25 May 1999 - DRAFT
1 Introduction
2 Instrument Description
2.1 Instrument Overview
3 Mechanisms
3.1 Instrument Cover
3.2 Pointing
3.3 Fine guiding
3.4 Aperture Wheel
3.5 Filter Wheel
3.6 Shutter
3.7 Rotation
4 Systems and Subsystems
4.1 Structure
4.1.1 Cover
4.1.2 Telescope Structure
4.2 Mirrors
4.3 Camera
4.3.1 Clock drivers
4.3.2 Video preamplifiers
4.3.3 Video drivers
4.3.4 Clocks
4.3.5 DCS
4.3.6 ADC
4.3.7 D-I/O
4.4 Filters, Apertures and Shutter
4.4.1 Filter Wheel
4.4.2 Aperture Wheel
4.4.3 Shutter
The principal design goals of the APT are to achieve a dimensionally stable, photometrically precise optical telescope and detector. These requirements have lead to the design of a monolithic telescope/detector assembly with few optical surfaces. This design harmonizes nicely with a natural desire to achieve a simple, light-weight experiment package. The APT telescope concept, while dependent on some new technology, has its key components derived from previous NASA space experiments. The telescope and detector concepts are derived from the Gravity Probe B optical telescope, and MDI and Trace detector concepts. The telescope is housed in an outer aluminum shell which protects the telescope from contamination. The Structure also holds the telescope rotation and pointing mechanisms. Covers in the Al structure will be opened in flight, to expose the telescope optics. Multiple redundant filters on one or more filter wheels are used to select wavelength regions of scientific interest.
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Table 1: Instrument Characteristics
An instrument cover protects the APT during launch and orbit insertion. Once on orbit, a single command will deploy the instrument cover. This is a one off operation, and the cover will then remain open for the duration of the mission.
The pointing and guiding strategy is separated into two sections. A `slow' mechanism provides Sun pointing for the instrument. The purpose of this mechanism is to point the APT precisely enough for the fine guiding mechanism to acquire the Sun. The slow mechanism will compensate for post-launch instrument depointing, and on-orbit spacecraft pointing errors. The error signal for the slow pointing system will be derived from the same sensor as for the fine guider.
The fine guiding mechanism is an active secondary mirror with a pointing range of several arcseconds. The fine guider mechanisms have flight heritage in the Solar Optical Universal Polarimeter (Title, 1984), MDI (Scherrer et al., 1995) and Luminosity Oscillations Imager instruments (LOI) (Appourchaux et al., 1997). The secondary mirror is mounted on 3 piezo-electric stacks. Varying the voltage to the stacks changes their length, and will provide image stabilization and focus control in flight. A piezo-electric actuator with built-in redundancy was developed for the LOI instrument, and a similar technique could be applied to the APT stacks.
A custom InGaAs detector will provide the guide error signal. The detector will have redundant limb sensors that provide the gude signal for a closed loop servo system with the acive secondary. A central detector element will measure the IR flux at 1.6µm. Four separate diodes at the edge of the substrate are covered and used to measure the detector temperature and temperature gradients in flight. A scheme similar to this is used by the VIRGO/LOI flying on SoHO.
| Actuators | PZT |
| Operating Voltage | 0 - 100 V |
| Pointing range | 22" |
| Displacement range | 10 µm |
| Displacement sensitivity | 10 -3 " |
| Stability | drift < 10 -3 " / min. |
| Servo bandwidth | 20 Hz (TBD) |
The Aperture wheel selects a single off-axis aperture at one of four angular positions. This provides direct information about the instrument focus.
With prpoer focus the solar image will not shift as the off-center pupil masks are individually selected. Changes in the system focal length would cause the image radius to change. Any such focus or detector shift is distinguishable from a scale change in the solar image produced from opposite pupil masks will be shifted by a few millipixels
The design of the filter wheels will be based on flight proven hardware from MDI and other optical instruments. Provision for atleast 6 filters will be made in the wheel. This will provide two filters at each of three wavelengths for redundancy and to check filter degradation. Filters will be cycled into the observing beam with a cadence of approximately one change per hour. A closed loop servo system of redundant drive motors and optical encoders will select the filter to be used and allow constant knowledge of which filter is in the optical path.
A rotating shutter at the entrance aperture provides the timimg of the image exposures. A closed loop servo system similar to that used for the filter wheels will drive the shutter.
In order to calibrate optical aberrations and their changes the telescopr-detector assembly must be rotated.
A drive mechanism similar to that of the Ultraviolet Coronograph Spectrometer (UVCS) instrument of SoHO (Kohl et al., 1995) is forseen, though the APT version can be significantly simpler as the moving mass is much less. A flexible cable will bring out signals from the rotating telescope as this is simpler, more reliable and less electrically noisy than slip rings.
| Number of pixels | 2048 x 2048 |
| Pixel size | 14 µm square |
| CCD full well | > 150 ke- |
| Digitization | 16 bits |
| Readout noise | < 30 e- |
| Exposure time | 100 ms |
| Exposure control | Aperture plane shutter |
| Summing | 20 images |
The camera design will be based on previous flight systems, or systems currently under development. Table 3 provides a baseline for the camera development based on the Thomson THX 7899 detector used by the Precision Solar Photometric Telescope (PSPT). Preamplifiers and clock drivers will be mounted close to the camera head on the rotating telescope assembly. However, the bulk of the camera electronics and associated image aquisition and processing hardware will be housed in a separate electronics bay that does not rotate with the detector.
Radiation damage to the CCD detector through the mission will be minimized by detector design, operation mode and shielding. Trade off of these parameters will ultimately depend on the mission orbit.