NASA’s Lucy mission spacecraft was launched on 16 October 2021 and will perform the initial in situ investigation of the Jovian Trojan asteroids (Levison et al. 2021, 2024). The Lucy LOng Range Reconnaissance Imager (L’LORRI) is a panchromatic visible light (420–795 nm, 50 V ≈ 20.4 at spatial resolutions far surpassing that available from Earth. This paper describes the L’LORRI instrument design and the requirements that drove the design. We present results from L’LORRI’s ground calibration campaign, summarize the L’LORRI in-flight calibration plan, and describe typical L’LORRI operations scenarios during the Trojan flybys. We also present an analysis of in-flight data taken during the first year of Lucy operations, which show that most aspects of L’LORRI’s performance are nominal (i.e., as predicted), but the telescope’s point spread function is slightly degraded relative to pre-flight predictions. Nevertheless, L’LORRI is still expected to fulfill all of its scientific objectives, which should revolutionize our view of the Jovian Trojans.

Although originally developed for astronomical applications, the space qualification and availability of the Teledyne HAWAII detector make it appealing for high-precision Earth-observing systems such as the carbon monoxide correlation radiometer required for GEO-CAPE. In this shot noise-limited application, the signal-to-noise ratio of a co-averaged measurement is driven by the detector's temporal stability. To assess the stability, we operated the H2RG under monitored blackbody illumination. The Teledyne SIDECAR ASIC provided 16-bit digitization and clocking for integration times faster than the frame conversion time. With proper application of reference signals, the co-averaging of hundreds of frames is possible. Integrations of one-quarter of the full well depth can attain precision to the 200 ppm level in the co-averaged result. For integrations above three-quarters the well depth, the precision reaches 111 ppm.

The MErcury Surface, Space ENvironment, GEochemistry, and Ranging (MESSENGER) spacecraft, launched in August 2004 and planned for insertion into orbit around Mercury in 2011, has already completed two flybys of the innermost planet. The Mercury Dual Imaging System (MDIS) acquired nearly 2500 images from the first two flybys and viewed portions of Mercury's surface not viewed by Mariner 10 in 1974-1975. Mercury's proximity to the Sun and its slow rotation present challenges to the thermal design for a camera on an orbital mission around Mercury. In addition, strict limitations on spacecraft pointing and the highly elliptical orbit create challenges in attaining coverage at desired geometries and relatively uniform spatial resolution. The instrument designed to meet these challenges consists of dual imagers, a monochrome narrow-angle camera (NAC) with a 1.5° field of view (FOV) and a multispectral wide-angle camera (WAC) with a 10.5° FOV, co-aligned on a pivoting platform. The focal-plane electronics of each camera are identical and use a 1024×1024 charge-coupled device detector. The cameras are passively cooled but use diode heat pipes and phase-change-material thermal reservoirs to maintain the thermal configuration during the hot portions of the orbit. Here we present an overview of the instrument design and how the design meets its technical challenges. We also review results from the first two flybys, discuss the quality of MDIS data from the initial periods of data acquisition and how that compares with requirements, and summarize how in-flight tests are being used to improve the quality of the instrument calibration.

The LOng-Range Reconnaissance Imager (LORRI) is the high-resolution imaging instrument for the New Horizons mission to Pluto, its giant satellite Charon, its small moons Nix and Hydra, and the Kuiper Belt, which is the vast region of icy bodies extending roughly from Neptune’s orbit out to 50 astronomical units (AU). New Horizons launched on January 19, 2006, as the inaugural mission in NASA’s New Frontiers program. LORRI is a narrow-angle (field of view=0.29°), high-resolution (4.95 μrad pixels), Ritchey-Chrétien telescope with a 20.8-cm diameter primary mirror, a focal length of 263 cm, and a three-lens, field-flattening assembly. A 1,024×1,024 pixel (optically active region), thinned, backside-illuminated charge-coupled device (CCD) detector is used in the focal plane unit and is operated in frame-transfer mode. LORRI provides panchromatic imaging over a bandpass that extends approximately from 350 nm to 850 nm. LORRI operates in an extreme thermal environment, situated inside the warm spacecraft with a large, open aperture viewing cold space. LORRI has a silicon carbide optical system, designed to maintain focus over the operating temperature range without a focus adjustment mechanism. Moreover, the spacecraft is thruster-stabilized without reaction wheels, placing stringent limits on the available exposure time and the optical throughput needed to satisfy the measurement requirements.

The Mercury Dual Imaging System (MDIS) on the MESSENGER spacecraft will provide critical measurements tracing Mercury’s origin and evolution. MDIS consists of a monochrome narrow-angle camera (NAC) and a multispectral wide-angle camera (WAC). The NAC is a 1.5° field-of-view (FOV) off-axis reflector, coaligned with the WAC, a four-element refractor with a 10.5° FOV and 12-color filter wheel. The focal plane electronics of each camera are identical and use a 1,024×1,024 Atmel (Thomson) TH7888A charge-coupled device detector. Only one camera operates at a time, allowing them to share a common set of control electronics. The NAC and the WAC are mounted on a pivoting platform that provides a 90° field-of-regard, extending 40° sunward and 50° anti-sunward from the spacecraft + Z -axis—the boresight direction of most of MESSENGER’s instruments. Onboard data compression provides capabilities for pixel binning, remapping of 12-bit data into 8 bits, and lossless or lossy compression. MDIS will acquire four main data sets at Mercury during three flybys and the two-Mercury-solar-day nominal mission: a monochrome global image mosaic at near-zero emission angles and moderate incidence angles, a stereo-complement map at off-nadir geometry and near-identical lighting, multicolor images at low incidence angles, and targeted high-resolution images of key surface features. These data will be used to construct a global image base map, a digital terrain model, global maps of color properties, and mosaics of high-resolution image strips. Analysis of these data will provide information on Mercury’s impact history, tectonic processes, the composition and emplacement history of volcanic materials, and the thickness distribution and compositional variations of crustal materials. This paper summarizes MDIS’s science objectives and technical design, including the common payload design of the MDIS data processing units, as well as detailed results from ground and early flight calibrations and plans for Mercury image products to be generated from MDIS data.

The Compact Reconnaissance Imaging Spectrometer for Mars (CRISM) is a hyperspectral imager on the Mars Reconnaissance Orbiter (MRO) spacecraft. CRISM consists of three subassemblies, a gimbaled Optical Sensor Unit (OSU), a Data Processing Unit (DPU), and the Gimbal Motor Electronics (GME). CRISM's objectives are (1) to map the entire surface using a subset of bands to characterize crustal mineralogy, (2) to map the mineralogy of key areas at high spectral and spatial resolution, and (3) to measure spatial and seasonal variations in the atmosphere. These objectives are addressed using three major types of observations. In multispectral mapping mode, with the OSU pointed at planet nadir, data are collected at a subset of 72 wavelengths covering key mineralogic absorptions and binned to pixel footprints of 100 or 200 m/pixel. Nearly the entire planet can be mapped in this fashion. In targeted mode the OSU is scanned to remove most along‐track motion, and a region of interest is mapped at full spatial and spectral resolution (15–19 m/pixel, 362–3920 nm at 6.55 nm/channel). Ten additional abbreviated, spatially binned images are taken before and after the main image, providing an emission phase function (EPF) of the site for atmospheric study and correction of surface spectra for atmospheric effects. In atmospheric mode, only the EPF is acquired. Global grids of the resulting lower data volume observations are taken repeatedly throughout the Martian year to measure seasonal variations in atmospheric properties. Raw, calibrated, and map‐projected data are delivered to the community with a spectral library to aid in interpretation.

Domingue, F. Morgan, R. Greeley, C. Paranicas, L. Prockter, D. Roth, T. Roush, K. Strohbehn, P. Thompson, M. Wirzburger, The Johns Hopkins University Applied Physics Laboratory, 11100 Johns Hopkins Rd. Laurel MD 20723 (scott.murchie@jhuapl.edu) , Department of Geological Sciences, Arizona State University, Box 87104, Tempe AZ 85287-1404 (greeley@asu.edu), NASA Ames Research Center, MS 245-3, Moffett Field, CA 94035-1000 (troush@mail.arc.nasa.gov).

The LOng-Range Reconnaissance Imager (LORRI) is an instrument that was designed, fabricated, and qualified for the New Horizons mission to the outermost planet Pluto, its giant satellite Charon, and the Kuiper Belt, which is the vast belt of icy bodies extending roughly from Neptune's orbit out to 50 astronomical units (AU). New Horizons is being prepared for launch in January 2006 as the inaugural mission in NASA's New Frontiers program. This paper provides an overview of the efforts to produce LORRI. LORRI is a narrow angle (field of view=0.29°), high resolution (instantaneous field of view = 4.94 μrad), Ritchey-Chretien telescope with a 20.8 cm diameter primary mirror, a focal length of 263 cm, and a three lens field-flattening assembly. A 1024 x 1024 pixel (optically active region), back-thinned, backside-illuminated charge-coupled device (CCD) detector (model CCD 47-20 from E2V Technologies) is located at the telescope focal plane and is operated in standard frame-transfer mode. LORRI does not have any color filters; it provides panchromatic imaging over a wide bandpass that extends approximately from 350 nm to 850 nm. A unique aspect of LORRI is the extreme thermal environment, as the instrument is situated inside a near room temperature spacecraft, while pointing primarily at cold space. This environment forced the use of a silicon carbide optical system, which is designed to maintain focus over the operating temperature range without a focus adjustment mechanism. Another challenging aspect of the design is that the spacecraft will be thruster stabilized (no reaction wheels), which places stringent limits on the available exposure time and the optical throughput needed to accomplish the high-resolution observations required. LORRI was designed and fabricated by a combined effort of The Johns Hopkins University Applied Physics Laboratory (APL) and SSG Precision Optronics Incorporated (SSG).

The LOng-Range Reconnaissance Imager (LORRI) is a panchromatic imager for the New Horizons Pluto/Kuiper belt mission. New Horizons is being prepared for launch in January 2006 as the inaugural mission in NASA's New Frontiers program. This paper discusses the calibration and characterization of LORRI. LORRI consists of a Ritchey-Chretien telescope and CCD detector. It provides a narrow field of view (0.29°), high resolution (pixel FOV = 5 μrad) image at f/12.6 with a 20.8~cm diameter primary mirror. The image is acquired with a 1024 x 1024 pixel CCD detector (model CCD 47-20 from E2V). LORRI was calibrated in vacuum at three temperatures covering the extremes of its operating range (-100°C to +40°C for various parts of the system) and its predicted nominal temperature in-flight. A high pressure xenon arc lamp, selected for its solar-like spectrum, provided the light source for the calibration. The lamp was fiber-optically coupled into the vacuum chamber and monitored by a calibrated photodiode. Neutral density and bandpass filters controlled source intensity and provided measurements of the wavelength dependence of LORRI's performance. This paper will describe the calibration facility and design, as well as summarize the results on point spread function, flat field, radiometric response, detector noise, and focus stability over the operating temperature range. LORRI was designed and fabricated by a combined effort of The Johns Hopkins University Applied Physics Laboratory (APL) and SSG Precision Optronics. Calibration was conducted at the Diffraction Grating Evaluation Facility at NASA/Goddard Space Flight Center with additional characterization measurements at APL.

The Compact Reconnaissance Imaging Spectrometer for Mars (CRISM) will launch in 2005 on the Mars Reconnaissance Orbiter (MRO) mission, with its primary science objective to characterize sites with aqueous mineral deposits hyperspectrally at high spatial resolution. CRISM's two Offner relay spectrometers share a single entrance slit with a dichroic beamsplitter. The IR focal plane contains a 640 (spatial) x 480 (spectral) HgCdTe FPA with a 980 nm to 3960 nm spectral bandpass. It is cooled to 110 K to minimize dark current, and coupled to a 28 min long cold shield to minimize thermal background. The spectrometer housing is cooled to -90C for the same reason. A three-zone IR filter consisting of two broadband filters and a linear variable filter overlays the IR focal plane, eliminating multiple grating orders and providing additional. attenuation of the thermal background. The visible focal plane contains a 640 (spatial) x 480 (spectral) silicon photodiode array, with a 380-1050 nm spectral bandpass occupying approximately 106 rows of the detector. A two-zone filter comprised of two different Schott glasses eliminates multiple grating orders. The two focal planes together cover 544 spectral channels with a dispersion of 6.55 nm/channel in the VNIR and 6.63 nm/channel in the IR. The optics and focal planes are gimbaled, and a pre-programmed slew can be used to remove groundtrack motion while superimposing a scan across a target. CRISM will operate in two basic modes: a scanning, high resolution mode to hyperspectrally map small, targeted areas of high scientific interest, and a fixed, nadir-pointed, lower resolution pixel-binned mode using selected wavelength channels to obtain near-global coverage to find targets. Preliminary performance of the CRISM instrument is presented, and is compared with prior system design predictions.

CRISM (Compact Reconnaissance Imaging Spectrometer for Mars) is a hyperspectral imager that will be launched on the MRO (Mars Reconnaissance Orbiter) spacecraft in August 2005. MRO’s objectives are to recover climate science originally to have been conducted on the Mars Climate Orbiter (MCO), to identify and characterize sites of possible aqueous activity to which future landed missions may be sent, and to characterize the composition, geology, and stratigraphy of Martian surface deposits. MRO will operate from a sun-synchronous, near-circular (255x320 km altitude), near-polar orbit with a mean local solar time of 3 PM. CRISM’s spectral range spans the ultraviolet (UV) to the mid-wave infrared (MWIR), 383 nm to 3960 nm. The instrument utilizes a Ritchey-Chretien telescope with a 2.12° field-of-view (FOV) to focus light on the entrance slit of a dual spectrometer. Within the spectrometer, light is split by a dichroic into VNIR (visible-near-infrared, 383-1071 nm) and IR (infrared, 988-3960 nm) beams. Each beam is directed into a separate modified Offner spectrometer that focuses a spectrally dispersed image of the slit onto a two dimensional focal plane (FP). The IR FP is a 640 x 480 HgCdTe area array; the VNIR FP is a 640 x 480 silicon photodiode area array. The spectral image is contiguously sampled with a 6.6 nm spectral spacing and an instantaneous field of view of 61.5 μradians. The Optical Sensor Unit (OSU) can be gimbaled to take out along-track smear, allowing long integration times that afford high signal-to-noise ratio (SNR) at high spectral and spatial resolution. The scan motor and encoder are controlled by a separately housed Gimbal Motor Electronics (GME) unit. A Data Processing Unit (DPU) provides power, command and control, and data editing and compression. CRISM acquires three major types of observations of the Martian surface and atmosphere. In Multispectral Mapping Mode, with the gimbal pointed at planet nadir, data are collected at frame rates of 15 or 30 Hz. A commandable subset of wavelengths is saved by the DPU and binned 5:1 or 10:1 cross-track. The combination of frame rates and binning yields pixel footprints of 100 or 200 m. In this mode, nearly the entire planet can be mapped at wavelengths of key mineralogic absorption bands to select regions of interest. In Targeted Mode, the gimbal is scanned over ±60° from nadir to remove most along-track motion, and a region of interest is mapped at full spatial and spectral resolution. Ten additional abbreviated, pixel-binned observations are taken before and after the main hyperspectral image at longer atmospheric path lengths, providing an emission phase function (EPF) of the site for atmospheric study and correction of surface spectra for atmospheric effects. In Atmospheric Mode, the central observation is eliminated and only the EPF is acquired. Global grids of the resulting lower data volume observation are taken repeatedly throughout the Martian year to measure seasonal variations in atmospheric properties.

A filtered imager, the CONTOUR Forward Imager (CFI), was designed, fabricated, and qualified for the Comet Nucleus Tour (CONTOUR) Discovery class mission. The CONTOUR spacecraft was launched July 3, 2002, and failed during injection to heliocentric orbit on August 15, 2002. This paper provides an overview of the efforts to produce CFI. The CFI imager was designed to perform optical navigation, comet nucleus imaging, and comet coma imaging. CFI was complemented in the CONTOUR payload by the CONTOUR Remote Imager and Spectrometer (CRISP). The emphasis in the CFI design was on high sensitivity at moderate to long ranges from the comet nucleus, while CRISP was designed for high-speed observations in close to the nucleus. A unique aspect of CFI was the requirement to image multiple comets after being exposed to high-velocity cometary dust on the previous comet flybys (which damages and contaminates the forward looking optics). The first optical surface was replaceable between comet encounters, using a mirror "cube" mechanism, to alleviate the dust damage. Another challenging aspect of the design is that the spacecraft was thruster stabilized (no reaction wheels), placing limits on the available exposure time to accomplish the high sensitivity observations required. CFI utilized ten filters covering from 300 to 920 nm to image onto a backthinned 1024 by 1024 element CCD. The Ritchie-Chrietien telescope provided a clear aperture of 62 mm, a full field of view of 2.5 degrees, and a pixel field of view of 43 microradians. CFI was designed and fabricated by a combined effort of the Johns Hopkins University Applied Physics Laboratory and SSG Precision Optronics. The CONTOUR mission was lost prior to CFI being powered on in flight.

The CONTOUR Remote Imager and Spectrometer (CRISP) was a multi-function optical instrument developed for the Comet Nucleus Tour Spacecraft (CONTOUR). CONTOUR was a NASA Discovery class mission launched on July 3, 2002. This paper describes the design, fabrication, and testing of CRISP. Unfortunately, the CONTOUR spacecraft was destroyed on August 15, 2002 during the firing of the solid rocket motor that injected it into heliocentric orbit. CRISP was designed to return high quality science data from the solid nucleus at the heart of a comet. To do this during close range (order 100 km) and high speed (order 30 km/sec) flybys, it had an autonomous nucleus acquisition and tracking system which included a one axis tracking mirror mechanism and the ability to control the rotation of the spacecraft through a closed loop interface to the guidance and control system. The track loop was closed using the same images obtained for scientific investigations. A filter imaging system was designed to obtain multispectral and broadband images at resolutions as good as 4 meters per pixel. A near IR imaging spectrometer (or hyperspectral imager) was designed to obtain spectral signatures out to 2.5 micrometers with resolution of better than 100 meters spatially. Because of the high flyby speeds, CRISP was designed as a highly automated instrument with close coupling to the spacecraft, and was intended to obtain its best data in a very short period around closest approach. CRISP was accompanied in the CONTOUR science payload by CFI, the CONTOUR Forward Imager. CH was optimized for highly sensitive observations at greater ranges. The two instruments provided highly complementary optical capabilities, while providing some degree of functional redundancy.

We have developed a general purpose data compression algorithm using Chebyshev polynomials which preserves the needed information for quantitative scientific analysis. The algorithm is computationally simple yet provides compression factors much greater than two. This method is suitable for compression of science data from many types of space instruments, including spectral data, particle or photon count rate data, and magnetometer data. Presently available data compression schemes are not well suited for such scientific data streams and are not generally used. We present the status of our development effort of this algorithm and its potential implementation on future space qualified flight processors.

Sealed channel plate photomultipliers with multiple discrete anodes for use as photon counting detectors in the image plane of Fabry-Perot interferometers are described. The influence of design and construction on performance of completed devices is discussed. Effects on spatial resolution, lifetime, and counting efficiency are described. It is shown that devices can be optimized for particular applications. The results should be generally applicable to resistive anode and wedge and strip anode types of sealed detectors.

This paper addresses the imager designs for an upcoming NASA mission to examine comets at distances from the sun fairly similar to that of the earth. The general constraints on design typical to a small spacecraft such as low mass and operation in vacuum apply. In addition the transient nature of the data which are presented as spacecraft moves by the comet at high speed must be considered. Two imagers with different tasks to perform will be flown on this mission. Design choices for electronics and selection of charge coupled device image sensors are discussed. Some early results from breadboard tests are presented.

This paper addresses the imager designs for an upcoming NASA mission to examine comets at distances from the sun fairly similar to that of the earth. The general constraints oil design typical to a small spacecraft such as low mass and operation in vacuum apply. In addition the transient nature of the data which are presented as the spacecraft moves by the comet at high speed must be considered. Two imagers with different tasks to perform will be flown on this mission. Design choices for electronics and selection of charge coupled device image sensors are discussed. Some early results from breadboard tests are presented.