This paper provides an overview of the satellite based Sapphire Payload developed by COM DEV to be used for
observing Resident Space Objects (RSOs) from low earth orbit by the Canadian Department of National Defence. The
data from this operational mission will be provided to the US Space Surveillance Network as an international
contribution to assist with RSO precision positional determination.
The payload consists of two modules; an all reflective visible-band telescope housed with a low noise preamplifier/focal
plane, and an electronics module that contains primary and redundant electronics. The telescope forms a low distortion
image on two CCDs adjacent to each other in the focal plane, creating a primary image and a redundant image that are
offset spatially. This combination of high-efficiency low-noise CCDs with well-proven high-throughput optics provides
a very sensitive system with low risk and cost. Stray light is well controlled to allow for observations of very faint
objects within the vicinity of the bright Earth limb. Thermally induced aberrations are minimized through the use of an
all aluminum construction and the strategic use of thermal coatings.
The payload will acquire a series of images for each target and perform onboard image pre-processing to minimize the
downlink requirements. Internal calibration sources will be used periodically to check for health of the payload and to
identify, and possibly correct, any pixels with an aberrant response. This paper also provides a summary of the testing
that was performed and the results achieved.
A detector consisting of two biased microchannel plates and a P20 or P43 phosphor screen was illuminated with 500-eV electrons in order to characterize the size and amplitude of individual microchannel plate (MCP) firings as a function of phosphor-to-MCP distance d and voltage Vacc. The P43 phosphor was significantly brighter (71%) than P20 at Vacc = 4250 V, and brighter but less so (4%) at Vacc = 8000 V. Events were Gaussian-shaped with full widths at half maximum of order 0.27 mm for Vacc = 4250 V and d = 2 mm. Widths decreased by 23% to 28% when Vacc was increased to 8000 V, and increased between 9 to 34% when doubling d, depending on Vacc and phosphor type.
Satellite-based remote sensing of atmospheric CO2 holds the promise to greatly improve our understanding of the processes which regulate atmospheric CO2 and the global carbon cycle. However, the required precision and resolution of such measurements needed to characterise sources and sinks of CO2 on regional scales presents strong instrument design challenges. One type of remote sensing instrument which has been proposed to measure the integrated-column concentration of CO2 is a Gas-Filter Correlation Radiometer (GFCR). As a technique, a GFCR is a radiometer which uses a sample of the gas of interest as a spectral filter for that gas in the atmosphere. In this paper we present a "strawman" design for a GFCR satellite instrument to remotely sense atmospheric CO2. This design, which includes multi-pass CO2 and O2 gas cells with path lengths of up to 10 metres, demonstrates that such an instrument can be built within the constraints of a satellite environment.
KEYWORDS: Tunable filters, James Webb Space Telescope, Sensors, Electronics, Mirrors, Stars, Space operations, Interfaces, Control systems, Optical filters
The science instrumentation for the James Webb Space Telescope (JWST) has concluded its Phase A definition stage. We have developed a concept for the JWST Fine Guidance Sensor (FGS), which will form the Canadian contribution to the mission. As part of the JWST re-plan in early 2003, the FGS design was recast to incorporate a narrow-band (R~100) science-imaging mode. This capability was previously resident in the NIRCam instrument. This FGS science mode makes use of tunable filters and filter wheels containing blocking filters, calibration sources and aperture masks. The science function of the FGS Tunable Filters (FGS-TF) remains complementary to the NIRCam science goals. Narrow-band FGS-TF imaging will be employed during many of the JWST deep imaging surveys to take advantage of the sensitivity to emission line objects. The FGS-TF will also provide a coronagraphic capability for the characterization of host galaxies of active galactic nuclei and for the characterization of extra solar planets. The primary function of the FGS remains to provide the sensor data for the JWST Observatory line-of-sight stabilization system. We report here on the overall configuration of the FGS and we indicate how the concept meets the performance and interface requirements.
The MOPITT instrument operates on the principle of correlation spectroscopy where the incoming signal is modulated by gas filter and chopper mechanisms and synchronously demodulated within the signal processing system. The performance and flexibility required by the MOPITT instrument resulted in the development of a novel timing control and signal processing design. This design synchronizes modulation and demodulation from a central programmable timing control unit. The data collection system performs a highly linear sigma-delta analog-to-digital conversion prior to signal demodulation. The demodulation operation includes data averaging which reduces the sampled signal bandwidth and extends the signal to noise ratio of the data to in excess of the analog-to-digital converter's rated 16-bit dynamic range.
The Measurements Of Pollution In The Troposphere (MOPITT) instrument is one of five instruments flying on NASA's Terra (formerly EOS-AM1) spacecraft that was launched in December 1999. This paper describes the MOPITT instrument mechanical configuration and how it was derived based on system considerations and spacecraft interfaces. These system level considerations include contamination control, EMC/EMI, thermal, optical and structural behavior. The key spacecraft interfaces include mechanical mounting, optical field-of-view (FOV), and thermal transfer. In addition, a detailed discussion is provided for the cryogenic region of the instrument that contains detectors, cold optics, warm optics, and active coolers. Special test fixtures were designed and incorporated in this region of the instrument to permit cooling of the detectors during ambient atmospheric conditions. Some of these test fixtures were designed to fly due to the difficulty in removing them. This utility of operating the instrument's cryogenic detectors within the laboratory environment was extremely beneficial during the instrument optical alignment, EMC testing, and special optical system tests. Final configuring (or closeout) of the instrument's cryogenic region for flight was performed to balance contamination and EMC risks. On-orbit data about the effectiveness of this closeout is provided.
Space flight optical instruments and their support hardware must reliably operate in stressing environments for the duration of their mission. They must also survive the mechanical and thermal stresses of transportation, storage and launch. It is necessary to qualify the hardware design through environmental testing and to verify the hardware's ability to perform properly during and/or after some selected environmental tests on the ground. As a rule, flight electronics are subjected to thermal, mechanical and electromagnetic environmental testing. Thermal testing takes the form of temperature cycling over a temperature difference range (Delta) T of up to 100 degrees C for a minimum of six cycles, with additional performance verification testing at the hot and cold extremes. Mechanical testing takes the form of exposure to random vibration, sine sweep vibration, shock spectra and static loading on a centrifuge or by sine burst on a vibration table. A standard series of electromagnetic interference and electromagnetic compatibility testing is also performed.
The fine error sensor (FES) system is a star tracker like device that is used within the Lyman far ultraviolet spectroscopic explorer (FUSE) instrument. The FES extends the performance achievable from the conventional star trackers by providing very low noise guidance reference signals. In addition the FES is used for scene and target identification for the UV system operation. This is accomplished by the FES looking through the Lyman FUSE telescope system by means of a pick-off mirror at an intermediate image. This paper describes the proposed readout electronics for the FES CCD and the CCD clock driving system. The clock driving system is unique in that software can reconfigure the hardware required to generate the various CCD clocks and the timing signals for the readout electronics. An adaptable method used to define the location of the six sub-rasters, used to identify the selected guide stars, is described. A variable clocking rate is utilized to minimize readout time while minimizing read noise. The benefits of using RAM based field programmable gate array (FPGA) technology for flight programs also is highlighted.
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