements, which can be used for cooling or heating depending on the current direction. Loss energy is removed by water cooling. The actual temperature is measured by thermopiles mounted at the base points of the sensor equipment and the current of the Peltier elements is controlled by an external control unit. The sic simulation computer is a SEL dig. computer, which forms integral part of an EAI 3200 hybrid installation. It is a true 32 bit machine with 600 nsec basic memory cycle time, the memory being four times overlapped, which means that a memory request can be initiated every 150 neec. The hybrid interface is used for interrupt control and data conversion. The dynamic sic simulation program is written in FORTRAN IV and is executed together with the interface control

program and including data recording on magnetic tape within 8 msec. 4.5

Simulation Software

When performing closed loop Dynamic Bench Tests a simulation model of the sic dynamics in all operational modes has to be established an thorougly checked for validity. The sic dynamic equations of motion are expected to be well known or may be looked up in the relevant literature and will therefore not be derived or repeated here. (see for instance "The Attitude Determination and Control Subsystem of the I.V. sic", Proceedings of AOCS Conference, Nordwijk, Oct. 1977, ESA SP-128, Nov. 1977). Only the simulation model used for closed loop testing of the station keeping attitude control system will be elaborated upon in some more detail, in view of the fact that this loop is carried through the paper as a particular test example. The emphasis of the discussion to follow will rather be concentrated on the methods applied for simulation software testing. 4.5.1

Simulation Model of sic Dynamics in Station Keeping Control Mode

A block diagram representing the sic dynamics in the station keeping control mode - for the pitch (8) axis including rigid body and torsional panel bending (~) and

(t,

- for the rolllyaw rOlllyaw ~) axes including rigid body, normal panel bending mode (qB) and angular momentum (H) coupling is shown in fig. 4.s. This block diagram is a good representation of the essential characteristics and the associated equations of motion can be readily reconstructed. 4.5.2

Simulation Software Testing

Once the sic simulation program has been established check runs have to be performed, which provide a means for testing the DBT simulation software. Four types of tests have been executed: - Rate cross-coupling and kinematic equations This test provides a check of all rate terms in the equations - Normal mode equations Normal mode rolllyaw equations are tested. - Disturbance torques Proper influence of roIVyaw disturbance torques is considered - Wheel loop The dynamics of the wheel loop is checked. As to the panel flexibility only the first normal bending mode has been implemented. For the rate test the panel is coupled with the yaw axis only. In the normal mode rolllyaw dynamics there are three possible modes: - the panel mode - the nutation - the orbit mode

Closed Loop Dynamic Testing Considering the panel mode, which is especially of importance for the station keeping manoeuvre the orbit mode can be neglected. On the other hand if the orbit mode is checked, the influence of the panels can be neglected. For the investigation of the influence of the disturbance torques two types have been considered - disturbance torques with orbit frequency - disturbance torques with multiples of the orbit frequency The first type of disturbance torques causes increasing oscillation amplitudes while the latter one results only in limited amplitudes. The wheel loop, which is the only closed loop in the simulation program is decoupled from the roll/yaw axes. The time constant of the wheel speed adaptation has been checked. In order to have an easy means for checking all cases, initial conditions and if necessary special parameter values have been assumed, which ensure that the resulting solution consists only of one mode, whose amplitude and frequency is predicted for each state vector component by analytical solutions. 4.5.3

-

~

Fig. 4.6 a: Yaw phase plane plot, excitation of nutation mode only

- 6 CASS-OH's + 1 CASSE (red.) 1 TOSSA (red.)

-

- 2 TOIRS'+ TOIRSE's. Because of the limited space available on the inner platform of the FMS and the limited number of sliprings, only those sensors being in operation are mounted on the table. All others are installed in the stationary temperature chambers, together with their associated electronics and the ADCE. The FMWA is tested in a second temperature chamber because of the different temperature test conditions defined for this equipment. There exist 3 significant sensor configurations according to the different miss.ion phases: A)

As

be observed from figs. 4.5 a and b not only one mode was excited, as intended but also a second one, but with very small amplitude.

4.6

c~ ..

Sensor Arrangement for DBT's

The Subsystem Components, involved in the DBT consist of the following units (for a fully redundant system) - 2 ADCE"s

The typical synchron orbit sensor configuration includes two (or three) GEOIRS and the four FOSS-OH with electronics on the table. This configuration allows testing of the following manoeuvres: Normal mode (including roll/yaw control, flywheel operation) Small offset pointing Antenna measurement mode Station keeping mode Transition from station keeping to normal mode,yaw acquisition mode Box switching

B)

The sensor configuration used for acquisition mode consists of one (or two) GEOIRS", 6 CASS-OH + Electronics and 2 ARA"s. This configuration allows the testing of the following manoeuvres Rate Damping Mode Sun Acquisition Earth Acquisition (Roll) Earth Acquisition (Pitch)

C)

Fig. 4.6 b: Yaw phase plane plot, excitation of panel mode only Fig. 4.6 c: Roll/yaw plot, excitation of orbit mode only.

FDSS-OH's + FDSSE (red.)

- 2 ARA's

Results of Simulation Software Testing

As a particular example for the type of (predicted) results of simulation software testing obtained when following the software test procedure outlined above, the yaw motion has been selected. The three modes of oscillation, nutation mode, panel mode and orbit mode individually excited by proper initial conditions have been recorded in phase plane plots or roll versus yaw attitude respctively, because they demonstrate the sinusoidal motion with constant amplitude and small amplitude variations can be recognized much easier than in time history plots. Fig. 4.6 represents respectively:

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Transfer Orbit Manoeuvres use the 2 TOIRS and the internal redundant V-Slit Sensor (TOSSA). This manoeuvre is not tested on the FMS.

The mechanical Sensor arrangement on the table for configurations A and B can be found in fig. 4.7 und 4.8 respectively.

5.

Test Data Recording of Hardware and Simulation Signals

According to the different objectives of the data evaluation, three different types of recording have been implemented for Intelsat V ADCS-DB-testing: - Strip chart recordings - Analog tape recording - Digital tape recordings

- 2 FMWA"s - 3 GEO-IRS"

5.1

Strip chart recording (SCR)

The main purpose of SCR is to provide conti-

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nuous analog data recording of the most significant signals for a quick look assessment of system operation and performance. The selection of the signals to be recorded during a test depends strongly on manoeuvre type. In the case of a station keeping manoeuvre the following data are recorded: - Torque Signals (ON/OFF) - Computed angular velocities - Commanded table positions

5.3

Analog tape recordings

In addition to the above mentioned recording techniques, a 14-channel magnetic tape recorder was used to record the most important analog data signals for the purpose of trouble shooting. In the case of anomalies, a play back of the signals in question is possible with the possibility of time scale expansion. If no anomaly was recognized on the strip charts, the tapes have been erased.

- Actual table positions 6.

- GEO-IRS-Signals - Earth presence

(ON/OFF)

- FDSS-Signal - Sun presence (ON/OFF) - Simulated wheel speed - Real wheel speed - Wheel unload signal (ON/OFF) - Wheel torque command - First bending mode of flexible panels - Time code (ON/OFF) - Time tacks (ON/OFF) - Event marks (ON/OFF) The time synchronization of the different recorders is achieved by recording identical IRIG-time code signals on each of the 6-channel recorders. 5.2

Digital tape recordings

The digital tape recordings are used as a basis for automatic off-line data evaluation. The recording comprises 5 diffenent types of records: - Initial conditions (rates, positions, system parameters, local time etc.) - Analog data (all analog signals as rates, positions, sensor signals etc. are recorded at a minimum sampling period of 120 msec. ) - Thruster firings are recorded in an asyncroneous manner, whenever a firing occurs. The minimum sampling time is 8 ms. - Telemetry data, mainframe and subframe are recorded at a rate of 512 msec. This frequency is identical to the on board data read out frequency. - End conditions are recorded (similar to initial conditions) at the end of a test and are closed by an end of file mark. The extend of recorded data is selected such as to enable the test to be continued at the same state where it had been stopped (if necessary) . The recorded variables are the same as for strip chart recordings except additional telemetry information.

6.1

Data Evaluation Quick look assessment

As mentioned above, the analog striP chart recordings (24 channel~ provide a measure of quick decision on correct operation of the subsystem and/or test facility respectively. The performance of the system operation can easily be reviewed and in case of anomalies the sources for malfunction can be traced throughout the loops, in particular whether the malfunction originates from a component of the ADCS-hardware or from the test set-up. In addition during long time runs, the identification of short time disturbances (spikes, " g litches") is facilitated. The quick look assessment is further used for control and interpretation of automatic data evaluation program results. 6.2

Automatic data

r~duction

The basis of a computer aided evaluation of test data is digital tape recording as described in para 5.2. The evaluation is performed off line and does not restrict the continuation of the tests.The tape containes a set of files each of which corresponds to one test run. The program first selects the desired file (test) to be evaluated. After identification of the different modes of operation during this particular rlill, (normal mode-station keeping-normal mod~, regulator gain, sensor gain, gain changes etc.) the program starts the data evaluation of each mode of operation or section with the manoeuvre specific subroutine. In case of a station keeping manoeuvre the following information is derived from recorded test data: - Disturbance torque level - Control torques - Controller gains - Number of thruster firings

(for each axis)

- Mean pulse duration - Total fuel consumption (for each axis) - Fuel efficiency - Steady state error - Variance - Max. and min. errors - Histogram

Closed Loop Dynamic Testing - Comparison between sensor output (digital telementry) and actual table position (for determination of bias errors). A typical example of data evaluation results is given in the following chapter 7. 7.

Example of Test Results

To demonstrate typical test results, the particular station keeping manoeuvre which has been described in section 3.4 has been selected. The test conditions were as follows: - Ngrth/south correction manoeuvre, testN- - 4/5-4 - 14.00 hrs local time (west looking sensor system) - FOSS-gain 1, 75

(/::J

-compensation)

- ADCE-gain 1,15

(/3

-compensation)

- GEO-IRS in "north scan inhibit" mode (more noise) Regulator gains "medium" in all axes - Temperature of ADCS-hardware "hot" - Initial conditions as defined in table 3.4.1 under test no. 4/5-4 - BOL torque levels Disturbing torques were maximum specified values of individual axes: -1 roll = 5,22 ~110-2 Nm; pitch = 1,53·10 Nm yaw = 2,9.10 Nm - The disturbance torque sweep starts at 120% of nominal max. value and is reduced in steps of 20% - As a modification to the manoeuvre sequence presented in section 3_4.2 it is omitted here to switch starting from "normal mode" to "station keeping mode" then to "transition mode" and back to "normal mode". - Furthermore the duration time of the manoeuvre has been extended and the regulator gain switching was reduced to a single "medium gain" state in all axes. The strip chart recordings of this case are presented in fig. 7.1 up to fig. 7.3. The length of the records is limited to the first two disturbance torque steps. Data associated with one axis are concentrated in one figure. Flywheel data are omitted because of minor importance for station keeping manoeuvres (wheel in speed mode). The variables recorded and the scale factors are indicated in the respective figures. For comparison an excerpt of the digital evaluation of these two examples is presented in tab. 7~1, each example consisting of two separately evaluated periods of disturbance 'torque t.orque steps. The evaluation results compare very well to a quick look assessment from the strip charts.

8.

391

Benefits of I.V. DBT for ADCS functioning

Testing the various models of the I.V. ADCS hardware, different emphasis is layed on the results to be obtained. In the development phase (development and engineering model tests) the performance of the subsystem is major subject of tests, as well as verification, or adaptation of mathematical models used so far for control system design and sensitivity analysis. In this respect, the development phase DBT yielded the following major results: - The GEO-IRS turned out to be very susceptible to acoustic noise. This was not only a special test environment problem, but had to be regarded on spacecraft level requiring a redesign of some significant mounting brackets. - Comparison of the results of sensitivity analysis and first DB-tests of station keeping manoeuvres showed considerable discrepancies in the dynamic behaviour of the loops, especially sensitivity to panel oscillations. A detailed investigation disclosed a phase shift of the FOSS up to 40 deg. depending on the input amplitude. As a consequence of this result, the sampling frequency of the sensor has been increased by a factor of roughly 30 times. The prototype tests represent the final qualification of the ADCS~hardware under worst case environmental conditions and long time operation. During these tests - a temperature drift of the GEO-IRS pitch output signal of about 0,3 deg over a temperature range of -25° to +50°C has been recognized. This effect could be eliminated by changing the design of the scan mirror suspension within the sensor. - An error in the sun rise/sun set switching logic caused a spike in the output signal when north- or south scan was inhibited by ground command. This effect could not be accepted for normal mode operation and has been eliminated by changing the switching logic. The performance requirements of the subsystem has been proved in long time runs and under extreme temperature conditions as well as during temperature transients (2°C/min) to meet specified values.

9. Summary The closed loop Attitude Determination and Control System (ADCS) test philosophy described in this paper has been developed for SYMPHONIE and has been extended and improved for INTELSAT V subsystem testing. Verification of sensitivity analyses versus test results, subsystem performance under realistic and extremum operational conditions ~long­ time runs, temperatu~e range) versus specifications and requirements are essential characteristics of such test metfi6ds,-which esta-

392

H. Bittner et al.

bli~h a high confidence level into proper in-

sent an optimal possibility for the training of ground station personal and enable an interpretation of the actual spacecraft behaviour by means of the "operational footprint" established during testing.

orb1t,sUbsystem operation. Additionally ADCS handllng by means of the telecommand unit and ~bservation of the telemetry data under rea11stic ground station conditions repre-

10. FIGURES AND TABLES

\ '/~ ,£ E L SPlli>,JUP

\

""'0

(O"y,,"'D TO

NO"""L

\

MOO~

Fig. 2.1. Mission phases covered by dynamic bench testing GAIN SWI1CHING

PULSE BIT SIZE

}

GROUND COMMANDS

0

J.l

t .-Jo

r

SPACE

I

I

CRAFT

~

DYNAMICS

I

RIGID

~

BODY;

!

STRUCTURAL


I

~AD:LAG

C~T=

FLEXIBIL:TY

-J

COUPLING

t--I

r_! __r -__I I

o

TESTPOINTS, I

I

0 0

CV

~=tl. I

GYR~~PiC I

O

0 0 0

I

! I

! i

O-'=-J=:O-S--"s-~=T--

~:-..-1 :CNSOR

o CCUPLlNG

__1i--_Y_A....

~L.-:----------=O=---II.-G_ER_OO_-L_~~_S _II-.-~l-'-
PITCH

I I II I I

SERVO TABLE DYNAMICS

._O~~

_

Fig. 2.2. Functional block diagram of I.V. station keeping control loops

__'

I

Closed Loop Dynamic Testing TABLE 2.1

Attitude Control Concepts

AT'[T~~:';·: ~)L'f.

ort:HA1'10-

~~~f:

Ir :l~:

1:')I.L

~ :....

~RANSF't:R­

TOIRS, TOSSA

'

PRBIT

.'\CCE~,I:Rt)~:!';n_:,;.

I,

c-.)~r~·i'<';L C'::~;-

X:":'!'L>

J.. . '.. '-,

1r-5_ _h·'-.a.....,....__,·'-' ....··~~~~~~·I~

TABLE 2.1

CLPT 1 "TC:LS,\T 'J ~~~:",ol

i6CO_...,.,.... c;t-.;-.::-l~~;:,"::on

;- St'1n .

(It..>~

45 q;r:d

-

..

1- Act :VC T..:tat.io'1 damping

I.>

'; 'TOfRS, 1'(-551\ ACC2~EROr.!E!'t:i-:

l\rOGSE

1A~OEUVRr:.: ~

I

DESPHi

--..h----:

.i\TT!TUDE: c.o~nr.,J~

DURING

KE:O:PING)

i

r;CQUISIION AND 'ARTH 5EARCH

CASSA AAA

IC,\SSA ARA

--CASSA ARA

ARA

...

.~~~~_ ROL~ GEO-IRS 'ITI0N

CASS.~.

I I'

_ _ _ GEO-IRS ARJ\

'----l.. FLYWHEEL

~

RUN-UP

:

CAS SA ARA

I ~

FOSS ARA

~--~.--+-__-

(ARA)

_*_~_oQ~_U_IS_I_-___ ~()R:-~;·,I

~T

3-axis attitude stabi-

'~~oOJdeg.

(,,;1:;<)-1

I /.

'p""e,Uo"

l·ff!et of 12,Sdc proC:'.:c,::d by two succes'=' siv,::: yaw torque pulses i:t half nutation control) -Angl.llilr rr.o:'lcnLlI~ vecto in sl?a::e rn';l.lntained d'Jc co gyrosr;o;>ic

GEO-IRS

~ rotated)

s':if~ness

-Roll «ngle variation between' 12,5 deg over one-orbit (24 h) r~riod

-f'itch attitude control wit!;' flys~.ecl loop, sweepin.g by ~€";:s::Jr bias co;nma'lds fro:TI ground -F:e:Jcquisit.ion .-ifter 24 pet· iod

of nacmaI

i'lt"("~

_·ni~tl:')!.

..... ~r.::~.

£ ixt.::.j r-:),"":h-:n t'.1f:'" .....·~-.0,...: in tOl"q;lc :.;~dc ! It..:.lJ/ 1.3.'1 C0:-1:'lo..!I1S.~tio:l, r':0t .:ork an~ noise f'lt-"r! f'~tch

.:l.~C;L:lilr

.;W::~::t\l::-'

onboard l·:Jg:c 0':: c;rc:un cOIiJ',;and - Roll cor,to:,:< '::h:-.:stf'r.:> lncli~:'': ny :i,5" '.:,)

~~~d~~~p~;~~:: y ~~.:: t~~.

I

roll c~)-:~r('l tora'J~ - !
/'!1·':",C:..:

L: tcr

- ;~utatL:'n att~:;'j,,':ion

I

Sensor Equipment Characteristics

fer. j'd'N acquIsiti.on No valO attit'Jdc rc..fcrE:'~.c~ ::-0'-iuired

~

i

TABLE 3.1

l__·_--+_~_l~_~_e_(_~_~~_\_lb_~_'a_;._r_~_~_~_tp_l~_;_}-f.

r,~,

~.~~~~~~ n~r1~c ~~~~c as t

-----

_I

~011

-

- usbg station k~ef.in c;)ntrol loops (.::nly attitude re£. & lead lag cOlT'~ens.) - (or acqu~sit",on COIltrol lcvc's (<.tt .. t-u':;e and rate r.1caSUrC::lent - !"tatf~ r~gulilt')'·~))

d~l!'pin1 ~y ,'ll:C;·I.J;11('::~.

c·

__

~GEO-IR:.;

.....-d-u-l-a-tc-r.-------...

FDSS

-

A.<.I.S

-

'iorr.::l.i. :nodc clttir.~Ge .- cont!:"oJ conceot in :'l.Lch <.::0 r;oli/y.J,.... ~'ith GE()-Il-:S bias c':n!i'.,:1d L1 ;... it~·h ar.d roll

earth crossings Attit'lce d"r1 rat.E: {.:3ta te regu12tor) fee::,rnr:k. for con~rol -Yaw -isun:.)-reTe;:u,c~from FC-SS f
(ARA)

,

~ ,;:::0- tr~s

I

%

:

ac~ion jets

I

I

r-:ODL

i

G!':O-J~S

iTl.J.t.c:'

--I ....----~---_+_------4-:---__4~.-----_.-

;~~~~i~~f~~~~c:U~- r:-

. GEC-I'S [0-"5

ON-,-JHi\

I\NGLE OFFSET POINTING

~

~ GEO-IRS

Rc,ll ri,:c cur:-.n:ar.d ::-'~-

I

~

;,

i

t.O

11

GEO-IRS/ GEO-IR2

(ARA)

YAW

,

~~~~~r~o a~~;~o,~~r;~r

ARA

GEO-IRS ARA

~U'-"'5/clXi. .5)

I S~':ALL

"r.

i'<:'rfo~ ~<,nc'=! to ra:1ge of d.i.st. torque !.c=vels Off-::\od'.llation of att..i. t'.l':;e control thrus teTs i;; E/W primiiry 3. ~/S b?cl<-up hloue - ~;tal.Jilization nor",.Jl .:>anel ber-ditlc os.::il:atious via (fast) ya'~' cor,trol loop - S\-I.i tchi.ng o~ ninirnu;n P.·/f'F modulator pulse bi t by grour.d CO:~""C.iHd (transttion to nor:'1a)i

i,

< )O~/sec

'J-

\; -..lr:;~_....

? ~ tCfl I:IO;';':;~. tu:n ·..·~.:::e 1. sneed none - 'l'i:rt>e axes atti:".lce ref-=rence (eartr., ~un) - Lead/lao COr.1DE:'nSi'u::"on d~ild zo;'e and P\';PF ft0d .... latol' for cO:1trol - Gc.il~ s'"itclling \) V:'\-

-

t

- Sun scarch ';:)j roll ratE' (lv '" Or5°/~,ec) x Attit'Jde and rate f-2ed Di\ck 100ps (state r2Sul?tor) for sun capture Fully auto~la\:ic :":lode by onocard locfJ c and CASSA (:,Hth pr'.'s0n.ce signals Earth :;r>arcr. by rotatio:-: 3,!:;o'.lt x-axis ( l"J x '" 0,5 deg. / sec. )

CASSA

.~RA

• ARA

_ PITCH

CAS SA ARA

FDSS

e

----------

~_ _ '

~UN

GE()- I RS

~

~~~~,~~;;O'

Yaw thruster operaticI'. by ground comrnc:nd rate

., . . • • , ,

1',:

...+.__...

....c..

~

!

thr'.lster~

Ra:.€ fe·~cbac:.c leaps with prOj.OJ:ti-:>nal ::·.:mt:-ol 3.:1'':;' ;":'i~F t:'.Jd'J':'ators to ·,)·)crate reacton jCLS Dead zone to im')!:ove nei:'".€' (\('rfQrman~'=!

~GEO-IRS

ORBIT

~ ~::::~:":fc:::,::m, ;r~i~~:;

Continued

!l-'l,.~;_...U1;.:•....

Spin a:
ARA

393

r

GfO-IRS

OLL:+i4° (gcostat. L-.- PITCH: ';~'j0

j : , . ..,. ....... r 1

coOn",,,,,

I

x: 2 , 2 0/ () , 5 0 12 ..,. ')

I

"

')

c / ~

?.

:!

~!J, 01

0,I

deg

H. Bittner et al.

394

;;':~'2a IHT!LSAT I~

V

A 'I E U V E R

CTil- .. /5

4/5- ...

T::5T NU:-IBER

LOOL SUN ?OS I

Nf.s eT... •

coa.a. I '+IS

TABLE 3.4.1

'115- 5

Jilt· 00

T!:'E

TlO'1

I

Tabl.3.3.1 .l/;j-COE:R

A8.00

t-i;HJrlcf-L... rNr.-+-----::-.::'O'"--+--~-+-----If_--+-­ 20-

o·

(1')

INITIAL RHES

t'"

6.)xo

.~o ~~~

'r x ~;lIllJ

r.ONT!lOL

~y ~~:~ ~Y

HO~~HTS

~; &::;~

I~

/t.O 0.811

0.0 0.0 0.0

0.0 0.0 0,0

0.0 0.0 0.0

0·0 0.0 0.0

0.~29

TORQUES

HITERTU

1.75 0.911

Dtga J

INITIAL VP"'l fL VIIHF.EL SPl:.:EO

0, ~2.S

~'.m

i·m

2747 1360 2154

"'3'. Z-i SIt

3500

3.500

2.~Cf.l

F)ISTUR BA~/CE

1.

TOROUE PROfILP. NO .sELECTED THRUSTER SVSTEM

nPERATI:lG COI~POrJF:rlT"

Initial Conditions and Parameters for North/South Orbit Correction Manoeuvre

B

ADCE GEO- IRS fDSS

~

FMWA

..1,

A

1

2

oi ;t,

Af'4BIEt.ll

TEMPt:RATURE

"OT

~:.Itu~· '~~~~U8 ~W""P

~Q1'"

-----------=------,-

._------'":---=.-.-:-~~--

TABLE 3.4.2

Comm&nd Sequence and Action List of a N/S Orbit Correction Manoeuvre

MA N E U V E H ACTION

No

~lo

~

1

11

2

42

3

26

4

-

~

I

N/S-4/5- 4

PROPORTIONAL CMD LSD MSB

•

t

COMMAND SEQUENCe AND ACTION LIST Q

TELEMETRY DATA MONITORING O(OFF) L(ON)

00111

EXPLANATIONS / REMARKS

n::

w

>

~OCE 1 on; Automatic ADCE condi tioning rel ated to CMO 11 Sec Remark B of List 5.4

15/5

Sel ect Synchron Test Orbit Tr1 Data (DGT)

76 00010

fLl H

u... H

..

Synchron Orbi t Data switched to ana 109 TLM channel s. Further Preselection described in action No 25

13/2-3 15/3 6/4

Power on to GEO- IRS 1 by appropri ate push button at TCU

Word 1 t word 2. 47

I ~

CMD

:

5

08

-

Word 3. 13/8, 55

Reduce threshold for FOSS-Input intensity by push button at TCU

6

7

FOSS 2 on. (n1 monitori ng only if sun is in Fov)

15

10000

00000

75 62

73/74

GEO-IRS 1 Data sel ected for control (if necessary with respect to previuos state) GEO- IRS 1 Oa ta sel ected for Tr-l

8

16

00100

00000

9

19

00000

00000

10

20

00000

00000

Roll Bias Set to Zero

11

21

00000

00111

Yaw Bi as Set to Zero, ADCE 9a in for FOSS selected to t,15

Pitch Bias Set to Zero

13/4-6

Word 1 word 2

12 13

----- -

J

~O

Zero adjust of GEO-IRS output signals by FNS position biassing in pitch and roll see 6.4.1 Positionin~

- - --- -------- ----_._ ..

--

-

.--

and C

z

+

of sun simulator to 14.0a local time

00 (see 6.4.1)

i-"

395

Closed Loop Dynamic Testing TABLE 3.4.2

A

~i

Continued

E U V ER:

cm:~AllD SEQUENCE

I

N,IS-4/5- 4

AND ACTION LIST

Start recording Yiith scaling indicated in table 6.4.1.-1.

!

15 16

22

10100

10100

17

29

00001

00000

18

30

01010

10000

19

31

01010

10000

15/7

15/8

F~'WA 1

Word 4

16/1

Speed mode selected, cor.unanded wheel speed 3500 RP!':

67-70

8/4 8/6 8/8

3/3 8/5 8/7

10/4

10/3 10/5

10/6 10/8 20

32

00000

00000

21

33

00000

00000

on by push button at TCU

Safe mode selected, provision against incorrect GiD during store & execute system conditioning 26, 38, 6B enabled for automatic mode

4 ~ 5 B,

]iB

enabled for automatic mode

10/7

11/1-8 8/2 15/4

Di sab 1e of manua 1 thrus ters disable manual thrusters

12/1- 8

22

23

00000

10000

Word 7 + 8/1

23

24

00000

00000

Word 5 + 6/1

24

25

10000

00000

10/1

25

26

00010

00111

6/4 15/3 13/2-3

26

27

11111

11000

21-23

27

28

00001

01001

16/2 50, 52

28

40

29

44

Thruster pul se duration

j

=0 Word 9

See display

~

0

Number of thruster firings-+ 1 (each fi ring counted) Pulse Duration frequency 0.12 Hz ST/E-10gic (MD referenced FOSS gain ~ 1. 75 All thruster firing history possibil ities selected for counting

6/8 49, 51

For transient phase high gain in all axes preselected. Autom. ~~heel unload enable. ISS-logic enabled

66

30

Pul se Delay

13/1

Linear range of GEO-IRS-DAC low

TLM pattern

Check ADCS status according to

Check temperature of ADCS envoronment accord i ng to conditions of section 6.4.5

31

32

Initial ize Computer

33

Switch

34

TLi'l-display-patt~rn

N/S-4/5- 4

29

11010

00000

67,69/70 16/1

68

F~jS

position control to cOr.lputer

I\DCS swi tched to norma 1 mode Computer Start

35 36

35

37

29

Word 3 01101

00000

68,69 2/3 ,4/5 6/7 at rea 1 time testdisp.

38

36

15/1

39

39

24 4/5 at

67,iO 16/1

28

00000

01001

50, 16/2

41

28

00001

10001

FOSS-Reset Station keepi ng, control by thrusters 4/5, disturbance torques AS in norm. mode and in addition cosscoupl ing Torques (4/5 = on-modulated)

Manual thruster WO enabled Thruster pulse (ST/EXEC) start, start. disturbance torque profile No 1 applied, as long as man. thruster on. (Afterwards TO = Zero) (4/5 off-modulated during thruster pulse on time)

test rea 1 timedis play 40

Normal mode disturbance torques

49,51-52 6/8

About 30 sec after action 39: gain switched to med. ga i n in a 11 axes

51-52-

50,49

16/2

~/8

About 60 sec after action 39: gain switched to med. gain in roll & yaw, low gain in pitch

I

H. Bittner et aZ-

396

TABLE 3.4.2

I , ;

H

.~CTION tlo

Continued

ANE

:J

V EH:

PROPORT:ONAL CM!)

010 No

LSE

i

COMMAND SEQUENCE AND ACTION LIST :..... fJ-l

T::LE~-!F.:TRY

H

u...

DATA

MSB

EXPLANATIONS / REMARKS

H

MONITORING o(OFF) LeON)

t

t

j

I

I

N/S-4/5- 4

re: w

:>

4?

28

00001

00001

52 16/2

49-51 6/8

About 4.5 min after action 39: gain switched to low gain (large imp. bit) in all axes

43

28

00001

11001

50-52 - .16/2

49 6/8

About 5 min after action 39: modulators switched to s~all imp. bit (25 msec)

44

29

01011

00000

67.69 16/1

45

29

11010

00000

167

Transition to normal mode

68 .. 70

.69.7~

After dead beat impulses and at least 1 nutation period 'fJithout pulses-tswitch to normal mode

68

16/1

45

13/8 1317

8

~~ord

55_

47

29

43

34

oooeo

00001

FOSS-OFF (at rCU) demons tra tes actual contra 1 transfer

3

67-70 16/1-2

Safe mode CMD termi nates test run ISS-logic reset and disabled

66

TLM - SUBFRAME DATA BIT

Iwzlw3lw[.

Wl

W5!W5jW7

lwa

W9 IW 1olwnlw1zlw13

w"

W15 W15

BIT

! W~itnf~l: i ~1 j

TLM - MAINFRAME DATA

~~+ir) ~o \:J:' ~2 hl~+?~l¥-7:6-n:::5B ~::~9 ;{~f ~>r ~2 ~/6:3 BIT

I

65

!/}

I

66 167 168 169 170 17'

~

():

i(j: - ~::D:

NOMINAL V'cUE

RELf="VAt--JT

\:6 C{:

~

~:{::

0 I:: U V R E

73 I 7[.

~o

(>

*

•

Itv

I

78

1::;:::::

79

(::;:

L;t !

_§!.L

I

f.O i BIT

I::::::::

LOGICAL "ONE" ~ L.OG,.Ic.... L ,. ~ER.O

'I

ONLy

Re..sf'orvs.~ To

~ R ~ f" A R. \. t-.....J G..- Ac..T le /'v S

MANO~uVRE

77

k\

o

If'Jb\CATE:b PA'TEQ.~

!

i 75 I 76 ~

ACTU", V'LUE

STATV\S

TLK 1:>.AT,A »\SPLA'1'

M A-fv

'72

~ ~ : A - '2. ~

o ()

'--------------.,..,-,-----------_._--..-_.._-

Closed Loop Dynamic Testing

TABLE 4.1

Specifications of the Dynamic Motion Simulator

Fig. 4.1.

The dynamic Motion Simulator

397

H. Bittner

398

C~

av.

Fig. 4.2. Test room with Dynamic Bench Test Facility (DBTF)

.!-

--~

~~-==~

r~ r

,--

-

~_7--;-

~

--.

~

1"-

r-

t-

;..

~--_.

f ~.~. .-.~~~ . =.et> (OM. O.025°/L.

q>"CT 0.011

oIL

~Q> 0.00055 oIL

~ct>

4> ACT

ct>COM

0.011 oIL

0.015 oIL

0.00055 oIL cPCMAX:

Fig. 4.3. Dynamic Performance of U-shaped (middle)

frame

f

!O.lSo

= 0.1 H-z

Closed Loop Dynamic Testing

399

'.AXIS

••• -." ".. F"tftJCTION GEN AND EOUATIONS MOTION

PICK l,;PS

.

Ace .,ce

Of

V~_l_!-

GYROS

'lE L !'DS 0 _ 'lE L POS

ros

a

ANALOG C~TROEaUf""'ENT

1'AOLE

Sl:NSCRS

tl) A~PLlFIE~ FILTE~S

COMPE NSATION P~E T.... OrlK$ tIt

•

ACTUATORS 0,; ANLJ HyoraUlIC

CONTROL COMP.

TEMPERATlfflEl CHAMBERS

TCU

I

.!!U

Fig. 4.4

AND

STrltP CHAnT REGI5TPI.T1CN

BACKGROUND: Ft;NCTION

-:

II--~~·~

GENERATION

I

~}

M1ALOG AMPLlrrC:.TION

FOREGROUND: TAaLE CONTROl

'--_~"':C"':O';';N":"T":"RO;:;';L~

I

:=--_-~~~

~

{

---J

Block diagram of the I.V. dynamic bench test set-up

DISTURB. TORQUE

FWLL

TDX

DYNAMICS

Tex

M~GULAR

TDt

MOMENTUM COUPLING

YAW

DYNAMICS

DISTURB. TORQUE

Fig. 4.5

Simulation model of

sic

dynamics in station keeping control mode

H. Bittner 0t al.

400

.

~

,

~



t

t

-i--~ct>

Fig. 4.6

Check of simulation software

~i

I

~i

NI

;V I G I T A L SENSOR ELECTRONICS

1==·=R=~=':>:'=J=_t:lr=__=c. =-~=_)-E~: : :; 1---------.. U

'

- J-' -to I

~--...L..-l

t~

c) Orbit mode

b) Panel mode

a) Nutation mode

I

r-------<

I

·_~-ti 1--

t:tTt--\:~j

~.

-1-

I

)(-Ao.S~_

INFRARED SENSOR"

Fig. 4.7

Sensor configuration for station keeping manoeuvres and normal mode

~; N,

~

¥

f

..

,fc:::~il !

1;

/ANALOG SUN SENSOR

ELECTRONICS

I ~-------.

---~ - - -tl~ -- ----t -i ~-----_-I---I

~":': - _:;:;

/

)

_.I:;..

I

;

\._ \

-r__ .:1- &\-...L _LI!J-__\.Q,l"':L'~~~E_ I 11

I

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r~-~--Ll,

I:

r-

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\

\

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l[J I

I

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S~NSOR "

I

Fig. 4.8

I

illj

IL

;.

: dj'

_

Sensor configuration for acquisition manoeuvres

Closed Loop Dynamic Testing

401

~E'O

- /R.S

R.oL<- Sl <;'NA-L

0.00-:"5

Fig. 7.1.

N/S- Correction 4/5-4

YL

Roll axis, local time 14h

H. Bittner et al.

402

0.00-1. ef~

Fig. 7.2.

le....

N/S-Correction 4/5-4

o.oos Pitch axis, local time 14h

YL.

-.--...

--

.. ~ ... -

'::.-

-=_.

::--1

=

.-

:=-.--

-_.

--:::: .r:

.__

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.

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.-

-

c-' .:::

-_.~_.-

._- ..

,

=-.b- ..

.- =

=

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... ...;.

-_.-

-\.

{~II -.-

- --J _.'=-:.-

L

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,- ---

t

+:-+--

o

+:-+--

Results of Digital Data Evaluation

TABLE 7.1

NUMBE~ NUMBEk

MANEUVE.RINORTH/SOUTH-COHR MANEUVERINORTH/SOUTH-COHR CTH-4/S RuN NR. 17

DATt 30.03.7Q DATl 10.03.79

TIME

MANOEUVHE TIME TH-

PITCH

o

NUMBEH NUMBER OF THRUSlEH THRUStEH

OF THRUSlER THRUStER FIRIN~S FIRINGS ROLL 719

PITCH

lAW YAW 0

157

41b

0.000 175.29b 175.2 fHl

111.00 14.00

INITIAL POSITION(Xl,X2,Xl,Xll) POSITION(Xl,X2,XJ,XII) O.OOOE 00 O.OOOl OU O.OOOE 00 O.OOOl 00

0.000

MEAN PULSE DURATION 0.0000 0.2438 0.2436 0.0000 U.I00E. 01 O.IOOE.

fUll EFf IClENCY ICIENCY 100.0000

23.344 U.1487 0.1487

Hl 0.000 ON OFF ~ODULATI0N ~OOULATION Ul- 1.000 1.000Hl

FLY~HEEL(XQ,XI0,ICO,ICU,KW,'MY,lH,F)1 FLVWHEEL(X9,XIO,ICO,lCU,KW,tMY,lH,F)1

O.OOOE 00 0.500E-02 0.SOOE-02

-O.~OOE -O.qOOE O} 0.9S5E-Ol 0.955E-Ol

-O.lbbE -0.3bbE. 0] 0) 0.100E O.IOOE 00

-0.J30E. .O.l~OE. 03 O~ O.ooot O.OOOl 00

DISTURBANCE TORQUES(TTX,lTV,TTZ,T9X,TSV,lSI,lA)1 DistURBANCE TORQUES(TTX,lTY,TTI,T9X,TSV,T81,lA)1 0.522l-01 0.153E 0.290E 0.522£-01 0.151E 00 0.290E. 00 0.522E-Ol 0.lS3E 0.290E O,OOOE 0.522£-01 0.IS3E 00 0.290£ 00 O.OOOE 00 MOMENTS MOME.NTS OF INERTIA(IX,IY,IZ)1 INERTIA(lX,ly,lZ)1 O.27470E 0.11bOOE 0.21470E 04 0.13bOOE 04 SAMPLING TIMt TIM~ H.

0.21S'l0E. 0.21SqOE 04

0.0060

b3.728

0.1532

0.000 145.280 14S.280

0.000

CONTROL CONlROL TORQUES lORQUES 0.929

U.4ge 0.498

BEGIN BE.GIN OF Of LIFE MED.GAIN,LARGE MED.GAIN,LARGl IMPULSE BIT

2.';)bO 2.SbO

0.800

MEAN PULSE PULSE. DURATION 0.0000 0.2252 O.lbOO

0.0000

FUEL EFFICIENCY 100.0000 100.0000

100.0000

0.1547

0.000

52.784

20.272

0.1534

0.0000 100.0000

92.4010

OVERSHOOT PHI,THETA,PSI 0.0931 0.1159 0.1993

STEADY STlADY STATE STAlE ,VARIANCE,MAXIMUM ,VARIANCE.,MAXIMUM ,MINIMUM 0.0747 0.1133 0.1855 0.0147 0.0131 0.0097 0.0039 0.0991 0.12b8 0.21b3 0.0482 0.1021 0.1460 0.1027

STEADY STAT~ STATt ,VARIANCE,MAXIMUM ,VARJANCE,MAXIMUM ,MINIMUM 0.Ob25 0.1013 0.1593 0.0091 0.0040 0.0155 0.084b 0.1111 0.1940 0.0432 0.0910 O.llbl 0.11bl

HISTO~HAM HI3TOl;HAM

HISTOGRAM

::r: ::r:: tJj to .....

~.

rt rt

rt

0,0000 0.0000

O.ObOO

0.0000

0.0100 0.0700

0.0000

0.0200

::l ~

(0 et>

t'1 1'1

0.0000

0.0500

0.0000

O.ObOO

0.0000

(\)

~

0.0100

0.0000

0.0700

0.0000

0.0900

0.0000

0.0100 0.0.500

0.0000

O.ObOO

0.0000

0.0800

0.0000

0.0200

0.0000

0.0800

0.0000

0.1100

0.0000

0.0400

0.0841

0.0700

0.0000

0.1000

0.0049 0.00119

0.0300

0.0000

0.0900

0.0000

0.1300

0.00119 0.0049

0.0500 O.OSOO

0.3bOb 0.3606

0.0800

0.0000

0.1200

0.0922

0.0400 O.OqOO

0.0058

0.1000

0.1918

0.1500

0.1158

O.ObOO

0,3192 0.3192

0.0900

0.3574

0.1400

0.4331l 0.4.BII

0.0500

0.05bl

0.1100 0.11 00

0.7b67

0.1700

0.5117

0.0700

0.2028 0.2026

0.1000

0.b342 0.b3112

O.lbOO

0.3603 0.3b03

O.ObOO

0.22l1 0.22'11

0.1200

0.039b 0.039b

0.1900

0.3381 0.:B81

0.0800

0.0:B3 0.03.B

0.1100

0.0084

0.1800

0.1093

0.0700

0.428b

O.13UO 0.1300

0.0000

0.2100

0.0295

0.0900

0.0000

0.1200

0.0000

0.2000

0.0000

0.0000

0.1000

0.0000

0.1300

0.0000

0.2200

0.0000

0.2400

0.0000

0,2203 0.1400 0.0000 0.2203 0.11100 SYSTEM 0.0800 SYSTfM A OR B 1 I NMOD 0 KRY. 52500TSK= 1.0001EOlC 0.9bOTBOl= 525QOTSK= 1.000lEOle 0.9bOTBOL= 2.8201M2.820lM- 1.200 D~ 0.Ob20 0.1500 0.0000 END lIME 1211009 0.0900 TIME IA(I)25 IA(I)=5 1271009 STATION KlEPING STATION KEf.PING TBtGIN= 1123.11[NUs 843.10 (SEC) TBEGIN: 423.IT~NU~ l~lGIN. 3.91ENOa l~~GINs 3.9lENU- 1123.10 423.10 (SlC) (S~C) LONG TIME liME. RUN DISTURBANCE TORUUlS lTXa TTYD TORQU~S lTX- O.Oblb 0.Ob2b TTYm 0.1841 COEF OF -0.200 COEf OF MULTIPLICATION 1.20

l44 344

131

OVERSHOOl PHI,THlTA,PSI PHI,TH~TA,PSI 0.0529 0.1531 0.1718

0.0000

CON1ROL CONlROL TORQUES(SMX,SMYSMZ)I 0.1I17E 0.2UO£ 0.103~ O.417E 00 0.200E 00 0.103~ 01

5

~lwI~bS ~lwl~~8

°

INITIAL RATE(XII,X5,Xb RATE(X4,X5,Xb (RAD/SlC»1 (RAO/SEC»I O.QOot O.OOOE 00 O.OOUE OU O.OOO~ 00 O.OOOE 00 PANEL STATE + PARAMETERS (X7,X8,GO,AN,BN,DN,EN)1 (X7,X8,GO,AN,BN,ON,EN)1 O.OOOE 00 0.000£ 0.100E-0~ O.OOOE 00 0.100E-03 0.170E 0.15bE 0.14bE 00 0.78QE. O.IIOE 02 0.15bE. 01 0] 0.78GE 02

YA" YAw

NUMBER OF PULSES PEH PEW MINUTE MINU1E o 92 0 18 49 0 TOTAL GAS CONSUMPTION ~OR lACH EACH AXIS

NUMBER OF PULSES PEH PER MINUTE o 119 00 2b b9 00 TOTAL GAS CONSUMPTION FOR fOR EACH AXIS

14119141 Iq,19141

o

ROLL ROll b45

0.2300 0.2500

0.0000

LON~ TIME RUN LON(; TTl 0.3481UlSTURBANCl Tlx: 0.0522 TTY= 0.3q81UlSTURBANtE. TOHUUES TOHYUES TTY: 0.1534 COlF -0.200 COlf OF Of MULTIPLICATION MUL1IPLICATION 1.00

CONTROL TOHUUlS TOHQUlS 0.929

0.498 0.4198

BEGIN OF OF lifE MED.GAIN,LARGl IMPULSE BIT

2.5bO

0.1100

Tll T1I O.ZQOl 0.2QOl

0.0000

0.1400

0.0000

ct~ ~