SO96.4.06
ISO MISSION CONTROL

Juergen Faelker

esa villafranca del castillo satellite tracking station (iso/scc),
p.o. box 50727, 28080 madrid/spain,
e-mail: jfaelker@vmprofs.vilspa.esa.es; fax: +34-1-8131139

ABSTRACT. ESA's Infrared Space Observatory (ISO) is the first cryogenically cooled astronomical observatory and as such presents specific mission control requirements for spacecraft subsystems and scientific instruments. The 18 months cryogen limited lifetime (now predicted for 24 months) and an oversubscription for astronomical observations demands rapid but safe execution of satellite and instrument operations. Telescope pointing requirements, higher than specified, instrument detectors whose performance is highly variable and susceptible to memory effects, short lifetime and satellite susceptibility to damage present conflicting requirements on the mission control system and complicate the data processing of the observations. These are described and lessons drawn which will be useful for the design of future "extreme" missions.

                                                

1. MISSION OVERVIEW

ISO is an European Space Agency (ESA) scientific mission that performs observations of selected celestial sources at infrared (IR) wavelength from 2.3 to 200 micro-meter with the highest possible sensitivity and spatial resolution. The high-technology telescope facility has been designed and built in Europe for use by the scientific community in Europe, Japan and the United States of America. The ISO spacecraft was launched on 17 November, 1995 at 01:20 UT from Kourou as a single payload on an ARIANE 44P. The spacecraft was placed in a highly eccentric, geosynchronous transfer orbit with a 71,577 Km apogee, a 500 Km perigee, and an inclination of 5.25 degrees. Two days within the mission, the perigee was raised to 1030 Km by use of hydrazine thrusters to achieve a quasi operational orbit. Finally, the apogee was lowered to 70,611 KM on 24 November and the operational orbit was achieved.

Figure-1. The ISO Satellite

The ISO satellite (see Figure-1) is a three-axes stabilised spacecraft and consists of a service module (SVM) and of a payload module (PLM). The SVM carries the conventional subsystems required to support the mission objectives with respect of attitude control (AOCS) and measurement during different modes of operation, manoeuvre capability by use of the hydrazine system (RCS), power distribution (PCS) from solar arrays mounted in front of a sunshield and two Ni-Cd batteries to support eclipse phases. Furthermore, radio frequency (RF), on-board data handling (OBDH) and a simple and passive thermal control (THC) subsystem are provided. The PLM consists of a huge cryostat vessel and associated tank (CRYO) that was filled at launch with 2300 litres liquid Helium-II, the associated electronics and valves and a jettisonable cryostat cover.

Inside of the cryostat vessel the telescope, type Ritchey-Chretien, and the four " cold " focal plane units of the scientific instruments are located underneath of the 60 cm primary mirror. They operate at around 3 K. The extremely low temperature is achieved by boiling off the liquid Helium. The instruments were conceived and built by various scientific institutions external to ESA. The four instruments carried on- board ISO are: The Infrared Camera (CAM), the Long Wavelength Spectrometer (LWS), the Photo- Polarimeter with imaging capability (PHT) and the Short Wavelength Spectrometer (SWS). The cryostat cover was jettisoned on 27 November and from this moment onwards, the scientific instruments started to receive infrared radiation from space. The Helium bath achieved thermal equilibrium at 1.724 K some three weeks after launch.

The ISO satellite has no on-board data storage. Therefore, to carry out any scientific observations, continuous real-time contact with a ground station is essential. In the absence of ground station contact, the
satellite has a variety of safe and autonomous modes enabling it to survive up to 72 hours until contact is re-established.


2. MISSION PHASES

The ISO Mission Phases are described briefly in the next paragraphs, since they play an important role with respect to mission control aspects. The Flight Control Team (FCT) is responsible for the conduct and control of the flight operations for ISO. Overall spacecraft security and safety including the scientific instruments rests with the Directorate of Operations (D/OPS) ESOC, while the operations of the scientific instruments rests with the Directorate of Science (D/SCI). The core FCT team consists of 6 Spacecraft Engineers, including 2 Mission Planning Engineers and 2 Spacecraft Analysts. They were supported initially by 3 Spacecraft Controllers. The core team was responsible for the bulk of the work involved in the preparation of the flight control operations. Some seven months before launch the complement of Spacecraft Controllers was increased to 12, required to support ISO operations 24 hours a day 7 days a week.

2.1 LAUNCH AND EARLY ORBIT PHASE

During Launch and Early Orbit Phase (LEOP), ISO was controlled from the Operations Control Centre (OCC) of the European Space Operations Centre (ESOC) in Darmstadt, Germany. The ESA ground stations Perth, Kourou and Villafranca supported LEOP. The principal LEOP activities were to undertake initial spacecraft acquisition operations, to perform initial subsystem check-out, to achieve a quasi operational orbit and to make the spacecraft ready for handover of operations t o Villafranca.

The LEOP activities were conducted according to the Flight Operations Plan (FOP). Beside a temperature
anomaly on the cooling of the CCD of Star Tracker-1 (STR), which was resolved shortly after, all LEOP activities were executed as planned. After three days, the first part of the Flight Control Team moved to the ESA Villafranca Satellite Tracking Station (VILSPA), near Madrid/Spain, where the ISO dedicated Spacecraft Control Centre (SCC) and the Science Operations Centre (SOC) are co-located. During LEOP, the SCC and SOC were in the so called "Listen- In" mode, i.e. only telemetry was passed on in parallel to the OCC. No instrument operations were conducted during LEOP. Handover of operations from the OCC to the SCC took place as planned on 21 November, i.e. with the beginning of revolution 4, which concluded LEOP.


2.2 SATELLITE COMMISSIONING PHASE

As of the beginning of the Satellite Commissioning Phase (SCP), ISO is controlled from the dedicated SCC and SOC facilities at Vilspa. The major SCP activities were to commission spacecraft functions and the scientific instruments, to achieve the operational orbit, to jettison the cryostat cover, to make the satellite ready for scientific observations and to verify the overall observatory functionality. As of this phase two ground stations, Villafranca and the NASA/JPL Goldstone Station, are supporting ISO for some 22 hours per revolution. The SCP was carried out according to the Satellite Commissioning Plan, and where applicable, according to the FOP, between 21 November (Rev- 4) and 9 December (Rev-21) 1995.

The SCP showed that the overall spacecraft status is excellent and that all subsystems performing within or above specifications. Although the cryo cover was not released yet, the electrical switch-on phase and check out of the scientific instruments commenced. On 24 November, close to perigee 7, the apogee lowering manoeuvre was performed and ISO reached its mission orbit, as shown in Figure-2 . The orbit determination result on 24 November, include the effect of the 39 min. 20.5 sec. Delta-V manoeuvre, executed at 02:50:04 UT prior to perigee 7. Epoch 24 November at 03:45:29 UT (perigee 7) . The major event during this phase was the cryostat cover release on 27 November at 10:27 UT. Verification of the jettisoning was immediately seen in telemetry by monitoring the gyroscope's outputs. Shortly after, the temperatures of the cryo subsystem indicated that ISO was indeed viewing cold space. The cool-down


Height of perigee 1004 Km
Height of apogee 70611.1 Km
Semi-major axis 42374.6 Km
Eccentricity 0.825788
Inclination 5.1975 Deg.
Ascending node 308.7 Deg.
Argument of perigee 115.78 Deg.
True anomaly 359.98 Deg.

Figure-2. ISO Mission Orbital Elements

of the cryogenics was well in line with the predictions of the thermal model and equilibrium was reached some three weeks after launch. The Helium-II bath temperature settled at 1.724 K (spec: 1.7 to 1.9 K) and the temperature about the Optical Support Structure (OSS), was measured around 2.8 K (spec: 2.4 to 3.4 K). All nominal functional modes of the SVM have been successfully verified. After cryo-cover ejection the Quadrant Star Sensor (QSS) was switched on and calibrated and the first QSS/STR misalignment calibration performed to measure the offset between the telescope boresight and the Star Tracker (STR). A misalignment of 205 arcsec was measured, which was well within specifications. Since then, the QSS/STR misalignment calibration is performed once per
revolution to determine variations due to thermo-elastic effects. The misalignment measured is then applied automatically to correct the attitude quaternions of the pointing requests.

The pointing performance of the Attitude and Orbit Control Subsystem (AOCS), as shown in Figure-3, is considerably better than specifications. Similar to the SVM, all nominal modes of the PLM have been successfully verified. The Helium-II flow rate is well within limits leading to an anticipated lifetime of 24 +/- 2 months, compared with the baseline of 18 months. Direct Liquid Content Measurements (DLCM) are required to determine the remaining He-II mass and hence, the lifetime.

Figure-3. ISO Pointing Performance

UNITS IN-ORBIT SPEC.
Relative Pointing Error (RPE) , 2 sigma-half cone arcsec 0.5 < 2.7
Absolute Pointing Drift (APD) arcsec/h < 0.1 < 2.8
Absolute Pointing Error (APE) arcsec < 4.0 < 11.7

The Focal Plane Geometry Calibration of the four instruments occupied a large part of the Satellite Commissioning Phase (8 revolutions), because the procedures used to acquire up to 5 stars for the PHT
and the SWS experiments with the Star Tracker using the restricted search and tracking mode, needed to be interleaved manually with an executing Central Command Schedule (CCS), whereby all groups (SOC/SCC/Instrument Dedicated Teams (IDT)/PROJECT and FLIGHT DYNAMICS) were required to work very closely together. Figure-4 shows the final offsets determined for each of the instruments and their relevant apertures.

INSTRUMENT UNITS Y-AXIS Z-AXIS
CAM arcsec 2.65 7.45
LWS " 13.5 4.5
PHT-1 " -9.5 -3.9
PHT-2 " -9.5 -3.9
PHT-3 " -6.8 -5.1
SWS-1 " 1.5 -19.6
SWS-2 " 4.7 -22.2
SWS-3 " 9.3 -24.5

Figure-4. ISO Focal Plane Geometry Calibration Offsets

2.3 PERFORMANCE VERIFICATION PHASE

The Performance Verification Phase (PV) was carried out in accordance with the Scientific Performance Verification Operations Plan between the 10th of December 1995 (Rev-22) and the 3rd of February 1996 (Rev-78). The principal milestones of this phase were to perform the core scientific calibrations, to validate the instrument observing modes (AOT's) and to verify the Observatory functions. The execution of the PV revolutions was planned in an automated way, using the Central Command Schedule (CCS) and one instrument per revolution only. However, due to additional requirements the CAM instrument was used in Parallel Mode during the majority of the 56 revolutions. The PV phase showed that the performance of the instruments is very good. The instruments are functioning very well and no problems in functionality of the mechanical "cold" units have been observed. The sensitivity of the instruments is affected by 'glitches', caused by high-energy cosmic-ray particles impacting on the instrument detectors. As a result of in-orbit performance tests, the originally 16 hours science window has been extended to almost 17 hours per revolution.

2.4 ROUTINE MISSION PHASE

The Routine Mission Phase (RMP) of ISO commenced on 4 February 1996 (Rev-79) and will continue until depletion of the liquid Helium. Throughout RMP planned operations will be conducted per Central Command Schedules (CCS's) , the vital products of the ISO Mission Planning System. The latter is presented in SO96.3.05. The aim of the RMP is to make optimum use of the available science observations time and to undertake the desired observations programme . To date some 33.000 observations from more than 1000 accepted proposals awaiting execution. Due to the very good performance of the space and of the ground segment and the pace of the observations, it is believed that some 22.000 observations will be conducted within the baseline lifetime of 18 months. For the anticipated mission extension beyond the 18 months, a second call for proposals will be made in the near future.

Baseline during Routine Mission Phase is that all 4 instruments will be activated and deactivated per schedule, disregarding that a particular instrument may not be used to observe during a particular revolution. Throughout all mission phases the relevant Flight Control Procedures (FCP's) and Contingency Recovery Procedures (CRP's) of the Flight Operations Plan (FOP) are applicable. It is noteworthy to mention that approximately 1000 FCP's and 500 CRP's have been written and validated before launch by the FCT Engineers, covering all spacecraft subsystems and as well system level aspects. The bulk of the procedures were validated with the platform simulator . Additionally, and under the responsibility of the SOC, more than 130 instrument control procedures and contingency recovery procedures have been written and validated with the instrument simulator. These are documented in the applicable Scientific Instrument Flight Operations Plan (IFOP). A considerable amount of these procedures was (re-) validated before launch during System Validation Tests and during End-to-End Tests with the ISO Flight Model .

3. MISSION CONTROL SYSTEM

The ISO Mission Control System (see Figure-5) performs all aspects connected with the operations and safety of the spacecraft, including safety monitoring of the scientific instruments. The hardware of the control system consists essentially of two VAX 4000-600 redundant Spacecraft Monitoring and Control computers (ISORT/ISODV), five associated SUN SPARC-20 workstations), its associated spacecraft control software, and the mission planning system software as far as Mission Planning Phase 2 (MPP2) is concerned. The system is designated as the ISO Dedicated Control System (IDCS). The Flight Dynamics System (FDS) consists of a set of five SUN workstations and its dedicated software. These systems are network-wise interfaced on a partially redundant OPSLAN to prevent single point failures and isolated against the outside world. Two redundant micro-Vax 3100-76 computers form the Operational
Figure-5. ISO Mission Control System Block Diagram

Data Server system (ODS-1/2). The ODS constitutes the interface between the spacecraft control system of the SCC and that of the SOC as far as science real-time data reception in form of Telemetry Distribution Formats (TDF) is concerned. The latter contains as well telecommand history data and specially provided derived telemetry parameters for instrument monitoring and control purposes required by the Real-Time Analysis (RTA) and Quick-Look Analysis (QLA) run on the four instrument workstations (one dedicated per instrument) of the SOC. The ODS as well provides the interface between the Mission Planning Phase- 1 (MMP1) of the SOC and that of the SCC (MPP2).

Furthermore, the ODS provides the short history archive of the science telemetry, which is accessible from the SOC Science Data Processing system. The NETWORK INTERFACE provides the connectivity of the IDCS with the ground stations through the Integrated Switching System (ISS), as part of the OPSNET. Support functions are provided for: Spacecraft Performance Evaluation (SPEVAL) required to determine all aspects of spacecraft performance which can impact the life of the mission and mission efficiency. Spacecraft on-board software maintenance for the AOCS, STR and the OBDH. Communications Services are provided to interface with the Ground Stations, and with ESOC for ranging and orbit-related activities. Two Spacecraft Hybrid Simulators are provided to support a variety of tasks such as procedure testing and validation, AOCS on-board software maintenance and validation, and spacecraft anomaly investigation.


4. MISSION CONTROL CONSTRAINTS

There are several spacecraft constraints that have to be adhered to in ISO mission control. Violation of any of the constraints indicated below will have serious consequences on the mission itself and under certain circumstances may terminate the ISO mission.

The solar aspect angle of the telescope boresight (X-axis) must be kept in the range 60 to 120 degrees. The earth constraint is determined by the opening angle of the sunshade (14 degrees). To avoid spectral reflectance of the earth by the sunshade into the cryogenically cooled PLM, the earth must remain outside a half cone top angle of 77 degrees centred about the X-axis. A 1 degree margin for safety is included in this figure. This functional constraint has been expanded to define two separate viewing regions for AOCS
control purposes, namely an earth warning region and an earth forbidden region. The gap between the two
regions is maintained to be 10 degrees. The SVM must support eclipses of 80 minutes durations. This is achieved by two 24 Ampere-Hour Ni-Cd batteries. By complying with the above constraints, the primary ISO mission requirements to maintain the low temperatures to the instruments and optics during a minimum lifetime of 18 months, and to point the telescope and hence, the scientific instruments towards selected celestial sources during a 10 hours maximum observation period, can be accomplished.

All scientific observations must satisfy the spacecraft constraints. Additionally, there are constraints in the use of the instruments and of the Star Tracker (STR). Due to the highly eccentric orbit with a perigee of 1000 Km, the satellite passes the outer and inner radiation belts of the earth, requiring the instruments to be switched on/off once per revolution, while the STR cannot be used +/- 3 hours around perigee. The instrument constraints are based on satellite altitude and are separated in the electrical use (26.500 Km leaving perigee to 21.200 Km towards perigee entry). Furthermore, scientific observations can be performed only within 43.200 Km leaving perigee and 37.250 Km prior to perigee entry, some 17 hours per revolution (orbit).

4.1 CENTRAL COMMAND SCHEDULE OVERVIEW

The Central Command Schedule (CCS) reflects the output of the ISO Mission Planning activities and contains virtually all commands and therefore operations of the spacecraft and instruments are performed
with minimum operator intervention. Only in this way can the science return from this short duration mission be maximised. A CCS overview of a skeleton revolution is shown in Figure-6. The CCS contains dedicated WINDOWS during which either spacecraft operations or science operations can be scheduled. Additionally, EVENT DESIGNATORS and KEYWORDS are defined which trigger certain command operations to be inserted in those windows, when required. In order to make maximum time available for the scientific observations, spacecraft activities and instrument activation and de-activation activities are placed along an orbit in such a way that they do not consume science time. Interleaved manual commanding from the stack is in principle only required to support Ranging and a few specific AOCS operations. Recovery from anomalies, either spacecraft, instruments or ground segment, the schedule offers HOLD, RESUME and SHIFT functions to minimize loss of observations.

Figure-6. ISO Central Command Schedule Overview

TIME
[relative] 		EVENT ACTIVITY
00:00:00 PSF_START Schedule start time; equals to perigee crossing time.
+ 00:15:00 AOS_TM Acquisition Of Signal - Villafranca Station. (+/- Y Antenna)
+ 00:15:00 AOS_CHK_OPEN S/C status check and re-configuration from perigee passage.
+ 00:35:00 AOS_CHK_CLOSE
+ 00:40:00 MOUT Message OUT: Next on-board antenna (+/- Y) change.
+ 01:00:00 PPL_LOAD_OPEN Pre-planned Pointing List uplink for 72 hours safe pointing;
+ 01:10:00 PPL_LOAD_CLOSE TT-Cmds for PPM entry and Instruments off.
+ 01:55:00 ACTIV_OPEN Activation window - 4 Instruments switch on and activation
+ 03:20:00 ACTIV_CLOSE begins at altitude of 26.500 Km.
+ 03:20:00 SOPS_OPEN Reaction Wheel Biasing for science window to prevent
+ 03:25:00 SOPS_CLOSE autonomous unloading and wheel zero-speed crossing.
+ 03:25:00 ACAL_OPEN Calibration window - Attitude update, transition PPM to
+ 04:00:00 ACAL_CLOSE Fine Pointing Mode (FPM), QSS/STR calibration.
+ 04:00:00 OBS_OPEN Observation window - Scientific observations commence.
+ 10:18:00 AOS_TM Acquisition Of Signal - Goldstone Station.
+ 12:46:00 OBS_CLOSE Mid-Orbit Calibration LWS/PHT/CAM (-17min) continues.
+ 12:46:00 HAND_OPEN Handover window - TM/TC from Villafranca to Goldstone
+ 12:56:00 HAND_CLOSE Mid-Orbit calibration LWS/PHT/CAM (+9 min) continues.
+ 12:56:00 SOPS_OPEN Reaction Wheel Biasing (as above)
+ 13:01:00 SOPS_CLOSE
+ 13:01:00 OBS_OPEN Mid-Orbit calibration LWS/PHT/CAM (+4 min.) continues.
+ 13:28:00 LOS_TM Loss Of Signal - Villafranca Station

Figure-6. ISO Central Command Schedule Overview (cont'd.)

TIME
[relative] 		EVENT ACTIVITY
+ 20:55:00 OBS_CLOSE Observations must stop at altitude 37.200 Km + PPM entry.
+ 20:56:00 STR_STOP At - 3 hours prior perigee crossing time.
+ 20:56:00 PPL_LOAD_OPEN New Pre-planned Pointing List uplink for 72 hours safety.
+ 21:06:00 PPL_LOAD_CLOSE
+ 21:06:00 SOPS_OPEN Reaction Wheel Biasing to avoid autonomous unloading
+ 21:11:00 SOPS_CLOSE during perigee passage - affecting orbit.
+ 21:11:00 DEACTIV_OPEN Deactivation window - 4 instruments deactivation and
+ 22:16:00 DEACTIV_CLOSE switch off for perigee passage.
+ 22:16:00 LOS_CHK_OPEN S/C status check and reconfiguration for perigee passage;
+ 22:31:00 LOS_CHK_CLOSE Instruments switch off (backup) at + 22:27:00 (21.000 Km)
+ 22:33:00 MOUT Switch S-Band antenna (+/- Y) required for next AOS+EOT
+ 22:46:00 PPM_ENTRY Latest possible time for PPM entry.
+ 23:46:00 LOS_TM Loss Of Signal - Goldstone Station.
+ 23:56:00 PSF_END Schedule end time; equals to perigee crossing time.

5. CONCLUSIONS AND LESSONS DRAWN

This paper has summarised the mission control for the very complex ISO mission with emphasis on those aspects and elements affecting the specific mission control requirements for spacecraft subsystems and scientific instruments. The careful preparation done by all parties has resulted in a very efficient and successful process to control and conduct the ISO mission, despite the complexities of the space and of the ground segment, which form the ISO Observatory. The fact that the Spacecraft Control Centre, the Science Operations Centre and one Ground Station are co-located in the Villafranca complex, contribute definitively to this success.

Future observatory type missions should adapt the ISO experience as a general guideline, although some areas have been identified, where improvements or changes would result in a more effective and cost efficient approach. The ground driven command schedule is highly susceptible to ground segment problems. This should be an on-board application task. On-board mass memory storage is required to reduce the costs for a second ground station, required to provide nearly 24 hours real-time support. The splitting of responsibilities between instrument operations (SOC) and spacecraft operations (SCC) is inefficient and a cost-driver. Instruments loosing their RAM patches after switch-off and hence, requiring the uplink of the said RAM patches once per revolution, is an inefficient and time consuming approach. The activation and deactivation process should be an on-board application task. The latter introduces a heavy load on the commanding system. The mission planning system (MPP1 and MPP2) requires streamlining. Too many parties are involved in the planning process, which makes the process slow and in some areas inflexible and hence, not efficient, although it works surprisingly well.