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6         FUSE Instrument Technical Appendix

6.1           Telescope Mirrors


6.1.1   Telescope Mirror Design


The four telescope primary mirrors were identical off-axis paraboloids, each with a rectangular 352 μm x 387 μm clear aperture, a 2245-μm focal length, and approximately 5.5 off-axis angle (Kennedy et al. 1996).


The primary mirror sections did not contain the vertex of the parent paraboloid.  The vertex was about 34.5 μm from the edge of the mirror aperture in x (See Figure 6-32). The different grating incidence angles required to optimize the spectrograph channels for the two wavelength bands resulted in slightly different off-axis angles for the SiC and Al+LiF coated mirrors. The  off-axis angles were defined by aperture stops placed over the surfaces of the mirrors.  Otherwise, the mirrors are identical except for the coatings.  The corners of each mirror were masked to match the grating apertures, whose outside corners were removed to satisfy space constraints of the Med-Lite launch vehicle fairing.  The resulting geometric area of each mirror was approximately 1330 cm2. These apertures were widely separated on the instrument optical bench, resulting in four parallel and separated optical axes. Selected mirror assembly specifications are listed in Table 6.1‑1. The point spread function (PSF) at the focal plane places ~90% of the light within a diameter of 1.5 arcsec.


Mirror type Off axis parabola
Substrate material Zerodur
Size of clear aperture 387′ 352 μm

Focal length 2245 μm
Off axis angle (to optical center) 5.3668 (SiC mirrors)
5.4678 (LiF mirrors)
Coatings 2 mirrors with SiC
2 mirrors with Al+LiF

Table 6.1‑1: FUSE Mirror Properties


The mirrors were fabricated from Zerodur, chosen for its low coefficient of thermal expansion (CTE). The blanks were aggressively weight-relieved: 70% of the substrate material was removed from each, leaving a triangular isogrid rib pattern with a 7.5mm-thick facesheet and a final mass of 7.7 kg (Figure 6‑1). This triangular rib structure provided a lightweight but very stiff substrate.  SVG Tinsley Laboratories lightweighted the blanks and figured the mirrors into parabolas.


Figure 6‑1: Left: FUSE mirror resting face-up on flexures prior to integration into the mirror assembly. Right: Backside of the FUSE primary mirror illustrating the aggressive lightweighting of the Zerodur mirror substrate.

Each mirror is attached to the front of a honeycomb-sandwich intermediate plate by means of tangential-blade flexures, which minimize mounting-induced stresses and distortions and isolates the mirror from stress induced by adjustments of the actuators (Nikulla, 1997). The flexures are oriented with soft axes radial to the center of the mirror (Figure 6‑1 Right), and consist of titanium alloy blades attached to low-CTE Invar fittings. Each Invar fitting is bonded to a mirror rib. Three stepper-motor actuators are attached to the rear of the intermediate plate.  This permitted independent tip, tilt, and focus control for on-orbit adjustment of each mirror (Figure 6‑2). The tip-tilt mechanism is used to provide rough alignment of the mirrors to the FPA entrance apertures.



Figure 6‑2:  Face-on view of mirror actuator assembly showing the three actuators and composite structure.


6.1.2   Optical Bench Structure


The instrument optical bench was constructed of hollow rectangular tubes and sheets composed of a graphite/cyanate ester composite material, designed to have a high strength-to-weight ratio, low CTE, and to insure dimensional stability over long integrations (200 kilo-second).


Individual pieces of the optical bench structure were bonded to form subassemblies that were joined by titanium fittings.  Optical components were mounted to inserts built into the structure.  Other components, such as the mounting plates for the electronics boxes, the baffles, contamination cover, etc., were mounted to the structure by flexures to avoid introducing any loads that might deform the structure and disturb the alignment of the optics. Exceptions to this practice, such as the lower baffle extensions, involved flexible materials that were not expected to apply significant loads. During thermal vacuum testing it was realized that mounting the contamination covers to the spectrographs resulted in undesirable motions of the grating bench. That mounting was modified and the new mounting validated in a second thermal vacuum test prior to launch. The decoupling of secondary hardware from the optical bench was not entirely satisfactory, however. The primary mirrors and gratings were found to move on orbital timescales once on-orbit. These motions, described in Section 5.2, were a significant complication to on-orbit operations.


Mechanical G-release, thermal expansion, and moisture desorption were expected to change the structure's dimensions significantly upon orbital insertion and slowly over the life of the satellite. Any errors in the placement of the mirror assemblies on the instrument optical bench would further misalign the mirrors with respect to the spectrograph (Conard et. al, 1999). Therefore to maintain alignment post-launch and throughout the mission, each mirror was equipped with precision actuators that permitted on-orbit mirror alignment.


6.1.3   Thermal Control


Thermal control is maintained by heaters attached to the intermediate plates, which radiatively couple to the rear of the mirror substrates.


The temperatures of the optical elements were regulated by heaters mounted on their enclosures (pie pans), and on the intermediate plate in the case of the mirrors. The instrument thermal control system was always configured to maintain the temperatures of the optics at least two degrees C warmer than their surroundings to minimize the likelihood of contamination.


6.1.4   Pre-launch Performance Specification & Evaluation


Optical testing and modeling of the FUSE primary mirrors were conducted toward a prediction of the on-orbit mirror point spread function (PSF) and its impact on spectrograph slit transmission. The image test was not meant to fully characterize the performance of the telescope mirrors in the FUSE bandpass. Rather, it was designed to insure that there were no severe problems with the flight mirrors and the implications of surface metrology data were understood. The test produced a data set which we used to validate our modeling and extrapolate a prediction into the FUSE bandpass with confidence.


Pre-launch measurements verified that the reflectivities of the optics exceeded their requirements of 32% for the SiC optics and 60% (λ > 1050) for the Al/LiF optics. Degradation of total system throughput of up to 20% per year was anticipated, but on-orbit performance was much better. The effective area of FUSE as a function of wavelength and time is presented in Figure 4‑2. Total throughput for the Al/LiF channels declined by less than 15% in the mid-wavelength band, and by ~25-30% in the long-wavelength band over the first three years, and was roughly stable thereafter. Performance degradation of the SiC optics differed: throughput declined at a roughly constant rate, falling by ~50% over 8 years. It is not possible to separate degradation of the primary mirrors from that of the gratings.


In addition to meeting reflectivity specifications (Oliveira et al. 1999), the fully assembled primary mirrors had an imaging requirement to contain  90% encircled energy (EE) at 100.0 nm in a diameter of 1.5 arcsec. This corresponded to a 16 micron diameter at the focal plane. To meet this requirement, FUSE established fabrication tolerances based on SOHO SUMER mirror heritage (Saha et al. 1996) and a modulation transfer function (MTF) analysis carried out at The Johns Hopkins University (JHU). The surface fabrication specifications were: figure error better than  λ/40 RMS and  λ /10peak-to-valley (P-V) at λ = 632.8 nm; midfrequency error less than 20 RMS over 10.0-0.1 mm spatial scales; and microroughness less than 10 RMS over 100-1 microns. These performance specifications were to ensure adequate transmission through the 1.25 × 20 arcsec, high-resolution spectrograph slit and instrument spectral resolution when using wider slits.


The in-flight mirror PSF performance was consistent with pre-flight measurements and model predictions (Ohl et al. 2000a, Ohl et al. 2000b). The mirror assemblies met the encircled energy requirement of 95% transmission through the 4 arcsecond MDRS aperture, and provided 90% transmission through the 1.25 arcsecond HIRS aperture, far-exceeding the 50% requirement.


In summary, the primary mirror assemblies were lightweight and adjustable in three degrees of freedom, maximized instrument reflective area in the bandpass, and met a stringent imaging requirement.

6.1.5   Mirror Positioning System Design

The position of each telescope mirror was adjustable through a Mirror Positioning Assembly (MPA).  Each MPA (Figure 6‑3) consisted of three actuators controlled by one of two Mirror Assembly Electronics (MAE) packages, two thermistors, and one strip heater, with redundant heater elements on the side of the intermediate plate closest to the mirror. The heaters held the assembly to 22 degrees +/- 0.5 degrees C.  Each MAE controlled a pair of mirror assemblies with one LiF coated mirror and one SiC coated mirror.

Figure 6‑3 Left:  Dummy aluminum mirror with actuator assembly (ref Ohl).
Right: Full flight mirror assembly, including pie-pan thermal enclosure and aperture stop.

Three actuators provided focusing and optical alignment for each of the four mirrors ( Figure 6‑4), enabling a focus resolution of 0.310 microns over a +/- 2.1 mm range and a tip/tilt resolution of 0.160 arcsec over a +/- 32.34 arcmin range. The A actuator controlled the rotation of the telescope mirror about the IPCS y-axis and consequently moved the target in the X-direction on the FPA (and the dispersion direction on the detector). The B and C actuators controlled the rotation of the telescope mirror about the IPCS X-axis.  The combined motion of the B and C actuators adjusted the position of the target in the y-direction on the FPA.  Equal motion of all three actuators moved the mirrors in the z-direction, or focus. Every movement of each actuator had to be tracked on the ground during the mission. Their range of motion is discussed in Section


Figure 6‑4: The actuator locations relative to the mirrors and the IPCS coordinate frame.


The mirror assembly positioning actuators, each contained a stepper motor and an optical soft stop sensor that electrically interfaced with the appropriate MAE.  The mirrors were positioned by an open-loop system. The initial on-orbit focus and alignment was determined from ground test. The MAEs had no memory, nor decision-making capability, nor were the absolute positions of the actuators telemetered. The specified stepper motor was commanded to move a certain number of steps in a specified direction, and the MAE attempts to comply with the command.  Neither the MAE or the IDS was informed if a hard stop was encountered. The hard stop was also not telemetered. To facilitate alignment in the event knowledge of the mirror position was corrupted, each mechanism had a "reference position".  This reference was a known position and, similar to the soft stops, was sensed using an optical disk encoder and therefore independent of commanding a known number of steps enabled.

In general, except for the "soft stop" function, the MAE did not know the current mechanism position. As a consequence, the ground software needed to maintain a record of the current expected actuator position and drive state of each of the twelve mechanisms. This record was updated each time one of the mechanisms was commanded.

To achieve initial alignment and focus, heavy usage of the MPAs was expected at the beginning of the mission, with only minor mirror adjustments expected later to compensate for structural changes from outgassing. Given this expectation for infrequent adjustments of the mirror positions, no position encoders were placed on the actuators and an on-board mirror motion manager, which could track mirror position through motion commands,  was not developed.  Post-launch, the four mirror channels (SiC1, SiC2, LiF1, and LiF2) were found to wander with respect to one another. The design savings and the lack of direct mirror tracking proved to be very costly in the operations phase where the frequent need for mirror adjustments required a significant modeling, planning, tracking, and analysis effort.

6.2           FPA Motion Control

The FPA electronics were functionally separated into two sides, one for each Rowland circle spectrograph.  Each side contained one LiF and one SiC FPA mechanism.  The complete subsystem consisted of four mechanisms, one electronics box, and the connecting harnesses. The FPA actuators provide motion of the slit plate along two axes: tangential to the Rowland circle (X) and along the radius of the Rowland circle (Z). These FPA axes are rotated by 30.5 degrees with respect to the IPCS axes, such that the FPA X axis is rotated towards the IPCS Z axis.

6.2.1    FPA X Axis

A High-Output Paraffin (HOP) actuator controlled via a closed loop system drove the X-axis.  Operationally, all motion commands to the X-axis mechanisms were issued to the IDS FPA driver task.  This task verified that the commanded position was within specified software limits, and then generated the appropriate FPA hardware command for the requested motion.  The data number issued by the driver was stored in a command register that was used by a digital to analog converter to provide a reference signal for the control loop.  The controller managed the temperature of the paraffin to maintain the FPA at the commanded position.  The X ‑ axis actuators were designed to allow a minimum of 10,000 individual motions and/or 1000 full displacements.  The commandable range of motion was 400 microns. The FPA X axis actuators were adjusted routinely for purposes of channel alignment and for FP-split procedures.

6.2.2   FPA Z Axis

The Z-axis was driven by a stepper motor and lead-screw, commandable in increments of 38 motor steps.  Each of these units resulted in a motion slightly less than 10m.  Limit sensors were used to prevent motion beyond the design limit of 280 m.  The practical operational limit, however, was 250 m.  The design requirement for each stepper motor was 80 in-flight motions.  The history of in-flight FPA motions is given in Table 6.2‑1. The first focus run was only partially-executed, but the second spanned the full range of -240μm to +240μm in steps of 40μm each. The final FPA Z positions were set on 24 March 2000, except that the LiF2 Z position was adjusted on 29 June 2005 to improve the focus of FES-B, and then back again at the end of the mission for the final series of airglow observations.


Launch 1999-06-24   60 20 0 0
I8150109005 1999-11-23 17:55:05 Focus run
P1020504001 1999-11-24 18:08:09 60 65 -100 -75
I8160101 1999-11-26 19:11:34 Focus run
I2070801001 1999-11-27 06:49:49 64 66 -100 -77
P1010502012 1999-12-12 14:09:50 64 210 -238 -225
S5130301001 2000-03-24 13:53:37 -35 145 -100 60
M1051402002 2005-06-29 11:40:36 -35 -250 -100 60
S1001701001 2007-08-15 14:26:06 -35 145 -100 60

Table 6.2‑1: History of FPA Z-axis motions. The first three columns give the observation ID, date, and start time for the observation immediately following a change in FPA Z position.


6.3           Detector Design and Pre-launch Characterization

6.3.1    Hardware and Software Description

The FUSE instrument included two Double Delay Line (DDL) detectors which collected incoming photons and measured their positions. Three FUV detectors were built at the Space Sciences Laboratory at the University of California, Berkeley. Unit FL01 remained on the ground as a spare and was used for ground testing, while units FL02 and FL03 were used in the instrument as Detector 1 (side 1 of the instrument, collecting photon events as part of the SiC1 and LiF1 channels) and Detector 2 (SiC2 and LiF2), respectively. The three detectors had identical physical characteristics, and were designed to be as similar as possible. Differences between them are due primarily to differences in the microchannel plates and the adjustments of the electronics to account for those variations. The differences in wavelength coverage between side 1 and side 2 of the instrument were determined by the mounting locations on the Rowland circle.


Each detector consisted of two segments. Mechanically, each detector was a single unit; electrically each segment was unique with most of its own electronics. Keeping the segments separated permitted each to be individually optimized. In addition, it ensured that a problem with one segment did not prevent its companion from being operated normally. Because of this design, one detector segment could be operated normally while the high voltage on the adjacent segment was turned off.


Light coming from one of the FUSE gratings to the detector first passed through a 95% transmission, +15 volt ion repeller grid, then through a 95% transmission QE Grid designed to improve the quantum efficiency of the system, before reaching the KBr-coated microchannel plates (MCPs). The photons striking the photocathode were converted to photoelectrons via the photoelectric effect, multiplied as they passed through the stack of three (Z-stack) MCPs, then proximity-focused onto the DDL anode. The DDL electronics determined the location of each charge cloud by measuring the time it took for the charge to propagate along the anode (for the X, or dispersion direction) or by charge division (for Y, or cross-dispersion). The top-level detector specifications are summarized in Table 6.3‑1.


Specification Description
MCP pore size
(pore diameter / center-to-center spacing)
10 m / 12.5 m (top & bottom plates)
12.5 m / 15 m (middle plates)
MCP pore bias angle 13
MCP Configuration Z-stack
MCP size 95 mm x 20 mm, 80:1 L/D
MCP resistance < 30 MΩ
Anode Type Double Delay Line
Photocathode KBr
Ion Repeller Grid 95% transmission, 1247 x 1247 m spacing, flat
QE enhancement Grid 95% transmission, 1042 x 1009 m spacing, curved to match MCPs
QE in FUSE bandpass 14 30%
Active Area 85 x 10 mm x 2 segments
Curvature of front surface of MCPs 826 mm radius
Number of Pixels 16,384 x 1024 per segment
Pixel size 6 m x 1017 m, depending on segment
Detector resolution ~20 m x ~80 m
Lifetime Specification > 107 events per 103 m2

Table 6.3‑1: Detector Specifications


Each detector subsystem was composed of three interconnected, modular assemblies. These were the Vacuum Assembly, Electronics Assembly, and Stim Lamp Assembly. The Vacuum Assembly was mounted in the spectrograph cavity and contained the detector imaging elements (grids, MCPs, anode, etc.) in a stainless steel vacuum housing, along with a high voltage filter module, charge amplifiers, timing amplifiers, a motorized door and mechanism, and ion pumps to maintain a high vacuum inside the vacuum box before launch. The Electronics Assembly, mounted to the instrument electronics baseplate in the electronics cavity, included the low- and high-voltage power supplies, Time-to-Digital-Converters (TDCs), Charge-to-Digital-Converter (CDCs), and a Data Processing Unit (DPU), along with an interface to the instrument computer the Instrument Data System (IDS). The Stim Lamp Assembly consisted of a mercury vapor lamp which was mounted to the spectrograph structure inside the spectrograph cavity and was powered and controlled via the detector electronics. Details on each of these assemblies are given in the sections below.


Each detector subsystem included thirteen thermistors (Table 6.3‑2) to monitor the temperature of the detector hardware. Although the temperatures of the anode and some parts of the electronics were known to affect the data, the thermistors information was only used as a general diagnostic, and detector stim pulses, described below, were used to account for temperature effects in the data.


Telemetry Mnemonic Location
I_DET1AMPATEMP Segment A Charge Amplifier
I_DET1AMPBTEMP Segment B Charge Amplifier
I_DET1DOORTEMP Detector Door
I_DET1DPUTEMP Data Processing Unit
I_DET1ELTEMP Detector Electronics Assembly
I_DET1HVFLTRTEMP High Voltage Filter
I_DET1HVMODLTEMP High Voltage Module
I_DET1PCTEMP Power Convertor
I_DET1PLTEMP Detector Backplate

Table 6.3‑2: Detector 1 Thermistors

More details on the design of the FUSE detectors can be found in Siegmund et al. 1997 and Sahnow et al. 2000.  Vacuum Assembly


Each Vacuum Assembly consisted of a stainless steel vessel which contained the grids, MCPs, and anodes used for collecting the photons, along with two 4 liter per second ion pumps; also included as part of this section and mounted outside of the vacuum box was much of the electronics used for calculating the locations of the photons. Since the MCPs and the photocathode had to be operated and stored in a vacuum environment, a recloseable door was included in the design. During ground testing in vacuum the doors were opened to allow FUV light to reach the MCPs. When ground testing was complete, the doors were closed and the vacuum maintained with the ion pumps. A pumping port allowed the vacuum box to be evacuated with a ground vacuum pump, and then an attached valve could be closed once the pressure was low enough for the ion pump to maintain the pressure. The doors were opened in orbit on July 16 and 17, 1999, after the spectrograph cavity had partially outgassed, and they were never closed again. Two sapphire windows (one for each segment) on each door allowed visible light to reach a portion of the MCPs when the doors were closed. This permitted some ground testing of the detector while it was at a safe pressure (< 10-6 Torr) despite the fact that the instrument was at normal atmospheric pressure.


Inside each vacuum box were two electrically independent segments mounted as a single mechanical and optical unit. Since the MCPs were part of the optical system, the MCP stackups were mounted to and precision aligned to the detector backplate. The backplate had an optical cube which was used to align the detector during spectrograph integration and test. The preflight specification was that the front surface of the MCPs be held to within 25 m of the 826 mm Rowland circle in order to ensure the highest resolution. During detector installation, shims were inserted to align the MCPs on each detector to the Rowland circle by using an optical cube on each backplate. The MCP positions had been previously determined with respect to the cubes to within ~37 m. During the alignment of Detector 1, a 6.3 arcsecond tilt was inadvertently introduced, but the instrument resolution was not limited by this tilt.


Each stackup contained a Z-stack of three 95 mm × 20 mm MCPs which acted as the photon-sensitive surfaces, although a front aperture mask limited the active area of each segment to 85 × 10 mm, with a ~7 mm gap between them. The top and bottom MCP in each stack had 10 m pores on 12.5 m centers, while the middle plates had 12.5 m pores on 15 m centers; these were different in order to help minimize the moir pattern which often occurs when identical MCPs are stacked (Tremsin et al., 1999).


The top plates were coated with an opaque KBr photocathode in order to improve the sensitivity to FUV radiation. Photons impinging on the front surface of one of the MCPs, which were held at a high potential, created photoelectrons which were accelerated down the pores due to a high potential maintained across the plates. Each time an electron struck the walls of one of these pores, multiple secondary electrons were generated, so that a single photon incident on the top plate resulted in a cloud of ~107 electrons which was several millimeters in diameter exiting the back side of the third plate. These electrons were accelerated across a 7 mm gap to a helical double delay line anode with a period of 600 m and an active area of 94 x 20 mm, where they were collected. The anodes were constructed on a flexible RT/Duroid substrate, and were curved to match the MCPs. The anode-MCP separation was 7 mm, and voltage difference was 550 volts.


Figure 6‑5: Expanded view of the detector stack mounting in the FUSE detector. The QE grid is held by the frame at the top, and the curved MCPs are mounted to a cylindrical surface to match the Rowland circle.

In front of the MCPs were two 95% transmissive electroformed nickel mesh grids to improve the performance of the detector. The front, flat grid, which was mounted on the vacuum box aperture, sat approximately 35 mm in front of the MCPs and was held at +15 volts; it was designed to exclude ions from the vacuum cavity. This ion (or plasma) grid was powered whenever the detector electronics were powered; it was not controllable in flight, but housekeeping telemetry was available to show the status of the grid voltage. In addition, a Quantum Efficiency (QE) grid, which was curved to match the curvature of the MCPs, was placed 6 mm above the front surface of the MCPs in order to improve the QE of the system by forcing photoelectrons generated by the MCP web back down the pores. The QE grid was held at 1200 volts above the front MCP. It was the source of the optical anomaly known as the worm (see Section 4.3.4).


The detector body assembly was bolted to a stainless steel backplate which provided the mechanical and thermal interface to the rest of the instrument. A mu-metal magnetic shield, which surrounded the body assembly and shielded it from external magnetic fields, was also mounted to the backplate. This entire assembly was enclosed inside the vacuum box.


Figure 6‑6:   A FUSE detector Vacuum Assembly mounted to the detector mounting bracket in the spectrograph cavity. The door assembly is at the top, with the light baffle protruding. Two ion pumps are visible at the front right, and the high voltage filter module, charge amplifiers, and timing amplifiers are visible behind them. The ladder-like structures at the top on either side of the detector are spectrograph baffles.


Also mounted to the backplate, but outside of the vacuum box, were two high voltage filter modules which provided the MCP, QE grid, and rear-field voltages. The backplate also supported the amplifiers, which amplified the charge collected by the anode before passing them on to the Detector Electronics Assembly. The Vacuum Assembly was thermally connected to the electronics baseplate via copper straps, which also provided chassis grounding. Figure 6‑6 shows a photograph of a Vacuum Assembly mounted in the FUSE spectrograph.  Electronics Assembly


There was one Electronics Assembly for each detector; each consisted of nine interconnected electronics boxes (Figure 6‑7). The major functional components of each were:

  1. The power supplies, which were responsible for providing power (high voltage to the high voltage filter modules and ion pumps and low voltage power for other components) to the detectors.
  2. A Delay Spool, used to delay the signal from one side of the anode in order to ensure that the START pulse always preceded the STOP pulse.
  3. Two 14 bit Time-to-Digital Converters (TDC) modules, which measured and digitized the signals from the vacuum assembly electronics into digital signals in order to measure the X (dispersion direction) coordinates of an incident photon.
  4. Two Charge-to-Digital Converters (CDC) modules to digitize the charge from the upper and lower delay line for use in calculating the Y (cross-dispersion) position of an incident photon.
  5. A Data Processing Unit (DPU), which processed the digital signals from the TDCs and CDCs, provided command and housekeeping functions, and provided an interface to the IDS computer and other parts of the instrument.


Functional block diagrams of the system are shown in Figure 6‑8 and  Figure 6‑9. The following sections describe some of these components in more detail.

Figure 6‑7: Electronics Assembly and Stim Lamp Assembly of the spare detector.


Figure 6‑8: Block diagram of the encoding electronics for the FUSE detectors.

Figure 6‑9: Functional Block Diagram of the Detector Electronics Assembly (hardware and software).   Digitizers


The timing digitizers determined the X (dispersion direction) position of the charge cloud to ~20 m by measuring the difference in time it took the charge pulse to propagate to the two ends of the anode, and then digitized the results to 14 bits (16384 pixels). The scale was adjusted so that the pixel size was very close to 6.0 m, although detector distortions meant that it varied with position. Deviations were generally only a few percent, except at the ends of each segment, where they could be much larger.


In the Y (cross-dispersion) direction, separate Y1 and Y2 signals from the upper and lower half of each DDL anode were digitized to 12 bits by the charge digitizers. The digitized X, Y1, and Y2 values were then passed to the detector DPU for further processing.     Stim Pulses and Thermal Stability


Since both anode and electronics temperatures were known to affect the measured location of photon events in the dispersion direction, it was necessary to provide a thermal correction to the data in the CalFUSE pipeline. Rather than relying on the measured temperatures, however, the positions of two stimulation (stim) pulses on each segment were used for this correction. These pulses were generated by injecting charge into the amplifiers so that pseudo-events are created near the left and right ends of each segment, beyond the active area of the MCPs. Since the stim pulses passed through the same electronics as incoming photons, this method provided a more accurate method of measuring the thermal effects on the data than relying on the temperatures. The stim pulses were normally commanded on for 59 seconds at the beginning and end of every exposure, although for snapshot exposures they were left on throughout the exposure. Stim rates varied slightly from segment to segment, but were approximately 45 counts per second per stim, giving a total rate of about 90 per second per segment.


Prelaunch ground testing using the spare detector showed that the stim pulses moved as a function of temperature. In particular, the temperatures of the anode (best tracked via the backplate thermistor) and the electronics (tracked with the TDC temperature) were both shown to affect the stretch (the separation between the two stims) and shift (position of the mean stim position) of the detector format, with the anode temperature dominating the former, and the electronics temperature dominating the latter.  On orbit stim pulse data show strong correlations between backplate temperature and stretch, and between the stretch of two segments on the same detector, which is consistent with the ground measurements. The inflight shift measurements, however, showed very different behavior between the segments (Figure 6‑10).


Figure 6‑10: Detector X shift, as measured by the change in position of the stim pulses, for all four segments during the mission. Long term trends appear to dominate short term temperature effects, particularly on segments 1B and 2B.


The positions of the stims in an exposure are measured by CalFUSE, and a linear shift and stretch is applied to the raw data in the dispersion direction in order to move the stims to their nominal position. In the cross-dispersion direction only a shift is applied, since their separation is too small to allow reliable stretch information.


Because the stim pulses are artificially generated, their shape does not exactly match the properties of events from the MCPs. In fact, the shapes of the stim pulses change as a function of count rate, with a sharp, single-peaked image at low count rates turning into a double-peaked distribution at high count rates. For normal FUSE count rates, the second peak was small and thus did not cause CalFUSE difficulty in measuring the position.


The correction for thermal effects based on stim pulse positions assumes that the corrections are linear over the detector, and that the thermal environment is changing slowly. Large temperature changes, such as those due to bringing the detector high voltage from SAA level to full, could result in rapid changes of the scale during an exposure, and those effects are much harder to correct for. In addition, this method assumes that the stims accurately track the overall detector format. While this is believed to be true in general, numerous unexplained changes in stim position that could not be correlated with any known external effects were seen (Figure 6‑10). These may result in uncorrected errors in the calculated wavelength scale. Some exposures are missing one or both stim pulses. In those cases, CalFUSE must assume positions based on a record of typical positions from other exposures.   Data Processing Unit     Overview


The detector Data Processing Unit (DPU) was responsible for the electronic processing of the collected photon events output by the digitizer. DPU memory contained two 16 KB regions from which code could be executed; these were known as the upper core (UC) and lower core (LC) memory regions. At boot up, a control program was loaded into memory from an onboard EPROM. This was a basic version of the DPU code (version 16200) which had been programmed into the PROMs before spectrograph Integration and Test in 1997. Other versions of DPU code (Section could then be loaded (from the ground or IDS memory) into UC memory, and execution switched there.


The main functions of the detector Data Processing Unit (DPU) were:

  1. Accepting commands from the IDS on the MIL-STD-1553 Instrument Data Bus and passing them on to the detectors.
  2. Collecting the photon event data (X, Y1, and Y2) from the digitizers and calculating the Y coordinate (Section
  3. Applying masks to the data (Section and counting events.
  4. Packaging the science data and sending it on to the IDS via the RS-422 Science Data Bus.
  5. Protecting the detector in case of count rates that were too high (Section
  6. Accumulating 7 bit pulse height histograms (Section
  7. Monitoring current draw from the high voltage and auxiliary power supplies, and reporting diagnostics if necessary (version 16500 of the code and above - Section
  8. Collecting housekeeping telemetry from the detector.
  9. Monitoring the detector behavior, and issuing a diagnostic code if an unexpected event occurred (Section
  10. Formatting housekeeping status, counters, and pulse height information and passing it to the IDS via the Instrument Data Bus.
  11. Continuously calculating checksums of specified regions of detector memory, and issuing diagnostic codes when the checksum changed.      Calculation of Y Position

After the output from the digitizers was passed to the DPU, the Y coordinate was calculated by taking the ratio of charge between the two halves of the DDL, as Y = Y1/(Y1+Y2), where Y1 and Y2 are the charge on the upper and lower half of the delay line. This result was digitized to 10 bits (1024 pixels), although the accuracy was only ~80 m. Y pixel sizes varied from ~ 9 to ~17 m in the raw data, depending on detector and segment. The pulse height, or total charge collected, was also digitized at this point as part of the y calculation. The result was digitized to 7 bits and used to construct a pulse height histogram (Section     Masks and counters


For each segment there were four programmable masks (shown in the block diagram in Figure 6‑9) which were used to monitor the count rate in a specific region of the detector. Each mask had an associated counter which counted the number of events in each of the regions defined by the masks. The masks and counters were the Active Image Mask and Active Image Counter (AIC), the SiC and LiF masks and counters, and the Autonomous Shutdown Mask and Autonomous Shutdown Counter (ASC). The latter was also known as the SAA counter, since it was designed to measure background counts in the South Atlantic Anomaly. Each of the counters recorded the number of events that fell in a particular mask region on each segment. DPU commands were used to enable or disable regions of the segment for each counter and mask in 16 x 16 pixel regions. A complete set of aperture-specific masks for all four segments was loaded near the beginning of each observation script via the set_det_mask_f flight script. This script cleared all previous masks, then loaded an Active Image Mask that covered the entire segment, an Autonomous Shutdown Mask that fell on an unilluminated portion of the segment, and SiC and LiF masks that were placed to collect essentially all the events from the aperture specified for that observation.


The Active Image mask was used to define the region of a segment to pass to the IDS, thus acting as a spatial filter. It could have been used to exclude a region around a hot spot, for instance, so that the Science Data Bus would not be overwhelmed by these events. However, these masks were left enabled everywhere on all four segments throughout the mission. Thus, the Active Image Counter recorded the total number of events passed from the detector to the IDS on the Science Data Bus. Each event which passed through all the thresholding was output as a 32 bit word containing the detector ID (1 bit - MSB), segment ID (1), x (14), y (10), pulse height (5), and a format bit (1) to identify this as a photon event. These words were passed on to the IDS for further processing via the Science Data Bus.


The SiC and LiF masks were used to define the regions of the detector to use for target peak up, so they were different for each aperture and each segment. The objective was to determine the count rate from the object in the slit, while minimizing the count rate from the other apertures and the detector background. These masks included the entire spectral range covered on each segment, so they included any airglow in the aperture.


The Autonomous Shutdown Mask was used to monitor a region on the detector which was not illuminated by any of the apertures, so it provided a way to monitor the detector background. The DPU code used the count rate for bright object protection; if the count rate over a specified time for a particular segment exceeded a predefined threshold, the high voltage on that detector segment was lowered to SAA level. This protection was designed to minimize the chance of scrubbing the detector due to an extremely high count rate at nominal gain. As originally envisioned, this was designed in to the system primarily to protect against passing through the SAA without lowering the high voltage (as a result this counter was also known as the SAA counter). In practice, most of the shutdowns were due to event bursts, which complicated the way thresholds had to be set (Section


The SiC, LiF, and Autonomous Shutdown masks were changed six times during the first two years of the mission (>Table 6.3‑3) in order to better align them with the position of the spectra and make them match the SIA tables (Section 6.3.3).

At Launch Prelaunch values
8/20/1999 Adjusted SiC1 and LiF1 based on first view of side 1 spectra
11/5/1999 Adjusted SiC2, LiF2 and ASC2 based on first view of side 2 spectra
4/1/2000 Clear ASC mask before setting to avoid adding default mask to new ASC
4/19/2000 Change ASC mask on segment 2A
11/18/2000 Reinitialize ASC mask to avoid SEUs
10/17/2001 Change size of ASC2 mask to avoid contamination from MDRS
11/19/2001 Match masks to SIA tables
4/18/2003 Special AICs for bright objects

Table 6.3‑3: Summary of detector mask changes during the mission


An additional counter, the Fast Event Counter (FEC also known as the Front End Counter) measured the total number of events over the entire segment at the output of the fast amplifiers. The number of counts lost before the FEC, e.g. due to the MCPs themselves, is very small at typical FUSE count rates, and was therefore ignored. The IDS code monitored the FEC rate for each segment and shut off the high voltage for that segment whenever the rate exceeded 45,000 counts per second for three seconds. These values were hard coded, but IDS scripts were written to poke the relevant memory locations in order to change them if desired. They were changed temporarily several times during the mission when bright objects were observed.


The Digitized Events Counter, or DEC, counted the number of events that were processed by the digitizer and reached the DPU FIFO (again, over the entire segment). This was typically less than the number of counts in the FEC, since there were electronics dead time effects which caused events to be discarded in the digitizer. The DEC and FEC were used in the calculation of the dead time of the instrument (Section 6.9).


A summary of the masks and counters, along with their functions is given in Table 6.3‑4.

Mask / Counter Name Mnemonic Operates On Bit in Mask Word Description
Fast Event Counter or Front End Counter FEC Full segment N/A Total counts reaching the detector
Digitized Event Counter DEC Full segment N/A Events above threshold
Preamp Reset Counter PRC Full segment N/A  
Active Image Counter AIC Modifiable 1 (LSB) Counts sent to the IDS
Silicon Carbide SiC Modifiable 2 Peakup in SiC channel
Lithium Fluoride LIF Modifiable 3 Peakup in LiF channel
Autonomous Shutdown Counter ASC or SAA Modifiable 4 (MSB) Detector background

Table 6.3‑4: Detector masks and counters


All counters were stored as 24-bit values (0 16,777,215) and were updated once per second. They counted continuously from the time the detector was turned on, so they would regularly roll over. They were typically sent to the ground via housekeeping telemetry once every sixteen seconds, although that rate changed in certain telemetry modes.


The FEC, AIC, SiC, LiF, and ASC counter values are saved for each exposure in the housekeeping ( files, and the first four of these are used to populate arrays in the Intermediate Data File created by CalFUSE. The FEC, DEC, SiC, LiF, and ASC counter values at the beginning and end of each exposure are also recorded in the engineering snapshot and then copied into the header of the science data files. CalFUSE uses the counter information for calculating dead time (Section 6.9) and Y blooming (Section effects. It should be noted that because the SiC and LiF counters change with aperture, the count rates for observations taken in different apertures are not directly comparable.     Pulse Height Histograms


As described previously, a 5-bit pulse height for each photon event is transmitted to the IDS along with the position information. For TTAG exposures, all of this information was sent to the ground, but for HIST exposures this pulse height information is discarded along with the timing information. However, for both TTAGs and HISTs data, the detector DPU also accumulated a 7-bit pulse height histogram over each detector segment and sent it (in pieces) to the IDS. The IDS accumulated these pulse height distributions for all four segments and downlinked a complete set of PHDs every 128 seconds. Although these full-segment PHDs could not be used to determine gain as a function of position, they were useful diagnostics which could be used to monitor the gain of the detector during high count rate observations, such as the stim lamp exposures or other HIST observations. The higher resolution was also potentially useful for setting the onboard charge thresholds, but they were never adjusted (except for testing) in flight.     Current Protection


An important function of the DPU code was to monitor the current draw of the microchannel plates (HVIA and HVIB for segments A and B), and the auxiliary current power supply (AUXI). The AUX power supply was used to power the ion pumps, the detector door, and the stim lamps.


More details are given in the discussion of High Voltage Transients in Section 6.3.5.     Detector Diagnostic Codes


An important job of the DPU code was to continually monitor the detector hardware and software to look for out-of-limit conditions. When a problem was identified, a diagnostic code was issued and other protective action was taken if warranted (Table 6.3‑5).  Since the two detectors were independent, they each issued their own diagnostic codes.


Diagnostic (hex) Description Effects on operations
16 AUXI threshold exceeded* (mini-crackle) None
18 CRC changed in lower RAM Associated CRC change triggers IDS process to reload DPU code
19 CRC changed in upper RAM Associated CRC change triggers IDS process to reload DPU code
1A CRC changed in ROM None
1B Watchdog reset as part of power up None
20 CRC changed in segment A mask None
21 CRC changed in segment B mask None
22 Counts in ASC counter exceed threshold in one of the segments (SAA shutdown) Drop HV on affected segment to SAA level.
27 HVIA threshold exceeded (mini-crackle) None
28 HVIB threshold exceeded (mini-crackle) None
2A HVIA threshold exceeded for longer than persistence time* (crackle) HV shut down on both segments.
2B HVIB threshold exceeded for longer than persistence time* (crackle) HV shut down on both segments.
2E AUXI threshold exceeded for longer than persistence time* (crackle) HV shut down on both segments.

Table 6.3‑5: Selected DPU Diagnostic Codes
* Not included in DPU code version 16200


The diagnostic codes in the table above can be divided into three types:

  1. Power on (1B) When a detector is powered on, a diagnostic 1B was issued.
  2. CRC changes (18, 19, 1A, 20, 21) A background process in the DPU code continually checked the detector code (stored in both RAM and ROM) and masks for changes. Although some changes were expected (when uploading new masks at the beginning of an observation, for example), many unexpected changes due to bit flips were seen during the mission (see Section 6.3.4).
  3. High count rates (22) When the number of counts in the ASC for a segment exceeded a defined threshold, the detector HV on that segment was dropped to SAA level. This behavior was designed primarily to protect the detector from passing through the SAA at full voltage levels, and it did serve that purpose several times during the mission (Section 4.4.8).
  4. High voltage problems (16, 27, 28, 2A, 2B, 2E) See Section 6.3.5.     DPU Code Versions


Six different versions of DPU code were used during the mission. The properties of each are summarized in Table 6.3‑6. Version 16200 was resident in the detector DPU PROM, and was loaded into lower core memory when a detector was turned on. It was used during detector ground characterization, spectrograph integration and test, and part of instrument and satellite integration and test. This version had no protection for high voltage transients (crackles).


Version 16500 was provided by UC Berkeley in early 1999 to add protection and diagnostics after the first detector crackle was seen during instrument testing at Goddard Space Flight Center. This version also decreased the range of upper core code memory checked by the CRC checking task. This decreased the susceptibility to upper core SEUs by about 25%, since only 10,544 bytes (instead of 14,336) were being checked.


UC Berkeley provided Version 16600 for improved crackle protection after initial on-orbit experience. It added downlinking of the UC CRC (in addition to the LC CRC which had been in previous versions) so that automated code loading could be implemented. It also extended the CRC change made in the previous version to the lower core code region.


After an increase in crackles in 2001, Berkeley provided version16603, which was functionally identical to 16600 except for an increase in the persistence from 20 msec to 60 msec.


Late in the mission, several additional versions were built at JHU with increased persistence values. Only two of those modified versions, 16710 and 16720, were ever used during normal operations.

Start Date Code Version Threshold[1]
(digital units)

Persistence (msec)[2] Notes
1999:181 [3] 16200 128 128 N/A N/A N/A Version in DPU ROM
1999:191 [4] 16500 128 128 20 20 255 Added crackle protection
1999:334 16600 128 128 20 20 255 Downloaded UC CRC
2000:323 16600 128/182 128 20 20 255 Change to flight script
2001:326 16603 128/182 128 60 60 255 Increased persistence
2006:290 16603 182 128 60 60 255 Change to ramp up script
2006:333 16710 182 128 128 60 255 Used for detector 2 only
2007:025 16720 182 128 128 128 255 Increased persistence

Table 6.3‑6: DPU code versions used during normal operations. Different values were used during stim lamp exposures.

[1] The threshold was set via ground command, but the values listed were rarely changed. They were set to higher values during HV ramp ups and stim lamp operations.

[2] Persistence values are only approximately milliseconds.

[3] Date first used on orbit. Version 16200 was resident in detector ROM and was used for much of the pre-launch testing

[4] Date first used on orbit. Version 16500 was delivered during satellite integration and test and was used for ground testing starting in late 1998.  Stim Lamp Assembly


The detector Stim Lamp Assemblies (one for each detector) included the mercury vapor stimulation (or stim) lamps, a mounting bracket, and a pinhole aperture to coarsely control the amount and direction of light reaching the detectors. They were mounted to a structural bracket in the spectrograph cavity, approximately 1.25 meters from the detectors. Their location allowed direct, quasi-uniform illumination of each detector, with count rates of ~2,000 to ~12,000 counts per second, depending on segment. The stim lamps were not designed to provide a true flat field of the detectors, but rather were included to provide general diagnostics of detector health. Before launch, the lamps were used to provide detector aliveness tests while the instrument was at atmospheric pressure by illuminating the MCPs through the sapphire windows in the vacuum doors. On orbit, they were used regularly throughout the mission as a means of monitoring detector performance, especially gain sag (Section 4.4.2).


The stim lamps were powered through the detector auxiliary power supply, which also powered the ion pumps.


6.3.2   Normal Detector Operations  High Voltage Ramp-up

Ramping up the high voltage was a complex process with many built in safeguards to ensure that the detector remained safe. The ramp-up scripts:


Only after all of these steps were taken would the HV be ramped above its turn on (2500 volt) level. The ground scripts used for ramp-up referenced a file which contained the proper voltages. Separate scripts used for raising the high voltage for testing purposes did not permit increases of more than 10 digital units at a time to guard against accidentally increasing the high voltage to an unsafe level.


4.n.2.2  High Voltage Management: Occultation Manager


The detector occultation manager was implemented during the mission in order to minimize the amount of gain sag at the location of airglow lines. At the beginning of the mission, the high voltage was normally left at FULL except when passing through the SAA. This meant it was always ready for a peak-up or an observation, but it also meant that any photons falling on the detector removed charge from the MCPs, even though they weren't necessarily being collected by the IDS. Once the differential gain sag problem became obvious, development began on an occultation manager, which lowered the MCP high voltage to SAA levels whenever it wasn't needed. This decreased the amount of unnecessary charge depletion by removing several sources of non-productive exposure:

(1)   Airglow lines illuminated the detectors during occultations and other times between exposures, causing significant localized charge depletion.

(2)   It was not uncommon for a slew to a new target and the acquisition of that target to occur well before an exposure began. This could be due to an intervening SAA or occultation, for instance. The target could remain in the aperture at full high voltage for many minutes before the exposure began. Before the implementation of the occultation manager, this time contributed to the detector exposure and gain sag, even though no data was collected.

(3)   Similarly, after an exposure was completed, the target sometimes remained in the aperture until it was time to slew to the next target. Before the occultation manager, this would also add detector exposure without the benefit of collecting any additional data.


The occultation manager was used starting on 20 November 2001, and used for the rest of the mission. It was only turned off during times when the high voltage was being ramped up, and during other special tests.


6.3.3    SIA tables


As described previously, the FUSE detectors were photon-counting detectors, and the DPU produced a location and pulse height for each detected event. These positions were passed via the Science Data Bus to the IDS, which could either send that information on to the spacecraft (after adding timing information) for eventual downlink to the ground (TTAG mode), or add it to a two-dimensional image (HIST mode), discarding the pulse height information in the process. Due to memory limitations, an image of all four 16384 × 1024 segments couldnt be stored in IDS memory. Instead, only the regions around the primary spectra were saved in memory, and they were typically binned by 8 in the cross-dispersion direction. The primary exceptions to this procedure were the M999 stim lamp exposures, which were binned 2 × 2, and for which the entire detector area was saved.


For each HIST observation, a Spectral Image Allocation (SIA) table was used to define the detector regions which were saved to memory. The SIA tables varied with the aperture chosen, and they also changed during the mission as a better understanding of the location of the spectra was gained. SIA tables were actually part of the IDS, but because they define how the detector data is saved during histogram exposures, they are described here.


SIA tables consisted of an 8 × 64 array for each segment, with each array element referencing a 2048 × 16 pixel region on the detector. Each array element could be set to 1 to save the data from the associated detector region in IDS memory, or 0 to discard it. In most cases, the observation scripts for histogram exposures called the init_bulk_memory_f script, which loaded the standard set of SIA tables for all apertures. A special set of SIA tables for histogram stim lamp exposures were uploaded from the ground when needed. The standard SIA tables were changed several times during the first few years of the mission ( Table 6.3‑7 through Table 6.3‑12). These changes were made primarily to adjust the positions to ensure that all data from the aperture was being collected. Note that the masks (Table 6.3‑4) were eventually set to be the same as the SIA tables, but they did not match early in the mission.


The SIA tables were downloaded to the ground with the science data, and they are stored in the primary data unit in the raw FITS files. Although they are not used by CalFUSE, they can be examined to verify that the expected mask was used and was not corrupted. The typical SIA table consists of up to 16 rectangular regions turned on, which include, for each segment, the detector regions where the SiC and LiF spectra fall, along with the area around each stim pulse. In many cases, one or more of the stim pulse regions overlap with the spectral regions. The tables below describe the regions of each detector segment covered by the SIA tables during the mission.


It is important to note that since the SIA tables operate on the raw data coming from the detector, they are not in the geometrically-corrected FARF, but in a distorted frame which is not corrected for geometric distortions or thermal shifts and stretches. Thus, the SIA table regions were oversized to avoid losing data.


Start Date
X Range
Y Range
Y Range
Y Range
Y Range
11/20/1999 0 - 16383 160 - 319 128 - 287 288 - 399 400 - 495 siahirs1x8b
5/17/2000 0 - 16383 160 - 319 112 - 271 288 - 399 400 - 495 siahirs1x8c,

Table 6.3‑7 SiC HIRS SIA Tables used during the mission


Start Date
X Range
Y Range
Y Range
Y Range
Y Range
11/12/1999 0 - 16383 288 - 447 240 - 399 208 - 319 224 - 335 siamdrs1x8b
1/13/2000 0 - 16383 288 - 447 240 - 399 208 - 319 352 - 463 siamdrs1x8c

Table 6.3‑8: SiC MDRS SIA Tables used during the mission


Start Date
X Range
Y Range
Y Range
Y Range
Y Range
10/11/1999 0 - 16383 16 - 175 0 - 175 384 - 495 480 - 591 sialwrs1x8b
3/4/2000 0 - 16383 16 - 175 0 - 175 352 - 463 448 - 559 sialwrs1x8c,

Table 6.3‑9: SiC LWRS SIA Tables used during the mission


Start Date
X Range
Y Range
Y Range
Y Range
Y Range
11/20/1999 0 - 16383 624 - 783 608 - 767 544 - 655 592 - 703 siahirs1x8b
5/17/2000 0 - 16383 640 - 799 608 - 767 560 - 671 592 - 703 siahirs1x8c
3/27/2001 0 - 16383 640 - 799 592 - 767 560 - 671 592 - 703 siahirs1x8d

Table 6.3‑10: LiF HIRS SIA Tables used during the mission


Start Date
X Range
Y Range
Y Range
Y Range
Y Range
11/12/1999 0 - 16383 768 - 927 736 - 895 464 - 575 544 - 655 siamdrs1x8b
6/20/2001 0 - 16383 752 - 927 720 - 895 464 - 575 544 - 655 siamdrs1x8d

Table 6.3‑11: LiF MDRS SIA Tables used during the mission


Start Date
X Range
Y Range
Y Range
Y Range
Y Range
10/11/1999 0 - 16383 480 - 639 464 - 623 624 - 735 656 - 767 sialwrs1x8b
3/4/2000 0 - 16383 480 - 639 464 - 623 624 - 735 672 - 783 sialwrs1x8c
5/18/2000 0 - 16383 496 - 655 480 - 639 624 - 735 672 - 783 sialwrs1x8d
5/1/2001 0 - 16383 496 - 655 464 - 639 624 - 751 656 - 799 sialwrs1x8e

Table 6.3‑12: LiF LWRS SIA Tables used during the mission


6.3.4  Single Event Upsets

The electronics for the two FUSE detectors were turned on six days after launch, on June 30, 1999.  On July 2, the Detector DPU code reported a change in the checksum value calculated by its Cyclic Redundancy Check (CRC) routine. Since this CRC calculation was done over a range of internal detector memory that contained the executing DPU code, this meant that the code had been corrupted. The following day, a similar error was reported on Detector 2, but neither had any apparent effect on the detector performance. The third reported CRC error, however, caused a loss of some Detector 1 telemetry. Only by rebooting the detector was the telemetry restored. These errors continued to be reported at a rate of one every few days throughout the mission. Their cause was later determined to be single event upsets (SEUs) in the DPU memory, due to radiation from the South Atlantic Anomaly (SAA). This susceptibility to SEUs was due to the choice of RAM in the DPU; although the memory chip chosen had a high threshold to total radiation dose and was immune to latchup, it had a relatively low resistance to Single Event Effects (SEEs).


Since the regions of memory being corrupted contained the DPU code that was controlling the detectors, this was potentially a very serious problem. Although the detector software was robust and it was believed to be unlikely that a single bit flip could cause any permanent damage to the detector, as a precautionary measure rules were developed to minimize the danger. An intensive investigation was conducted to formulate a plan for responding before the high voltage was first turned on. In order to minimize the risk to the detector, the decision was made to only operate the high voltage when a code image was loaded into both Lower Core (LC) and Upper Core (UC) memory. If an SEU occurred in the active memory region, a command was issued to jump to the alternate core on the next ground station pass, and an uncorrupted version of the code reloaded from the ground. If the error was in the inactive memory core, it was simply reloaded to repair it. If both cores were found to be bad during a ground station pass, the detector was rebooted, which shut down the high voltage on that detector. Since recovery from a detector reboot required close to 24 hours because the ramp-up steps were commanded from the ground, this was a major cause of inefficiency early in the mission. In addition, the IDS flight script det_hv_f was modified to reload several important detector parameters (SAA count rate limits, SAA integration times, SAA HV levels, SAA reduce time, HV current limit, and AUX current limit) after every SAA passage.


This procedure was performed manually until 8 January 2000, when version 16600 of the DPU code was loaded onboard, along with modified IDS scripts that autonomously detected and repaired the code. Although the code was initially loaded from IDS memory, starting in April 2000 IDS EEPROM memory was used to free up memory for IDS scripts.


Before the autonomous code repair was implemented, 122 SEUs had been detected on the two detectors and repaired manually. After these changes, the autonomous correction code permitted the IDS to identify an SEU and correct it autonomously within three minutes in most cases.


There were 1285 SEUs detected on the two detectors during the mission. These were divided as shown in Table 6.3‑13. The fact that they were roughly evenly distributed between Upper Core and Lower Core memory, and Detector 1 and 2, shows that the susceptibility was due to the choice of parts rather than a single bad chip (note that UC and LC for a particular detector were located on the same memory chip). No obvious correlations of the SEU rate with time were seen, aside from changes in the rate due to modifying the size of the memory area checked.


Detector Lower Core Upper Core
1 335 271
2 299 343

Table 6.3‑13: SEUs by detector and memory core.


Although most of the SEUs had no noticeable effect on the operation of the detectors, the consequences could be more severe when certain parts of the code were affected. In some cases, a detector watchdog reset occurred, causing the detector to reboot. Other times unusual behavior was seen, such as a loss or corruption of detector telemetry. Those cases usually required rebooting the detector from the ground to recover.


As expected, nearly all of the SEUs occurred while the detector was in or near the SAA. Figure 6‑11 shows the location of FUSE in its orbit when each SEU occurred. The figure shows that they were concentrated near the center of the SAA, and that the spatial distribution was similar for the two detectors. Only fifteen of the 1238 SEUs for which the location could be identified did not fall within the SAA contour shown.


The CRC calculation took up to 40 seconds to identify a bit flip, depending on where it is in memory. In addition, CRC values and diagnostics were reported in telemetry only once every sixteen seconds. As a result, the times reported are likely to be somewhat later than the actual time of the SEU, and the locations marked on the figure will be slightly misplaced from their actual position.


Figure 6‑11: Positions of SEUs during the mission. The dashed line shows the SAA region used by Mission Planning after 17 September 2003. No events occur below -25 due to the orbit of the satellite.



In addition to the DPU code, detector masks (Section were stored in the same memory chips, and thus were also susceptible to corruption. The CRC check done by the DPU code only checked mask memory when a mask was changed, however, so it was not possible to tell when this occurred. As a result, there were instances early in the mission where a mask bit was switched off due to an SEU, causing a loss of events in that region of the detector. Once this effect was indentified, the observation scripts were modified to reload the masks before every observation. This greatly decreased the likelihood of this problem.


6.3.5   High Voltage Transients (Crackles)

During satellite-level testing, the detector high voltage power supply current periodically had excursions that were large enough to potentially cause damage to the detectors. To minimize the likelihood of a detector failure on orbit due to one of these high voltage transients, or crackles, UC Berkeley modified the DPU software to monitor the high voltage power supply currents on segments A and B (HVIA and HVIB), and the Auxiliary Power supply current (AUXI), at a sample rate of approximately 1 millisecond. After these modifications, whenever these currents were greater than or equal to a threshold (a mini-crackle), a diagnostic was issued by the detector DPU. If this threshold was exceeded for a particular length of time (the persistence), the high voltage for both segments of the affected detector would be turned off and an additional diagnostic issued.


In addition, a portion of DPU memory was set aside for use as a circular buffer containing the last 1024 samples (approximately 1 second) for each of these three currents, along with three histograms that stored their distribution.


The thresholds could be set via ground command, but because the maximum current of most crackles was so high, changing the threshold had a limited effect. The persistence, however, could only be modified by changing the DPU code. Each change required modifying the code followed by extensive testing. Section describes the DPU code versions used during the mission, along with the thresholds and persistence values for each.


Figure 6‑12 shows the three currents during a typical crackle event, as saved onboard in the circular buffer. Although the details of each crackle event were different, the basic features were similar in most cases:

(1)   One of the currents (in this case HVIA, the high voltage current on segment A), showed some oscillatory behavior and then went off scale.

(2)   The AUX current followed a similar shape as the segment current, but had a much lower value. Since it did not reach its maximum, it often showed the oscillation more clearly.

(3)   The current draw of the other detector segment (HVIB) was also affected.

(4)    When the (HVIA) current eventually exceeded the threshold (182 in this case) for the persistence time (60 msec), the high voltage shut down, and the currents all dropped to zero.

Figure 6‑12: HVIA, HVIB, and AUXI during a crackle.


Figure 6‑13: Number of crackles and mini-crackles vs. time.

Table 6.3‑14 summarizes the numbers and types of crackles and mini-crackles for each detector segment during the mission, while Figure 6‑13 shows their distribution as a function of time. Recall that mini-crackles had no effect on operations, while crackles shut down the high voltage on both segments of the affected detector, which required a high voltage ramp-up from the ground. Since the thresholds and persistence values changed during the mission, the crackle rates at different times cannot be compared directly, but it is clear that the incidence of crackles is episodic, for reasons that were never understood. It is also thought that they were more likely to occur soon after a high voltage ramp-up.


Detector Mini-Crackles Crackles
HVIA (27) HVIB (28) AUXI (16) HVIA (2A) HVIB (2B) AUXI (2E)
1 695 67 114 6 18 0
2 456 0 239 75 0 8

Table 6.3‑14: Number of mini-crackles and crackles during the mission.
The values in parentheses are the diagnostic value (in hex) issued by the detector.


Crackle shutdowns were the most common reason that the high voltage on one of the detectors was not at FULL as planned during an exposure. Since the time to return the high voltage to its nominal value after a shutdown could be a day or longer, this was a major source of inefficiency. Plans to automate the high voltage ramp-up process were developed, but had not been implemented by the time the mission ended.


Similar current protection scheme is used on the GALEX and COS detectors, but the DPU code for those detectors provided more flexibility on setting thresholds and persistence values.


6.4           Fine Error Sensor Cameras

6.4.1   FES CCD Detector


The CCD detector was a modified version of a SITe SIA003A. This was a 1024 1024 pixel thinned backside-illuminated CCD die mounted on a 2-stage thermo-electric cooler (TEC) and sealed in a kovar package with a fused silica window.  The CCD quantum efficiency, full-well depth, charge-transfer efficiency, dark current, and readout noise were all in accord with pre-flight measurements.  Basic characteristics of the CCD are given in Table 2.7‑2.


Half of the CCD die was masked, so that the CCD could be operated in frame transfer mode. After a given exposure duration, the image would be transferred rapidly to the region under the mask and then read out slowly so that the readout noise would not affect the accuracy of the centroid calculation.


CCD thermal control was achieved through the use of an external radiator in conjunction with a thermo-electric cooler (TEC). The CCD package was coupled to the external radiator by means of a heat strap fabricated from flexible copper braid that was fused at either end into a copper block. In general, the spacecraft attitude was maintained so that the radiator of the active FES was shaded from the sun.  During in-orbit checkout the TEC setpoint was chosen to be -32C for FES-A and -30C for FES-B. Lower setpoints could be maintained at some attitudes, but the reduction in dark current did not justify the operational inconvenience of tailoring the setpoint for each attitude. With the advent of two-wheel and one-wheel operations later in the mission, which necessitated a much wider range of spacecraft roll angles, on-board scripts were developed that would autonomously adjust the TEC setpoint in response to the changing thermal environment. The resulting higher dark current and slightly-degraded FES performance was accepted, as the pointing performance was then limited by the torquer bars and not the FES.

6.4.2   Performance and Anomalies  FES-A Failure

The first in a series of intermittent spontaneous FES-A reboots occurred on day 105 (April 15) of 2005. When the reboots occurred, FES-A was left in Boot mode, and could not be commanded to application mode without first cycling the power.  In addition, the FES would frequently not recover unless it was left powered off for several minutes to an orbit. This led to the suspicion that some component in the controller had an intermittent problem with heat-sinking. We attempted to mitigate this problem by lowering the temperature of the FES-A surroundings and by having the IDS flight scripts autonomously handle the power cycling and rebooting sequence. We operated in this mode for several months, but ultimately decided to switch to FES-B on July 12, 2005. FES-A was not used for guiding during the rest of the FUSE mission.


The cause of this problem was not determined. FES-A had been subjected to survival temperatures (-5C), along with the rest of the FUSE instrument, between Dec 28 2004, when the Roll RWA failed, and March 22, 2005, when version T31 of the ACS software was loaded. The reboot problems began a few weeks later. It was suspected that this prolonged cold soak had precipitated the problem.  FES-B Focusing

Following the initial focusing of the LiF2 channel for optimal spectrograph performance in IOC, the PSF of star images in FES-B was found to be 23 times larger than desired (the FWHM was 46 pixels instead of 2 pixels), indicative of defocusing of FES-B.  This caused the light from any given star to be spread over many more pixels than would be the case if FES-B were properly focused, reducing the contrast between the signal and the background. 


Although FES had no internal focus adjustment mechanisms, the focus could be adjusted by moving the LiF2 FPA and/or the LiF2 primary mirror along the Z-axis (focus direction).  If the FPA was moved, the only impact to FUV data in that channel was to reduce the throughput of the MDRS and HIRS apertures.  If the LiF2 primary mirror could also be moved, the spectral resolution of the LiF2 FUV data will be degraded.  FES-B was used successfully with these large star images during periods when the FES-A was being annealed. When FES-B was made the permanent guiding camera, however, the likelihood of guiding problems with faint stars was too large. FES-B was refocused by moving the LiF2 FPA by 400 microns (See section 6.2). The resulting image quality was comparable to that of FES-A.



6.5           IDS: Instrument Data System

6.5.1   IDSACS Interface

The IDS was responisble for managing FES operations, for attitude determinations based on FES data, and for initiating slews. In essence, the IDS and FES were an integral part of an extended attitude control system architecture. Communications between the ACS and IDS were handled by Fine Pointing Data (FPD) packets (IDS to ACS) and ACS status messages (ACS to IDS) transferred across the SDB. Measured quaternions, covariances, and configuration flags were sent to the ACS once per second in FPDs for fine pointing management. In return, the ACS provided status messages back to the IDS.


The spacecraft was responsible for the health and safety of the satellite. Therefore the ACS was free to reject individual FPDs or ignore the IDS altogether. The IDS did not command the ACS, but changed the status of FPD flags to in effect request ACS activities. The ACS monitored state changes in the FPDs and responded accordingly.

6.5.2   Observation Sequencing and Fine Guiding

Slewing and target acquistion was controlled by scripts executing in the IDS. The scripts controlled sequencing of commands, while complex functions such as processing of FES images or attitude determination were performed by the embedded guidance task software in the IDS.  Except for centroid computations performed by the FES when in tracking mode, all fine attitude calculations were performed by the IDS. 


The main steps in a basic target acquisition sequence were as follows:   

  1. The IDS sends a new attitude to the ACS, initiating a slew, and then waits for the slew to complete.
  2. The IDS commands the FES to obtain & deliver a full FOV image
  3. The IDS processes the image and determines locations of up to 30 stars. 
  4. The IDS commands the FES to centroid 3 to 6 of the brightest stars at 1 Hz, computes quaternions, and transmits them to the ACS. These quaternions are flagged as unknown, because the stars have  not yet been identified. These data were used to maintain a precise fixed spacecraft attitude while the absolute attitude was determined from the FES image. (see Section 8 for a discussion of quaternions).
  5. The IDS compares the objects in the image to the star catalog uplinked for this observation and determines the absolute attitude.
  6. The IDS sends the absolute quaternion to the ACS, which initiates a slew to remove the residual attitude error.
  7. After the slew, the IDS commands the FES to centroid pre-selected stars surrounding the target, at a rate of 1Hz or 0.5 Hz. As each set of centroid data is received by the IDS, it computes quaternions from reported star positions, and sends them to the ACS for input to the attitude-determination filter.
  8. At the end of the target visibility period, the IDS commands the FES to stop tracking stars, and the IDS stops generating FPDs. The ACS maintains coarse tracking using gyros and TAMs.
  9. For multi-orbit observations, return to step 2 above when the target is next unocculted by the Earth, and the satellite is not passing through the SAA.


More complex acquisition procedures were needed for some observations. The main variations are described briefly here.


Offset acquisitions. In cases where the star field in the vicinity of the target was too sparse for the star-ID process to work, or if there were a very bright source in the FES field of view, the first five steps of the acquisition sequence above would be performed at an offset field. After the field was identified, the correction slew request would include the offset needed to move directly to the desired target position.


Peak-ups. Observations using the MDRS or HIRS apertures would include a peakup procedure folowing step 7 in the basic acquisition sequence. Peakups are described in Section 4.2.2.


FES-assisted acquisitions. In these acquisitions, the target would be left at the reference point position, away from the apertures, and the FES would be commanded to measure its position as well as that of the guide stars. The IDS would then compute the actual position of the target and then command a short slew to place the target in the proper spectrograph aperture. This mode was initially envisioned for objects such as comets that might have poorly-known coordinates, but in practice is was often used for routine observations.

6.5.3   IDS Thermal Control  Thermal Design

The IDS was responsible for the operation of the instrument thermal control system used in normal operations. Temperature data from instrument thermistors were collected from the IPSDU, averaged, and then used to control 32 heater zone switches in the IPSDU. Overall management was specified by a user-defined thermal control table. The table identified which thermistors were associated with each thermal zone and the upper and lower temperature thresholds for the zones. Up to 4 thermistors were monitored for a zone and, when the average temperature fell outside the zone thresholds, the heater was turned off or on. By executing one full cycle of the algorithm every 16 seconds, the IDS guaranteed that the temperature of each thermal zone was maintained to +/- 0.25 C of the desired temperature. Space for two tables was provided to facilitate the loading of a new table while under thermal control of another.  Thermal Performance

The deadbands in the thermal control tables were set at +/-0.25 C. The control system was able to maintain temperatures within this deadband, with the exception of the equipment panels which were controlled to within 1C, which was adequate for the electronics modules mounted on these panels.

Under nominal three-wheel operations instrument temperatures were well-controlled about a daily mean temperature that exhibited no long term trends.  In particular, the mirror bench temperatures on the shaded (+X or LiF) side of the instrument exhibited no change in daily mean temperature over the mission. During the subsequent two-wheel phase with roll-offsets the temperatures were also, in general, well-controlled.   However, for some of the heater zones in the lower sections of the instrument a gradual increase in the daily mean temperature was evident - including the SiC1 mirror bench.

The behavior of the zones in the upper half of the instrument was slightly different. This includes the spectrograph structure, the GMA shrouds, and the gratings. Each of these zones maintained constant or nearly constant temperatures until the pitch wheel failure in December 2001. Following restoration of 3-axis stabilization, the mean temperature in each of these zones was slightly (<0.1C) higher.  The temperatures of these zones exhibited a few slight mean temperature changes/offsets later in the mission and the SiC grating temperatures increased gradually by 0.05C over the last few years of the mission.

The LiF grating temperatures were well-controlled. The combination of controlling the grating pie-pans in conjunction with their thermal mass was successful in providing a stable environment for the optics.  Under normal operating conditions the grating temperatures were controlled to within 1C.  However, the grating temperatures, in particular for the SiC temperatures, varied with solar beta angle.

The zones within the instrument were controlled very well throughout the mission. The slight changes that were seen were presumably due to minor degradation of the multi-layer insulation with exposure to solar UV radiation and atomic oxygen. The distinct step in temperatures seen at the time of the roll wheel failure in particular may be the result of such exposure, as this was the first time in the mission that exterior surfaces of the instrument not covered by silver teflon blankets were exposed to the Sun for an extended period.

The temperatures of the telescope baffles were not controlled, as there were no functional or survival requirements to heat the baffles. The baffle mounting to the structure was designed to decouple mechanical stresses from differential thermal contraction of the baffles with respect to the structure. However, the significant variations in telescope alignment found on orbit that correlated with the Solar beta angle and orbit pole angle, and with orbit phase, suggests that this decoupling was not adequate. The temperatures of the door closure and unlatch HOPs, located near the top of the baffles, are seen to vary on orbital timescales by 4-5 C on the LiF side and by 6-8 C on the SiC side.  These temperatures vary with spacecraft attitude as well: typical variations early in the mission were 3-5 C on the LiF side, and 12-15 C on the SiC side. Typical temperatures at the tops of the baffles were 40 C and below on the LiF side, and 20 to -40 C on the SiC side.

6.6           Instrument On-orbit Performance

6.6.1   Telescope Focus  Post-launch Focus Assessment Details


The FUSE instrument was focused pre-flight with the provision for in-flight focus adjustments for the mirrors and FPAs.  On-orbit focus adjustments were expected due to (1) the unavailability of a laboratory FUV source with a light beam collimated to roughly one arcsecond accuracy and (2) the anticipated focus changes associated with gravity release and changes in the positions of the optical elements resulting from moisture desorption from the optical bench structure.


The on-orbit instrument focus procedure was essentially a two-step process. First, the telescope was focused by adjusting the mirror to FPA distance for each channel. Then each spectrograph was focused by adjusting the distance from the telescope mirror to the spectrograph grating for each of the four instrument channels.  The FPAs were then re-adjusted to maintain the previously determined telescope focus.


The telescope focus was determined through a series of knife-edge tests performed by scanning a target across the edge of  the FPA slit. The knife-edge test was repeated for a set of FPA positions to determine the location along the optical axis where the light cut-off was sharpest.  Then the FPA for each channel was moved to place the aperture at the best telescope focus as indicated by this test.  The FPA motions executed to attain the telescope focus for each channel  are presented in Table 6.6‑1.



FPA FPA Motion
LiF1 no change
SiC1 -99
LiF2 +48
SiC2 -81

Table 6.6‑1: Initial in-flight telescope focus adjustments made November 23, 1999.  Adjustments in the focus (Z) direction are limited to 10 micron increments of the FPAs.  Small residual errors account for the slight departures from integral 10 micron changes in the adjustment values above.  The true uncertainty in the magnitude of the computed focus adjustment was at least 30 microns.


Two programs were executed to determine the spectrograph focus. The first of these programs, I817, was executed during the December 7 December 9th, 1999 time interval. Multiple stellar spectra of HD208440 were acquired through the LWRS aperture for each of 5 mirror positions stepped in 150 micron increments along the optical axis.

Data from program I817 and the earlier knife-edge tests enabled a robust determination of the best grating to mirror distance (the spectrograph focus) for the LiF1 and LiF2 channels. However, the signal-to-noise of the I817 data was relatively low for the SiC channels resulting in a less robust determination of the best spectrograph focus position for the SiC1 and SiC2 channel mirrors.


On December 12, 1999 the mirrors and FPAs were adjusted to the best spectrograph focus for each channel based on the data obtained December 7th-9th. LiF2, SiC1, and SiC2 previously had their FPAs adjusted to achieve the best telescope focus (mirror to FPA distance) as determined from the knife edge testing. For these channels, both the FPA Z position and the mirror focal distances were adjusted to focus the spectrograph and maintain the best mirror-to-FPA distance (i.e. the telescope focus) determined by the knife edge scans. For LiF1, the adjustment of the mirror location for spectrograph focus matched the required adjustment for mirror to FPA focus; hence the mirror only was moved. The FPA and mirror position adjustments made on December 12th 1999 are presented in Table 6.6‑2 for each channel.


Channel Initial
Dec 12, 1999
Z Mirror
Z Mirror
Z Mirror
Dec 12, 1999


64 0





















none; -150





Table 6.6‑2: Spectrograph focus adjustments executed on December 12th, 1999 as a result of the I817 post-launch programs.


Given the low signal-to-noise of the spectrograph focus measurements conducted as part of program I817, a second spectrograph focus test was executed two months later as part of program I819. The target, WD0439+466: the central star of a planetary nebula, was observed in the LWRS slit. The spectrum exhibited many narrow molecular hydrogen lines, which were nearly ideal for focusing.  Quality data combined with careful analysis resulted in a much better measure of the spectrograph focus. On March 16, 2000, the spectrographs were brought to their best focus position by moving the telescope mirrors along the optical (Z) axis. The magnitude and direction of these motions are provided in Table 6.6‑3. These motions were generally in the opposite direction from the first spectrograph focus adjustment, but since the data were of significantly better quality, they were considered far more reliable. Corresponding adjustments were made to the FPA Z positions on March 24th 2000 to maintain the mirror-to-FPA separation for LiF2, SiC1, and SiC2; the LiF1 FPA Z position was adjusted by -100 microns (see Table 6.2‑1 for the history of FPA Z motions). These focus positions remained the default throughout the remainder of the mission.

Mirror Z Mirror Motion
(microns) Executed
December 12, 1999
Z Mirror Motion
(microns) Executed
March 16, 2000
Net Mirror Focus



No change














Table 6.6‑3: Spectrograph focus adjustments executed as a result of the I817 and I819 post-launch programs.
The spectrograph focus adjustments implemented March 16th, 2000, were used for nominal operations for the remainder of the FUSE mission.

6.7           Telescope Alignment Performance

During the IOC shortly after launch, it was realized that the telescope mirrors underwent periodic motions that shifted the targets image at the telescope focal plane and thus its spectrum in both X and Y on the detector. A source in either of the SiC channels could move as much as 6 arcseconds in a 2 kilosecond time interval. This motion had two effects on the data: first, flux was lost if the source drifted (partially or completely) out of the aperture; second, the was spectrum shifted on the detector, degrading spectral resolution for observations using the LWRS (30 arcsecond) aperture.


6.7.1   Initial Alignment

The procedure for the initial post-launch channel alignment was to perform a spiral search while guiding on the target using FES-A with LiF1 and the LWRS aperture. The FUV count rate was monitored while using the mirror actuators to locate the target position in the three non-guide channels. This method required use of a moderately bright FUV point source in a fairly isolated star field to provide sufficient counts and prevent confusion from nearby field stars.  Overall, this method worked well and was used to initially align channels to a 15-arcsecond  precision.

The data were analyzed to determine the mirror rotations required to remove the measured alignment errors at that attitude for each of the three non-guiding channels relative to the LiF1 channel.  As further alignments and observations were attempted, it was found that all of the channels were moving with respect to each other, with misalignments as great as 40 arcsec.

Finer co-alignment was then performed by stepping the roughly co-aligned images across the edge of the LWRS slits.  This method allowed for co-alignment to approximately 2 arcseconds.  Early attempts to use this method were hampered by the fact that the measurement and adjustment were not closed loop.  Scans were made and mirrors adjusted, but no confirmation of the image position was made.  FUSE would then slew to another target, which often resulted in co-alignment being lost due to changing the thermal environment of the instrument.  Early in the mission the source of the misalignment was unknown and the lack of a closed loop mirror adjustment initially masked the thermal misalignment problem.

6.7.2   Mirror Motion Anomaly

The FUSE  Image Motion was defined as the variation of the alignment of the SiC1, LiF2, and SiC2 channels with respect to the reference LiF1 channel used for guiding.  Although these motions were generally quantified as being associated with the non-guiding channels, they were caused by the motions of LiF1 in combination with the other channels.  The observed image motion was attributed to the thermal changes induced in the instrument by changes in boresight beta angle and orbital pole angle.  The boresight beta angle is the supplement of the angle between the satellite-sun line and the instrument line of sight and the orbital pole angle is the angle between the orbit pole and the instrument line of sight.

To mitigate loss of data due to channel misalignment while retaining observing efficiency, the LWRS aperture became the primary observing aperture for most programs.  For MDRS and HIRS observations, two peakups per orbit were executed to maintain channel alignment during an observation and obtain full spectral coverage while using these apertures.

6.7.3   On-orbit Mirror Motion Mitigation Strategy

A two-part strategy was developed to maintain alignment of the four mirror channels.  This strategy included both a predictive component based on empirical modeling of changes in mirror position as a function of boresight attitude and a periodic re-baselining of the alignment.  Mirror Alignment: Baseline Maintenance

The re-baselining component of the alignment maintenance was achieved by performing xy-scans on a stellar target and performing near real-time analysis to determine the direction and amplitude of the corrective mirror motions required to attain alignment at that particular attitude. The alignment baseline was valid for a range of beta and orbital pole angles in the vicinity of the target.  Alignment within a particular attitude range could be maintained for approximately two weeks, assuming that targets were scheduled within the allowed beta/pole zone.  Approximately every two weeks, xy-alignment scans were performed, the data analyzed, mirror motions calculated and executed to remove the accumulated thermal and temporal drifts and maintain the mirror co-alignment at the current attitude.  For targets within a beta/pole range with a different thermal environment over an orbital period, xy-alignment scans were required to establish the mirror co-alignment baseline at that attitude.  It was also necessary to establish an alignment baseline for every hemisphere crossing of the satellite.  Mirror Alignment: Predictive Modeling


Given the sensitivity of the thermally induced mirror motions to the boresight attitude, maintaining channel alignment by performing xy-scans for each target would have been prohibitively inefficient.  As a result, a model was generated that predicted the motion to be executed for each mirror to maintain channel alignment for each target observed.  This predictive model was part of the planning and scheduling process.  The mirror motion scripts were uploaded to the satellite and executed as part of the command sequence for each observation. 

This predictive model for the corrective mirror motions was derived from the analysis of xy-scans taken at multiple attitudes in combination with the analysis of peak-up data using the 4-arcsecond wide MDRS and the 1.25 arcsecond wide HIRS aperture.  The xy-scan data provided the magnitude and sign of the target mis-alignment in both the x and y direction. The peak-up data provided sign and amplitude information in the x-direction, but no y information was available.  In addition to the dependence of the mirror mis-alignment motion on spacecraft attitude, temporal dependencies were also found but not completely nulled by the predictive model.  Channel Alignment Operations


Empirically, it was determined that the secular drift of the predictive model relative to observed target position in the different channels required obtaining xy-scans every two weeks to re-establish the model baseline. Extreme attitude changes, such as hemisphere crossings which were driven by the sun angle and typically resulted in large beta changes, were not well-modeled and required xy-scans at the target attitude to determine the appropriate mirror motions for channel alignment. In addition, mirror motions were a strong function of attitude for targets at large beta angles


Executing the scans, analyzing the data, and uplinking the corrective mirror motions impacted observatory efficiency. Since the channel misalignment resulted from thermal changes in the instrument environment local to the mirrors, it was imperative that the mirrors were in thermal equilibrium prior to executing the xy-scans.  Thermalization periods of 6 -12 hours were required depending upon both the beta angle and the magnitude of change in the beta angle between targets.  Analysis of scans taken with extreme attitude changes exhibited an exponential dependence of the target (i.e. mirror) position as a function of time, consistent with the thermal non-equilibrium hypothesis. Achieving full positional equilibrium could require 11 hours; however, the target was usually within the 30 arcsecond wide LWRS aperture and monotonically approaching equilibrium within 6-8 hours.  As a result, to improve efficiency, observations were typically initiated before completion of the full thermalization period.  For time-tag observations the data are processed by the CalFUSE pipeline as photon lists.  Signal-to-noise permitting, this facilitates the comparison of data in temporal bins to remove secular drifts from the data.

The image motions for the two SiC channels were primarily in the X  (dispersion) direction, although small corrections in Y were also occasionally required.  For the LiF2 channel, the motion was exclusively in Y (perpendicular to dispersion), and significantly larger than the Y motion in the SiC channels.  Y motions required corrective mirror adjustments. In the X (dispersion) direction the FPA was used to accommodate small changes in the predicted position of the target because FPA adjustments were simple to implement and resulted in no loss of observatory time.   Due to the limited travel range of the FPAs, larger X motions corrections required mirror motions to achieve co-alignment.

Analysis of the  xy-scan data and comparison of signal levels obtained in each of the channels during routine observations enabled the formulation of a set of rules to maintain alignment between alignment re-baselineing observations.

These predictive model and alignment rules were incorporated into a planning spreadsheet that was executed prior to completing the detailed scheduling of targets. The primary alignment considerations were to restrict the changes in beta and pole angle to within 30 degrees of the previous target.  In addition, large attitude changes required extending the observation time on a target to include the thermalization time for that attitude and delta attitude to ensure obtaining full spectral coverage, if needed.  Hemisphere crossings and observations at high beta angle required long thermalization times and dedicated alignment activities.  Long-term temporal dependencies were not well-modeled necessitating the need for re-baselining  alignment scans approximately every two weeks.

Alignment activities were required at both low and high beta angles.  Motions of the LiF2 mirror at high beta angle (>90 degrees) were not well modeled an the consequently observations requiring LiF2 data were not executed at high beta angles without a dedicated alignment activity.  Due to the paucity of data at high beta angles, we were not able to improve the predictive model in this regime.  Pole angle was a less important consideration than beta angle for maintaining alignment.  Observations at a single beta angle for a large range of pole angles supported this conclusion. 

Later in the mission, planning constraints necessitated by operational changes arising from failures of the FUSE reaction wheels and associated with spacecraft safety took precedence over alignment constraints. This periodically resulted in either increased re-baselining activities to achieve mirror co-alignment, which was operationally inefficient, or acceptance of limited spectral coverage as a result of mis-aligned channels.  To ensure that observations were acquired with the necessary spectral coverage, Guest Investigators were required to specify the channels essential for their science program. This information was propagated through the MP database and factored into the scheduling and detailed sequencing of observations.  Orbital Motion

Image motion was also found to have several time dependencies.  A very repeatable orbital motion was found, and was well mapped for CVZ targets (Figure 6‑14 This source of this motion was attributed to the thermal cycle of the orbit.  For some targets, a 24-hour period motion of as much of 8 arcsec full amplitude was observed.  This motion appeared to be roughly sinusoidal in shape. There were also apparent long-term time dependences for some of the motions, but a paucity of continuous data inhibited characterizing this temporal dependence.

Figure 6‑14: Orbital Dependency of Image Motion for Selected Targets  Impact and Evolution of Image Motion Corrections

Image motion impacted FUSE science data in several ways.  (1) Channels were misaligned for targets at beta and pole angles near the edge of the range of nominal operations.  To mitigate this, the predictor set was continually expanded and improved.  As a result, channel loss decreased over time.  (2) Re-alignment activities reduced the time available for science data collection.  Thus, realignment activities were streamlined to use significantly less time than required early in the mission. (3) Motions of the spectral image could result in the loss of spectral resolution, especially for long exposures at high and low beta angles.  Consequently, short histogram exposures (<500 s) were used for bright targets. Time-tag data is corrected for this motion in data processing using an empirically derived correction.  A more detailed discussion is presented in the the CalFUSE pipeline paper (Dixon, et. al 2007). The bright limit for time-tag data was increased to allow more targets to use this mode.  (4) Tracking orbital motion for medium and high resolution aperture observations required additional FPA motions, and impacted efficiency.  Target Peak-up Strategy for MDRS and HIRS Observations

To maintain co-alignment of the channels at the arcsecond level required for the MDRS and HIRS apertures, an on-board peak-up procedure was employed.  This procedure slewed the telescope in small steps to move the images across the narrow aperture.  The FUV count rate at each point was used to calculate the position of the image centroid for each channel.  Next, FUSE was slewed to the peak position for the guiding channel.  Then, each of the other three FPAs were moved to the calculated location for each of their peak count rates.  In performing this operation, it was quickly determined that the images moved over an orbit while pointing at a single target.  Consequently, procedures were developed and implemented to perform peakups twice per orbit to maintain alignment when using the narrow apertures. The inclusion of mid-orbit peakups resulted in reduced observing efficiency, but enabled full spectral coverage using the MDRS and HIRS apertures.

6.7.4   Channel Alignment: Observations and Analysis

As discussed in Section, the real-time channel alignment process was time consuming. Large overheads were incurred as a result of the time required to thermalize at the target attitude, execute the scans, wait for ground station passes to downlink the data, perform the scan analysis, generate the mirror motion scripts, and then uplink the corrective mirror motions at the next available pass opportunity. This need for ground station communications for data downlink was both a time and a scheduling constraint. To eliminate overheads and streamline this process as much as possible, on-board detector counter data was acquired instead of science exposures.  Figure 6‑15 depicts  xy-scan counter data used to analyze an observation.  The first peak in the counts/second (occurring at between 400-100 seconds) is the y-scan; the second peak is the x-scan. The four channels are each graphed using a different color. From inspection of the graph, all channels are aligned in the y-direction.  However for the x-scans, the SiC data (blue peaks) are offset from the LiF1 guide channel data indicating a  misalignment of the SiC1 and SiC2 mirrors.

Figure 6‑15: xy alignment  scans as depicted using the Alignment Tool Graphical Analysis.  These data are discussed in Section 6.7.4.

With the loss of reactions wheels, the pointing jitter increased. Now, one of the primary observation scheduling constraints was the availability of torque authority, and, in one-wheel mode, using target sequencing to manage the spacecraft momentum. Minimizing the changes in beta angle and pole angle between subsequent observations became a secondary consideration. This change in planning strategy resulted in larger predictive mirror motions and larger deviations of the mirror position from the predicted position.  As a result of the larger alignment errors and the increased difficulty in executing maneuvers in one-wheel mode, the alignment scan pattern was changed both in its spatial coverage and from a step and dwell pattern to a series of continuous scans. This in turn necessitated a new alignment assessment tool, the Channel Alignment Tool (ChAT) shown in Figure 6‑16. In addition to the xy-scan pattern, diagonal scans were also executed to increase the diagnostic capability in one-wheel mode where the amplitude of pointing excursions was much larger than previously experienced when pointing was controlled using three reaction wheels.  The erratic scan trajectories in Figure 6‑17 and Figure 6‑18 exemplify the difficulty of fine-pointing with only a single wheel.

The graphical interface provided by ChAT for interaction with the alignment scan data greatly facilitated analysis of the scans.  Regions of pointing excursions and high background were now easily identified and excised from the alignment analysis. 

Figure 6‑16: Example of Channel Alignment Tool (ChAT) Results

Figure 6‑17: (Revised) One Wheel Mode Alignment Scan Pattern

Figure 6‑18 One Wheel Mode Alignment Scan Pattern Results

Figure 6‑19: Additional ChAT Sample Results

Confirmation of the successful execution of the commanded mirror motions was inferred by one of four methods: examination of the command sequence to verify that it had completed properly, execution of another set of X- and Y-scans, analysis of a follow-on peak-up activity, and/or analysis of the count rate of the next sufficiently bright target at a comparable attitude. Depending upon the availability of ground station contacts, addition scans or peak-ups could require several hours or more to perform the scans and downlink the data. Consequently, a confirmation of mirror motion execution was rarely performed.  In practice, confirmation of the count rate from the next bright target was the primary method used to confirm that the mirrors had been moved to their desired positions.

Examination of the telemetry over the eight year lifetime of the mission never revealed an instance where the mirror motion command only partially executed.  With one exception, the infrequent instances where the mirror motion did not appear to execute could be traced to a manual tracking error somewhere in the alignment maintenance process. Consequently, steps were taken to automate the procedure as much as possible.

For the first five years of the mission, LiF1 was the guide star channel and the LiF1 mirror actuators were not moved.  In 2005 due to anomalies with FES-A (see Section 4.5.3) LiF2 became the guide star channel.  The restriction on moving the LiF1 mirror was then lifted, and LiF1, SiC1 and SiC2 mirrors were moved to align with LiF2.

6.7.5   Mirror Motion Accuracy

There was no explicit test of the accuracy of the mirror motions and the mirror motions were not telemetered.  It could be inferred from multiple realignments that the inaccuracies of the motions are small compared to both the requirements and measured orbital motions.  Peakup test data showed motion accuracy about Ry for the SiC 1, LiF 2, and SiC 2 mirrors of better than 2 arcsec for small (< 6 arcsec) motions.  Moderate size motions (<30 arcsec) have been shown to have accuracies of better than 4 arcsec. This had significant positive impact on operations after the loss of the reaction wheels, when prioritized scheduling as a function of spacecraft momentum management was needed.  This required additional mirror motions for a greater number of observations and hence, greater flexibility in the mirror motion scripts.  Mirror Motion Tracking and Actuator Performance

Due to the thermally-induced relative motions of the four telescope mirrors (LiF1, LiF2, SiC1, and SiC2) with respect to one another (Sections 4.1.1,, and 6.7.2 - 6.7.4) the total number of mirror motions executed was much higher than the ~900 motions per actuator that was predicted before launch.

The actuators were designed for 1.0 × 106 motions, far more than could reasonably be done in even a very long extended mission. As a result, although the number of actuator motions for each mirror were tracked, there was no cause for concern in their usage for channel alignment.  Furthermore, any loss of a single actuator on any number of mirrors is a soft failure that will cause a minimal loss of data. 


Figure 6‑20: Time series of individual mirror motions executed to maintain co-alignment of the LiF channels.

Figure 6‑21: Time series of individual mirror motions executed to maintain co-alignment of the SiC channels.


The mirror actuators were not run over their entire range of motion in flight.  However, the defocus tests, which were performed to assess the feasibility of executing bright target observing programs, did move the actuators to within 20% of their full range of travel for the LiF2, SiC1, and SiC2 mirrors.  The LiF1 mirror was not defocused, since as the guide channel at that time, the risk was not justified.  No adverse effects were noticed and the mirrors returned to their nominal in-focus positions as commanded.

The secular  drifts of the actuator position associated with mirror co-alignment over the lifetime of the mission is shown in Figure 6‑22 and Figure 6‑23.  The position of the actuators for each mirror, the overall repeatability of the system, and the ability of the model to keep the mirror motions within a well defined range is illustrated. The use of LiF2, rather than LiF1, as the guide channel in early 2006 is evident from the plots.  In April 2006, the predictive model became less robust as is indicated in the frequent use of larger actuator motions to obtain co-alignment.

Figure 6‑22: The range of motion for each of the Lif1 (top) and LiF2 (bottom) mirrors illustrating that although to co-alignment position for each actuator exhibits a secular drift with time, this change is small and well within the range of travel for each of the actuators.


Figure 6‑23: The range of motion for each of the SiC1 and SiC2  mirror actuators.

6.7.6   Spectral Motion Anomaly: Thermal & Mechanical Analysis

During IOC an investigation team was assembled to determine the cause of the observed spectral motion. The functional form of the motion with orbital phase was suggestive of differential thermal heating of the instrument.  A full FUSE thermal math model was correlated to the thermal balance test data obtained during thermal vacuum test at GSFC (November 1998, January 1999) and input into a structural Finite Element Model (FEM).  This model was then used to infer temperature induced image motions.  Results of the analysis showed that, as expected, the temperature of SiC side of the instrument and spectrograph increased as beta angle increased.  These results were in agreement with flight data depicting the SiC mirror bench and spectrograph heater duty cycles decreasing by 11% and 7% respectively as the instrument's exposure to the sun increased.  LiF mirror bench duty cycles were only minimally affected by sun angle changes.  SiC baffle temperatures at the upper attachment flexure changed as much as 28 degrees Celsius when the beta angle changed from 70 to 105 degrees.  The corresponding change to the LiF side was only about 1 degree Celsius.  Quasi-steady-state temperature variations at these points, caused by orbital influences, were 1 degree Celsius.

Since the spectral motion was first discovered using airglow lines, the possible sources of the motion were isolated to the gratings, FPAs, and detectors.  Because the SiC and LiF lines incident on a common detector behaved differently, the detector was exonorated as the cause of the observed motions. The grating bench was eliminated as the cause of the spectral motion because the amplitude and directionality of the spectral motions are larger in X than in Y and it was expected that any motions caused by the grating bencyh would be larger in Y than in X. Neither were the FPAs regarded as a significant contributor to the spectral motion since observations using the LWRS slit, which is much larger than the image of a typical astronomical source, show similar motions in the spectrum of a point source and in airglow lines that fill the slit.

The FUSE Grating Mount Assemblies (GMA) were also investigated for any potential contribution to the in-flight spectral motions.  The existing thermal distortion analysis was reviewed for orbital variation for the GMAs.  A overview of the design was also performed.  The review of the GMAs was limited to the mount assembly since the FUSE on-orbit performance shows no evidence that the grating glass itself could be distorting.  During the instrument design phase, a thermal model of the GMA was constructed and coupled to the satellite model to predict an on-orbit temperature map for the grating assembly components.  That analysis predicted that for each individual piece part of the GMA, the orbital temperature change laterally in IPCS X and Y was negligible (< 0.02 Celsius).  The orbital temperature change in the mount in the IPCS Z is ~0.2 degrees Celsius, and the worst case orbital temperature change between the inner and outer tube temperatures is 0.2 degrees Celsius.  The FEM distortion analysis showed that translations of the center of the grating are within short-term stability requirements.

However, the model analysis was not regarded as conclusive given the large size of the mount and the relative flexibility of titanium, the primary material of construction. A test simulating the on-orbit environment was regarded as the best verification of the mount's thermal stability.

Consequently, pre-launch tests were conducted to determine if the gratings distorted with temperature.  The gratings were heated using their shrouds while monitoring their spectral images for changes.  The glass temperature was raised from 21 to 28 degrees Celsius while the shroud temperature peaked at ~40 degrees Celsius.  A post-launch re-examination of this data found that the image moved 19 pixels on the detector (~114 microns) for this temperature change in the direction of IPCS +X for the SIC1 channel.  Concurrent observations using the LiF1 channel, without heating its shroud, exhibit only a ~3 pixel motion.  These results and the magnitudes of the motions are supported by a FEM/thermal analysis done for FUSE by the University of Arizona.

The resulting hypothesis was that the direct conductive path from the shroud to the grating mount hardware was heating the "wedge" non-uniformly, possibly leading to a grating rotation.  While not conclusive, this test data illustrates the potential sensitivity of the grating mount assembly to temperature change.

Subsequently, an on-orbit test was conducted to test the hypothesis that the heaters may have been thermo-electrically coupled to the spectrograph GMA.

To evaluate this hypothesis, the two grating shrouds (LiF2 and SiC2) on the +y spectrograph axis were permitted to drift to a lower temperature while the other two grating shrouds (LiF1 and SiC1) were held at their nominal set point of 23 C. A temporary set point of 19 C was chosen for the heaters. This temperature was chosen to permit acquisition of Lyman beta airglow spectra for 2 orbits without the heaters cycling while keeping the gratings at or above the spectrograph temperature, a contamination related issue albeit not a severe concern after ~15 months in orbit.

Although the time required for the GMA shrouds to reach their test thermal set-point was significantly faster than expected, a few conclusions could be drawn. The motion of the spectrum in the dispersion direction was not alleviated by operating with the GMA shroud heaters off.  Changing the temperature results in an offset in the zero point position of the spectrum on the detector in the dispersion direction.  And, the functional form of the relative motion of the spectrum was the same at the lower grating temperature as it is at the nominal temperature. The spectral motion exhibited after the GMA shrouds were returned to their nominal temperature was the same as it was before the temperature was lowered.


Figure 6‑24: LiF2B data over an orbital period illustrating that the spectral motion observed with the GMA shroud at 19 C (top) is shifted/offset by ~5-6 pixels from the data acquired with the GMA shroud at the nominal 23 C (bottom).


6.8           Optical Design Specifications



Figure 6‑25 Optical element layout for LiF1, SiC1 channels.


Figure 6‑26: Optical element layout for LiF2, SiC2 channels.



Figure 6‑27: Optical element layout, top view.


Figure 6‑28: Line drawing of an FPA mechanism, showing the two-axis stage and aperture plate.

Figure 6‑29: Details of optics layout at detector surfaces.

Figure 6‑30: Positions of optical elements.



Figure 6‑31: Side views of instrument, showing the structure, optics, baffles, electronics, and radiators. Left: Y-Z view, Right: X-Z view.


Figure 6‑32: Section views of a primary mirror. Top: "pie-pan" enclosure, intermediate plate and bench are shown, with positions of the actuators. Bottom: mirror dimensions and vertex position are shown.


Figure 6‑33: Close-up view of light paths at the FPA-FES interface.

Figure 6‑34: A section view of a grating mount assembly and grating are shown.

6.9           Dead Time


There were multiple effects which caused photons that reached one of the detectors to be lost before they could be properly recorded. We refer to these collectively as dead time. Corrections for these effects, which are a function of count rate, are made in the CalFUSE pipeline. Figure 6‑35 and Figure 6‑36 show the flow of data through the detector places in the data flow where dead time effects occur. Information on the treatment of dead time effects by CalFUSE can be found in Dixon et al. 2007.


Figure 6‑35: Functional block diagram showing the flow of events through the detector electronics (1 of 2). Counts incident on the detector can be lost (1) at the digitizer; (2) due to counts falling outside the Active Image Mask; and (3) in the Round Robin, which combines the data from two segments on one detector.


Figure 6‑36: Functional block diagram showing the flow of events through the detector electronics (2 of 2). Counts can be lost (4) in the IDS if the count rate is above 32,000 cps (HIST only); (5) due to screening by the SIA table (HIST only); (6) in the IDS if the count rate is above 8,000 cps (TTAG); (7) if the FIFO fills (TTAG); or (8) between the spacecraft and the ground system.



The following effects, numbered to match the figures, are potential contributors to the dead time.




MCP response time: The first of these effects is the response time of the MCPs themselves. If the count rate is too high in a particular region of the plate, local gain sag may occur and subsequent photons may not generate enough charge to be detected by the electronics. The count rates for this to occur, however, are well above the count rates seen during the FUSE mission, so this has a negligible effect on the data.




Detector Electronics Dead Time: The inability of the detector front end electronics to process two events that arrive within a very short time interval leads to the electronics dead time. Analysis of the electronics design by UC Berkeley [memo reference] showed that this effect could be characterized by the following equation:

  Plive = exp(-Rta) / (1+Rtism -x10+x6+x2-4x),



where Plive = (1 Dead time); R is the input count rate, as given by the FEC; and ta and t-clk and tism are constants related to the design of the Input State Machine in the DPU ACTEL.


Analysis of science data for each segment was used to calculate the best-fit values of ta, tclk, and tism in the above equation so that they could be used by the pipeline. The correction applied could then be calculated for each segment of each exposure based on the measured FEC rate. The parameter values in CalFUSE are given in Table 6.9-1, which were based on data obtained early in the mission. A reanalysis using all the data from the mission resulted in updated parameters, given in Table 6.9-2. The differences between th two fits are negligible for most FUSE observations, but can reach a few percent for targets at or above the bright limit. Live-time curves computed from these parameters are plotted in Figure 6-38 and Figure 6-39.


Segment ta tism tclk
1A 2.76 11.68 1.50
1B 4.66 10.28 1.66
2A 5.23 8.37 1.27
2B 4.23 6.31 0.00

Table 6.9-1: Detector deadtime parameter values used by CalFUSE.


Segment ta tism tclk
1A 5.53 9.94 2.25
1B 4.28 14.31 3.11
2A 4.85 8.89 1.08
2B 4.90 11.30 3.45

Table 6.9-2: Updated detector deadtime parameters



Active Image Mask: An Active Image Mask (Section which excluded part of the detector would result in a loss of counts. Since these masks were left wide open for the entire mission, this had no effect.




Round Robin: The Round Robin combined the events from two segments on a single detector into a single data stream for passing to the IDS.




Data Bus Limit (IDS Dead Time): A maximum of 32,000 counts per second could be passed on the data bus in HIST mode. If the sum of the Active Image Counter rates on all four segments exceeded this value, CalFUSE scaled the counts appropriately. Stim lamp exposures were the most common causes of exceeding this limit.




SIA Table Screening: Use of an SIA table limits event storage to specified regions of the detector.




Data Bus Limit (IDS Dead Time): In TTAG mode, the maximum throughput is 8,000 counts per second.




FIFO Size and Drain Rate: A 9 MB (default) region of memory acted as a FIFO to buffer the data stream on the way to the spacecraft recorder. Once this FIFO was full, it could be drained at only ~3500 counts per second.  Thus long TTAG exposures at high count rates had regular data droputs (Figure 6‑35). As a result, TTAG exposures were typically limited to count rates of less than ~2500 counts per second [is this the right number?]




Transmission to the Ground: Data could be lost between the spacecraft recorder and the building of the raw data files.



Figure 6‑37: Apparent count rate as a function of time for exposure Q11401001, which was obtained in TTAG mode, despite having a count rate of more than 100,000 counts per second. For the first ~520 seconds the count rates on all four segments were constant, with their sum limited to the 8,000 cps TTAG data bus limit. For the remainder of the exposure, regular data dropouts appeared as the FIFO filled.


All of these effects have been considered to be independent of time in CalFUSE. There were suggestions in the data that the electronics dead time may have changed during the mission, but the effect, if present at all, is small.


Figure 6‑38: Detector live time calibration used by CalFUSE

Figure 6‑39: Updated live-time calibration curves.




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