THE NEAR INFRARED MAPPING SPECTROMETER EXPERIMENT ON GALILEO 



                                by 



        R.W. Carlson, P.R. Weissman, W.D. Smythe, J.C. Mahoney, 

                               and 
  
              the NIMS Science and Engineering Teams* 
       





* The Near Infrared Mapping Spectrometer (NIMS) Engineering and Science Teams
consist of I. Aptaker (Instrument Manager), G. Bailey (Detectors), K. Baines
(Science Coordinator), R. Burns (Digital Electronics), R. Carlson (Principal
Investigator), E. Carpenter (Structures), K. Curry (Radiative Cooler), 
G. Danielson (Co-Investigator), T. Encrenaz (Co-Investigator), H. Enmark
(Instrument Engineer), F. Fanale (Co-Investigator), M. Gram (Mechanisms), 
M. Hernandez (NIMS Orbiter Engineering Team), R. Hickok (Support Equipment
Software), G. Jenkins (Support Equipment), T. Johnson (Co-Investigator), 
S. Jones (Optical-Mechanical Assembly), H. Kieffer (Co-Investigator), 
C. LaBaw (Spacecraft Calibration Targets), R. Lockhart (Instrument Manager),
S. Macenka (Optics), J. Mahoney (Instrument Engineer), J. Marino (Instrument
Engineer), H.Masursky (Co-Investigator), D. Matson (Co-Investigator),
T. McCord (Co-Investigator), K. Mehaffey (Analog Electronics), A. Ocampo
(Science Coordinator), G. Root (Instrument Systems Analysis), R. Salazar
(Radiative Cooler and Thermal Design), D. Sevilla (Cover Mechanisms),
W. Sleigh (Instrument Engineer), W. Smythe (Co-Investigator and Science 
Coordinator), L. Soderblom (Co-Investigator), L. Steimle (Optics), 
R. Steinkraus (Digital Electronics), F. Taylor (Co-Investigator),
P. Weissman (Co-Investigator and Science Coordinator), and D. Wilson
(Manufacturing Engineer) 



This paper is dedicated to the memories of Gary Bailey and Hal Masursky. Gary
would have been very pleased with the excellent performance of his detectors
and Hal would have enjoyed the Galileo flyby of Venus, one of his favorite
planets.  Their crucial contributions to NIMS and Galileo will continue to be
apparent throughout the mission, and will be appreciated in whatever future
success we may enjoy. 





                         Table of Contents 



1. Introduction 

2. Scientific Objectives
    2.1 Venus
    2.2 Earth
    2.3 Moon
    2.4 Asteroids
    2.5 Jupiter Atmosphere
    2.6 Jovian Satellites 

3. Instrument Description 
    3.1 General Description and Operation
    3.2 Optical Design
    3.3 Detectors and the Focal Plane Assembly
    3.4 Radiative Cooler
    3.5 Mechanisms
    3.6 Electronics Design
    3.7 Thermal Design
    3.8 Instrument Contamination Protection 

4. Instrument Calibration
    4.1 Introduction
    4.2 Spectral Calibration
    4.3 Radiometric Calibration
    4.4 Spatial Calibration
    4.5 Spacecraft Radiometric Calibration Target
    4.6 Spacecraft Photometric Calibration Target

5. Operating Modes and Data Acquisition
    5.1 Instrument Modes and Operation  
    5.2 Instrument Commands
    5.3 Typical Encounter Operations 

6. Spacecraft Interactions
    6.1 Thermal Control
    6.2 Contamination Control 
    6.3 Instrument Pointing
    6.4 Spacecraft Obscuration 

7. NIMS Mission Design Aspects 




 
                       LIST OF ILLUSTRATIONS

          [Figures are in separate Postscript files named
           INSTFGnn.PS where nn is the figure number.]


Fig. 1            Photograph of the NIMS Instrument. 

Fig. 2            Schematic Diagram of the Instrument and Scanning             
                  Motions. 

Fig. 3            Schematic of the NIMS Optical Design. 

Fig. 4            Photographs of the Focal Plane Assembly.

Fig. 5            Focal Plane Assembly Components. 

Fig. 6            Radiative Cooler Assembly

Fig. 7            Electronics Block Diagram 

Fig. 8            Preamplifier and Autobias Circuitry

Fig. 9            Analog Signal Chain Electronics

Fig. 10           Instrument Timing 

Fig. 11           Spectral Bandpass Response 

Fig. 12           Spatial Response 

Fig. 13           Parameter Table Assignments

Fig. 14           Spacecraft Obscuration



                           LIST OF TABLES


Table  1         Instrument Parameters

Table  2         NIMS Standard Operating Modes





                              Abstract 



The Galileo Near Infrared Mapping Spectrometer (NIMS) is a combination of
imaging and spectroscopic methods.  Simultaneous use of these two methods
yields a powerful combination, far greater than when used individually.  For
geological studies of surfaces, it can be used to map morphological features,
while simultaneously determining their composition and mineralogy, providing
data to investigate the evolution of surface geology.  For atmospheres, many of
the most interesting phenomena are transitory, with unpredictable locations. 
With concurrent mapping and spectroscopy, such features can be found and
spectroscopically analyzed.  In addition, the spatial/compositional aspects of
known features can be fully investigated.  The NIMS experiment will investigate
Jupiter and the Galilean satellites during the two year orbital operation
period, commencing December 1995.  Prior to that, Galileo will have flown past
Venus, the Earth/Moon system (twice), and two asteroids; obtaining scientific
measurements for all of these objects. 

The NIMS instrument covers the spectral range 0.7 to 5.2 microns, which
includes the reflected-sunlight and thermal-radiation regimes for many solar
system objects.  This spectral region contains diagnostic spectral signatures,
arising from molecular vibrational transitions (and some electronic
transitions) of both solid and gaseous species.  Imaging is performed by a
combination of one-dimensional instrument spatial scanning, coupled with
orthogonal spacecraft scan-platform motion, yielding two-dimensional images for
each of the NIMS wavelengths. 

The instrument consists of a telescope, with one dimension of spatial scanning,
and a diffraction grating spectrometer. Both are passively cooled to low
temperatures in order to reduce background photon shot noise. The detectors
consist of an array of indium antiminide and silicon photovoltaic diodes,
contained within a focal-plane-assembly, and cooled to cryogenic temperatures
using a radiative cooler.  Spectral and spatial scanning is accomplished by
electro-mechanical devices, with motions executed using commandable instrument
modes. 

Particular attention was given to the thermal and contamination aspects of the
Galileo spacecraft, both of which could profoundly affect NIMS performance.
Various protective measures have been implemented, including shades to protect
against thruster firings as well as thermal radiation from the spacecraft. 


 


                            1. Introduction 

The Near Infrared Mapping Spectrometer (NIMS) experiment on the Galileo Orbiter
Spacecraft represents a combination of imaging and spectroscopic methods.  The
advantage of spectral imaging, as opposed to pure imaging systems or
single-field-of-view spectrometers, is the ability to simultaneously find,
identify, and map compositional units on planetary surfaces.  With such
information one can investigate the geochemical evolution of satellite surfaces
as well as the dynamical and compositional properties of atmospheres. 

The NIMS instrument possesses modest spatial and spectral resolution, and
operates in the near-infrared range of 0.7 - 5.2 microns.  This spectral range
is particularly diagnostic of minerals known or suspected to occur on planetary
and satellite surfaces, and also includes many observable features of
atmospheric species.  The spectral resolution was designed for investigating
the relatively broad bands seen in surface reflectance, yet is adequate for
identifying several major and minor atmospheric constituents. 
      
Galileo will be inserted into Jupiter orbit in December of 1995, commencing a
nearly two year investigation of the Jovian system, performing eleven orbits
around the planet during this period.  Prior to Jupiter arrival, Galileo has
flown by Venus (February 1990) and will fly by the Earth and Moon twice
(December of 1990, 1992) and two asteroids--Gaspra (October 1991) and Ida
(August 1993). NIMS will investigate all of these bodies.  The scientific
objectives for these measurements are briefly presented in the following
Section (2). 

The purpose of this paper is to describe and document the NIMS instrument
design and development, which posed many unique design and spacecraft
integration challenges, owing largely to the low temperatures required by the
detectors and optics. The instrument itself is described in Section 3, followed
by calibration aspects (4), instrument operating modes (5), and spacecraft
considerations (6). Mission design aspects which relate to NIMS are presented
in Section 7.  General aspects of the NIMS experiment have been briefly
documented by Carlson (1981), and a description of the Galileo mission and its
complement of experiments is contained in a volume by Yeates et al. (1985). 

Although this paper is devoted to a discussion of the instrument - i.e. the
hardware -  the corresponding analysis tools - the software - must also be
mentioned.  NIMS generates a great amount of such data in a short amount of
time, and this large volume of data can only be digested using
highly-efficient, visual, interactive computer methods.  These computer tools
are therefore an integral part of the NIMS experiment, and a separate
discussion is warranted.  These aspects are not discussed here.  In the
interim, refer to Torson (1989) for an overview of NIMS data visualization
capabilities. 





                     2. Scientific Objectives 

2.1 Venus 

During the Venus flyby (February, 1990), NIMS measurements concentrated on
spectral features which arise from surface and deep-atmosphere thermal
emission.  These features occur in spectral regions where CO2 is relatively
transparent, allowing one to probe far below the cloud region and to measure
the spatial variations of the intervening cloud extinction.  This infrared
radiation is observable on the nightside of Venus and was only recently
discovered (Allen and Crawford, 1984). Two types of measurements were obtained
from NIMS: (1) multiple spatial images at selected infrared wavelengths for
dynamical studies, and (2) detailed spectra at a variety of latitudes and
longitudes for chemical abundance information, specifically H2O and CO in the
deep atmosphere. The results from this flyby are described by Carlson et al
(1991).

2.2 Earth 

During the first flyby of the Earth-Moon system, NIMS performed both
atmospheric and surface spectral-imaging of the Earth and similar geological
investigations of the Moon.  For the Earth measurements, NIMS investigated
mesospheric airglow emission features and obtained geological maps of Australia
and Antarctica.  Exploratory global mapping of the Earth was also performed;
the first time ever in this spectral region.  Comprehensive lunar measurements
were also obtained by NIMS at multiple phase angles, but with relatively poor
spatial resolution. 

One of the primary goals of NIMS during both the 1990 and 1992 flybys is
investigation of mesospheric water, observable through limb scans of infrared
fluorescence in the 2.7 micron band.  It has been recently proposed (Thomas, et
al., 1989) that noctilucent clouds, and upper-atmosphere water in general, has
increased over the past century due to a larger amount of biological and
anthropogenic emission of methane, a photochemical source of water in the
mesosphere. The abundance of water in the upper atmosphere is poorly
determined. Microwave measurements are in general agreement with photochemical
models, but rocket ion spectroscopy and infrared measurements indicate a
greater water abundance, perhaps implying an additional source of mesospheric
water (c.f. Garcia, 1989). NIMS may provide an independent measurement of the
vertical water profile at several latitudes, extending to the summertime south
polar region. 

Other limb airglow emission features that were investigated are the O2 infrared
bands, the well-known infrared hydroxyl bands, the strong CO2 v3 band, ozone
emissions at 4.8 microns, and the tail of the NO (1-0) band, a
prime cooler of the thermosphere. The NIMS limb measurements occur within
roughly +/- 1/2 hour of closest approach, and sample both the night- and
day-side airglow. 

As Galileo receded from the Earth, NIMS investigated specific geographic
regions as determined by the timing of the flyby, illumination geometry, and
the gain characteristics of the instrument. These gain characteristics were
established for our ultimate goal, Jupiter at 5.2 AU from the sun, and tend to
be too sensitive for general Earth observations.  The Earth at 1 AU is simply
too bright. Nevertheless, by choosing favorable geometry, we were able to
perform spectral mapping of both Australia and Antarctica during the 1990 pass
and plan to investigate other regions in 1992.  Using ground-truth
measurements, we can corroborate and extend our remote sensing measurements to
both continental and planetary scales. 

2.3 Moon 

On the first pass through the Earth-Moon system, the closest approach of
Galileo was ~350,000 km, limiting NIMS resolution to ~170 km, in contrast to 10
km typical of spot spectrometer measurements obtained from ground-based
observations. This initial pass provided viewing of roughly a quarter of the
lunar surface (selenographic longitudes 90 to 180 deg)  not previously observed
spectroscopically.  The phase angle range of observations (30 to 150 deg)
extended the range of viewing geometry available from Earth, particularly for
sub-Earth meridians, where Earth observations are limited to 90 deg phase
angles.  The wavelength coverage was also extended by NIMS, terrestrial
observations being generally limited to below 2.5 microns by the atmosphere.
However, because the NIMS dynamic range is designed for measurements at
Jupiter, the lunar observations were saturated at many wavelengths,
particularly for small incidence angles; unsaturated spectra to 2.5 microns
generally require incidence angles greater than 80 deg. At wavelengths beyond
about 3 microns, thermal emission becomes important. Because of this
dynamic-range limitation, full NIMS spectra are available for only narrow
angular regions. Nonetheless, with multiple observations taken at many
geometries, some important lunar questions can be addressed, including: (1)  an
initial search for hydrated minerals - there is a remote possibility that
hydrated minerals may be present near the polar regions where low temperatures
may occur.  The high sensitivity of NIMS allows a search for the 3 micron
hydration feature during the second Earth-Moon encounter which passes over the
Moon's north polar region, with a spatial resolution of about 60 km. (2)
Spectral characterization of additional lunar areas - with the two passes,
about 20% of the Moon not visible from the Earth can be mapped with surface
resolution from 200 to 500 km. (3) Obtaining the first spectra in the 2.5 to
5.2 micron region.  The NIMS wavelength range will allow an accurate
determination of local surface temperatures  and a correction for thermal
emission for that portion of the spectrum that contains reflected radiation. 

2.4 Asteroids 

Galileo will fly by the main belt asteroids Gaspra (16 km diameter) and Ida (32
km diameter) in October 1991 and August 1993, respectively.  Ground based
studies have identified both asteroids as S-type.  The NIMS goal at each
encounter will be to acquire spectral-reflectance and thermal-emission images
of the asteroid at maximum possible spatial resolution.  Because of the high
velocity of the flybys, it will be possible to obtain resolved images of only
one hemisphere of each object.  Full disc, unresolved spectra of the other
hemisphere will be obtained during approach.  NIMS will identify and map
mineral species on the surface of each asteroid, and will seek to determine if
the surfaces are chemically heterogeneous. 

2.5 Jupiter Atmosphere 

Although NIMS was originally conceived for satellite surface spectral
reflectance measurements, the experiment is well-suited for a variety of
investigations of the Jupiter atmosphere.  This is due to: (1) The spectral
range available to NIMS, which includes signatures from several minor species
such as germane, phosphine, and water, which are produced in the deep
atmosphere of Jupiter and may serve as tracers of motions in this unexplored
altitude region of the planet.  (2) The NIMS spectral range also includes
absorption by the more abundant molecules: CH4, NH3, and H2. Due to variations
in absorption strength for these molecules and cloud layer variability, NIMS
can probe a large altitude range, ranging from the high-altitude polar
aerosols, far above the ammonia cloud deck, extending down into the water cloud
region at the several-bar-level.  (3) The spatial resolution of the experiment,
about 300 km, is sufficient to resolve the many dynamical features of the
atmosphere and to investigate their temporal changes.  Investigation of changes
over time scales of hours to years are possible with the repeated observations
available from Galileo.  Finally (4) a full range of phase angle coverage is
available with Galileo, enabling one to optically investigate the microphysical
properties of the diverse cloud layers.  Prior discussions of the NIMS
measurements at Jupiter can be found in Taylor and Calcutt (1984) and Hunten et
al. (1986). 

2.6 Jovian Satellites 

For the three large icy Galilean satellites, Callisto, Ganymede, and Europa,
the primary NIMS science objectives are to map the various surface
compositional units and to identify their elemental and mineralogical
composition.  An important aspect is to study these three objects as a
collection.  Voyager multispectral data suggests that there are common
compositional units across the three objects.  For example, the crust of
Callisto, saturated with scars of ancient impact structures, has 
the same albedo and color as the remnants of the oldest terrain 
on Ganymede.  Likewise, the much younger grooved terrains on Ganymede
appear to be similar to the linear markings on Europa.

A primary question concerning these units is the composition of the dark
components that are mixed with the dominant water-ice crusts.  Are these
materials silicates or organic-rich materials derived from primitive objects
such as comet nuclei?  If silicates, NIMS may detect bands due to olivines,
pyroxenes, or a range of iron-bearing minerals.  C-H features might be present
if the satellite surfaces contain dark organic components such as those found
on some asteroids.  Magnetospheric sources of implanted material (e.g. sulfur)
may be an important process, providing  chemically reactive species  which can
modify the surfaces (e.g. generating S-O from the S implanted in H2O). 

Another set of questions pertain to the formation and history of these
compositional units.  Are there systematic correlations of dark component
abundance or nature with geologic age and setting?  Were the dark materials
added early during the accretion or late during subsequent impacts?  Methods to
address these questions involve studying the global distribution of these
units, examining leading- versus trailing-hemispheres as well as the geologic
setting of such units.  For example, dark-rayed craters on Ganymede may expose
units of concentrated dark materials that exist as layers and lenses in the
subsurface.  Ejecta may show compositional correlations with size, age,
latitude, terrain type, or longitude that will provide insights into their
nature and origin.  Another important question concerns other volatile species.
These might be involved in active processes of eruption and transport, in
particular on Europa, or be found as cold-trapped species, for instance in
Ganymede's icy polar caps. 

The innermost Galilean satellite, Io, is of great interest to the NIMS
investigation, exhibiting not only a wide range of volcanic processes, some of
which are continuously active, but also an equally bizarre mixture of surface
chemistry, composition and mineralogy.  Active volcanic processes include: 1)
two classes of eruptive plumes thought to be driven by superheated volatiles in
the form of sulfur and sulfur dioxide vapor, 2) eruption of icy clouds of
material along  scarps and fractures that have been proposed to be due to the
escape of liquid sulfur dioxide to the surface, where it explosively forms
gases and ice, and 3) dark hot spots, ranging in temperature up to 400 kelvins,
which have been proposed to be crusted pools of liquid sulfur. All of the
processes likely involve a wide range of chemical components other than S and
SO2.  Other volatiles may be cold-trapped in the polar regions.  H2S could
exist in regions on the surface; it has been identified in ground-based
spectra.  Not only is there a plethora of potential spectral reflection
features from the exotic compositional components, but gaseous absorption
features may be detected over hot volcanic regions and the thermal output of
the volcanic activity itself will be monitored and mapped.  There is even the
possibility that molten silicates will occasionally breach through to the
surface. 





                       3. Instrument Description 

3.1 General Description and Operation 

Performing simultaneous spectral and spatial infrared mapping at Jupiter
necessitates use of high light-gathering power optics and sensitive detectors,
both operating at low temperatures to minimize detector and background noise.
The resulting NIMS instrument, shown in a photograph in Fig. 1 and schematically
in Fig. 2, consists of the following major elements: (1) a telescope with one
dimension of spatial scanning, (2) an optical chopper to provide a dark
reference, (3) a diffraction grating spectrometer which disperses the radiation
onto the focal plane assembly, (4) a focal plane assembly consisting of
multiple detectors, optical filters, and preamplifier circuitry, (5) a passive
radiative cooler which cools the focal plane to cryogenic temperatures, and (6)
signal processing and control electronics, not shown. 

The typical operating mode of the instrument can be described with the aid of
Fig. 2.  One dimension of spatial scanning is provided within the telescope
employing a "wobbling" secondary mirror, giving 20 contiguous pixels, each with
0.5 x 0.5 mrad resolution, for a total angular field of 0.5 x 10 mrad.  The
other dimension of spatial scanning is provided by slowly slewing the
spacecraft scan platform in the cone direction. During one half of the up/down
mirror scan, the grating is set at a fixed angle, with a corresponding set of
17 wavelengths striking the 17 individual detectors of the focal plane
assembly.  At the extremes of the mirror scan, the grating is stepped to a new
setting, and a new set of wavelengths are measured during the second half of
the mirror scan. The chopper frequency is synchronized to the 63 Hz spacecraft
timing, and the mirror and grating motions are synchronized to the chopper,
with motions taking place during the dark portion of the chopper cycle.  The
size and number of grating steps can be adjusted for specific encounter
conditions and scientific objectives, and the scan platform motion is matched
to the resulting spectral scan time.  Details of the instrument timing are
given in Section 3.6 and the instrument modes are described in Section 5. 
Additional information on many aspects of the instrument are contained in a
series of papers by Aptaker (1982a, 1982b, 1983, 1987).  A tabulation of
instrument parameters is given below: 

 


                                TABLE 1 
                   
     Instrument Parameters for the Near Infrared Mapping Spectrometer 

============================================================================

Angular Resolution:        0.5 mrad x 0.5 mrad (individual pixel size).

Angular Field:             10  mrad x 0.5 mrad (20 pixels, cross-cone direction.

Spectral Range:            0.7 - 5.2 u.

Spectral Resolution:       0.0250 u (lambda > 1 u),  0.0125 (lambda < 1 u ).

No. of Spectral Samples:   Variable; 408, 204, 102, ... ,1.
                                 
Spectral Scan Time:        Variable; 8 2/3 sec (408 spectral samples),
                                     4 1/3 sec (204 spectral samples),
                                     ........
                                     1/3 sec   (17 spectral samples).

Telescope:                 22.8 cm Diameter f/3.5 Ritchey-Chretien, 
                           800 mm equivalent focal length, 
                           spatial scanning via moving secondary mirror,
                           Operating Temperature ~ 150 kelvins.

Telescope Etendue:         1.1 x 10-4 cm2 steradian.

Spectrometer:              39 lines/mm plane grating spectrometer, 
                           400 mm focal length, f/3.5, Dall-Kirkham collimator,
                           200 mm focal length, f/1.8, flat-field camera,
                           Bipartite diffraction grating, 30% blazed for 1.9 u,
                                                          70% blazed for 3.8 u.
                           Operating Temperature ~ 150 kelvins.

Detectors:                 Seventeen individual photovoltaic diodes,
                           15 indium antimonide, 2 silicon,
                           Active area  = 0.2 mm x 0.2 mm,
                           Quantum efficiencies > 70 %.

Radiative Cooler:          Passive radiative cooler, 
                           Achieves 64 kelvins, with detectors energized.

Noise Equiv. Radiance:     7 x 10-9 W cm-2 sterad-1 per pixel and
                           per spectral resolution element (0.025u),
                           at 3 u and 70 kelvins FPA temperature.
                            
Noise Equivalent Albedo:   0.0002 at 5 AU.

Mechanisms:                Torque motor drives for spatial and spectral scans,
                           Tuning fork chopper,
                           Deployable covers for telescope, radiative cooler.

Electronics:               Seventeen channel signal chain,
                           Microprocessor controlled (RCA 1802).

Gain States:               4 ground-commandable gain states, detectors 1-14,
                           Automatic gain switching, detectors 15-17.
 
Protective Devices:        Covers for pre- and post-launch protection,
                           Continuous instrument purging through launch,
                           Heaters for continuing contamination control.

Mass:                      18 kg.

Power:                     12 Watts (average), 13 Watts (peak).

Dimensions:                83 x 37 x 39 cm (optics),
                           20 x 25 x 13 cm (electronics).

Data Rate:                 11.52 kbps

Data Encoding:             10 bits (0 - 1023)

Mounting:                  Scan platform,
                           Co-aligned with SSI, UVS, PPR

On-board Calibration:      Photometric Calibration Target (PCT), 
                           (a solar reflectance target), and a
                           Radiometric Calibration Target (RCT-NIMS), 
                           a blackbody radiator.




3.2 Optical Design 
       
A variety of considerations led to the ultimate design of the NIMS optics,
which primarily involved signal and instrument noise aspects. In order to
minimize the latter, one must use small-area detectors, as the noise varies as
the square root of detector area.  At the same time, maximizing the signal on a
detector requires large acceptance angles, i.e. use of a low f-number camera
system illuminating the detectors.  Similarly, adequate signal is obtained
through use of a large diameter telescope, consistent with the system etendue'
and angular resolution requirements.  The resulting optical design of the
instrument is illustrated in Fig. 3 and consists of all-reflective telescope
and spectrometer sections.  Much of the optical design and fabrication were
performed by the Perkin-Elmer Corporation.  Development and testing of the NIMS
optics have been previously described by Macenka (1983). 

The telescope is a 228 mm diameter, f/3.5 Ritchey-Chretien  design, with an
equivalent focal length of 800 mm.  The secondary mirror steps in 20 equal
increments, sweeping the image in the plane of the field stop in 0.5 mrad
increments and thereby providing one dimension of spatial scanning.  The field
stop is a 400 micron wide slit, which defines a 0.5 mrad field-of-view, normal
to the mirror scan direction and parallel to the plane of dispersion of the
spectrometer.  The angular resolution in the other direction is defined by the
projection of the detectors at the field stop, and is approximately 0.5 mrad.
Spatial response measurements are given in Section 4.3.  The slit was made
longer than required in order to ensure against misalignments caused by thermal
or vibration induced shifts.  For a distant point source, the telescope forms
an image with 90% of the energy contained within a circle of angular diameter
0.05 mrad, which is quite satisfactory when compared to the aforementioned 0.5
mrad resolution.  An InGaAs light emitting diode is mounted on the telescope
spider and is used for inflight wavelength verification of the spectrometer. 

The spectrometer employs a plane diffraction grating, illuminated by a
Dahl-Kirkham collimator and followed by a wide-angle flat-field camera which
focuses the entrance slit (the telescope field stop) onto the detectors.
Stepwise rotation of the grating allows the complete NIMS spectrum to be 
generated. 

The collimator has an effective focal length of 400 mm and a focal ratio of
f/3.5. The smaller mirror in the collimator has a slightly toroidal surface to
correct some system aberrations, primarily astigmatism from grating
anamorphism. 

The grating is a 39 lines/mm dual-blazed grating, with 30% of the area blazed
for 1.9 microns and the remainder for 3.8 microns.  The first order of the
grating is used for wavelengths greater than 1 micron (the InSb detectors), and
the second order for shorter wavelengths (Si detectors).  Between the blaze
configuration and use of multiple orders, reasonable efficiencies can be
obtained over the relatively large wavelength range of NIMS.  Measurements of
the blaze efficiency of the flight grating are presented by Macenka (1983).
Ruling of the master NIMS grating, replication, and measurement of efficiencies
were performed by the Perkin-Elmer Corporation. 

The detectors are widely spaced in the focal plane, requiring a wide-angle,
flat-field camera.  The two mirrors comprising the camera are both rather
extreme aspheric surfaces, with an effective focal length of 200 mm and a focal
ratio of f/1.75. 

The linear dispersion in the focal plane is 8 mm/micron in first order.  For
the active area of the detectors (0.20 mm square) the spectral width of each is
0.025 micron.  This matches the width of the entrance slit, yielding a
triangular spectral bandpass, only slightly broadened due to finite spot sizes
(see Section 4.2).  The grating can be stepped through minimum increments of
one-half of a spectral resolution element.  The spectrometer exhibits some
residual astigmatism and was aligned for the best spectral focus, consequently
the vertical spatial resolution profile is somewhat broadened (see Section
4.3). 

All of the mirrors and the grating were fabricated from fused-silica.  Because
of the large number of reflections, efficient infrared-reflective surfaces are
required; NIMS uses pure gold with no protective overcoats, obtaining
reflectivities of R = 96 to 98% in the spectral interval 0.7 to 5.2 microns.
Obscuration, mainly by the camera secondary mirror, reduces the incident energy
with a transmission factor of Tobs = 60%, giving an etendue' for the optics,
exclusive of reflection losses and grating efficiencies, of A  Tobs = 6.1x10-05
cm2 steradian.  Stray light is reduced with baffles and an interior finish of
matte black (Bostik-Finch Catalac Black). 

The optics and their housing must operate at low temperatures in order to
minimize photon shot noise from background thermal emission.  Furthermore,
initial alignment was performed at room temperature, thus thermal aspects are
an important part of the design, and are discussed in Section 3.7. 

3.3 Detectors and the Focal Plane Assembly 
    
There are seventeen individual detectors (15 InSb and 2 Si) contained within
the focal plane assembly (FPA), along with their associated spectral filters
and electronic preamplifier components.  Shielding from high energy particles,
necessary in the Jovian magnetosphere, is provided by a hermetically sealed
Tantalum case,  which includes a sapphire window for optical input.  A platinum
resistance thermometer provides a measure of the FPA temperature.  Fig. 4 shows
photographs of the FPA while Fig. 5 illustrates the various components and
their packaging.  The FPA was manufactured by Cincinatti Electronics
Corporation, and various aspects are described by Bailey (1979) and Smith et
al. (1982). 

Dispersed radiation from the spectrometer enters the FPA through two sapphire
windows, the first contained within the radiative cooler (see Section 3.4),
while the second is integral with the FPA and is anti-reflection coated for the
near infrared.  The radiation then passes through optical filters, whose
purpose is to reject higher-order radiation and to limit the amount of thermal
radiation incident on the detectors.  These filters are at the cryogenic
temperatures of the FPA and emit negligible amounts of thermal radiation that
can be sensed by the detectors.  A cold field-of-view limiting aperture is
placed just behind the filters, further limiting thermal radiation on the
detectors. 

Each of the seventeen photodiode detectors contained in the FPA has a photo-
active area of 0.2 x 0.2 mm , and each is anti-reflection coated for its own
individual spectral region.  Detector quantum efficiencies of 70% or greater
were measured for the coated photodiodes.  By operating the detectors at a
constant bias voltage, near zero, the detectors act as high impedance current
sources, with linear response to incident photon flux. 

Noise performance of the instrument is determined by the detectors and their
preamplifiers, each of which can be characterized by a noise-current spectral
density.  When the detectors are operated at near-zero bias, only Johnson noise
contributes, with a noise-current spectral density of 

               i = sqrt(4kT*Deltaf/RD), 

where RD is the junction resistance (Hall et al., 1975).  RD exhibits a rapidly
varying temperature dependence (Kruse et al., 1962), varying as  

                 RD oc T^(-1/2)*exp(1/2 eV/kT), 

V being the detector band gap (0.22 eV for InSb).  From this, the InSb detector
noise is expected to show roughly a factor of three change for a temperature
change of 10 kelvins, and this dependence was indeed observed for the NIMS
instrument.  Measurements of individual NIMS InSb photodetectors at 77 kelvins
show a noise current density of 1.8 X 10^-15 Amps/sqrt(Hz), with 0.9 x 10-15
Amps/sqrt(Hz) for the preamplifiers.  This implies that NIMS will be
preamplifier-noise limited for FPA temperatures less than about 70 kelvins.
NIMS noise levels measured in flight, as well as in the laboratory, are
entirely consistent with the above values. 

A hybrid dual junction field effect transistor (JFET) preamplifier front-end
and 1010 feedback resistors are mounted in close proximity to each detector. 
Locating these components within the FPA minimizes circuit noise contributions
by providing minimum capacity at the detector and by providing a low impedance
interface to the external electronics assembly.  The preamplifier circuitry
contains a differential source-follower circuit which was experimentally found
to give the best electromagnetic interference (EMI) performance. A commandable
automatic bias circuit is incorporated in order to maintain the detectors at
the aforementioned constant near-zero bias. This feature allows higher
temperature operation of the InSb detectors by enforcing nonsaturation of the
preamplifiers due to increased dark current and was included for the unlikely
event of high FPA temperatures. Experience in flight has shown that there is no
need for this precaution, although possible inflight contamination of the
radiative cooler could change this situation. Additional details of the
preamplifier electronics are given in Section 3.6 

Radiation shielding is provided by a 3 mm thick tantalum enclosure and by
shielding the back of the camera secondary mirror, providing complete angular
shielding. The shields were designed to limit the integrated exposure to < 10
krad during the nominal Galileo mission, to prevent radiation damage, while
simultaneously reducing noise caused by penetrating magnetospheric particles,
mainly electrons.  Tests were performed on the developmental model using
energetic gamma rays; from this one can predict the resulting performance at
Jupiter:  Jovian radiation noise will be most severe for the observations taken
during the Io close encounter just prior to Jupiter orbit insertion, where a
predicted signal-to-noise ratio of 10:1 is predicted. Corresponding results for
flyby observations of Europa and Ganymede are 30:1 and 100:1 respectively. 
Measurements from Callisto's distance are unaffected by ambient magnetospheric
radiation. 

High energy magnetospheric protons can also cause displacement damage, with
loss of sensitivity for InSb detectors, but such loss can be recovered by
annealing the detectors at roughly 300 kelvins.  For this reason, the radiative
cooler contains a commandable heater which can elevate the detectors to
annealing temperatures.

3.4 Radiative Cooler 

The indium antimonide detectors require cryogenic temperatures for operation
and this is achieved with a single-stage passive radiative cooler, illustrated
in Fig. 6.  The cold stage, containing the FPA, has a 627 cm2 aluminum
honeycomb plate which radiates energy to space and cools the detectors down to
64 kelvins.  The ultimate temperature is determined by input power to the cold
stage and arises from many sources.  The FPA dissipates 9 mW of electrical
power, and thermal conduction from the cable and support mechanism provides
another thermal path. Incident radiation is another source, and can arise from
the instrument and the cooler housing and shield, from the spacecraft and other
instruments, and from the planet itself.  Minimization of these sources is
discussed in the following paragraphs and in Section 6.1.  The cooler was
provided by the Santa Barbara Research Center. 

Mechanical support of the cold stage is accomplished through a suspension
system of fiberglas bands, similar to the mounting of a bicycle wheel hub.
This arrangement not only provides an extremely low thermal conductance path,
but also very stable mechanical positioning stability ( 0.01 mm). Electrical
access to the FPA is provided by a 60-conductor ribbon cable with narrow
stainless steel plated conductors in a thin polyimide sandwich, with a Ni
coating for shielding. 

Radiative loading from the instrument is minimized using a surrounding outer
shroud with very low-emittance gold surfaces.  The outer surfaces of the cold
stage are similarly treated giving very low radiative coupling.  Radiation
from the spacecraft is reduced using a shield around the radiator plate to
optically block emissions from the scan platform and portions of the
spacecraft, the latter dependent upon the scan platform cone angle. Radiative
loading to the honeycomb plate by the inner surface of the shield itself is
reduced by use of a low emissivity (<0.03), highly specular (99 % at 7 microns)
gold coating.  Additionally, the shield is thermally insulated from the cooler
body, and radiatively cools to ~ 120 kelvins. The surface of the shield which
faces the scan platform is thermally isolated with multiple layers of thermal
blanketing, while the outward-facing radiating surface is painted for high
emissivity. 

In order to maintain the cooler peformance, it is crucial that the low
emissivity of the inner shield surface be maintained.  Any contamination would
seriously increase the radiative loading; consequently the shield contains a 26
Watt strip heater to elevate the temperature and avoid condensation of
spacecraft outgassing and thruster products.  This heater is used nearly
continuously, except during observation periods, and elevates the shield to
about 300 kelvins.  Additional contamination protection, for the pre-launch and
the early post-launch period, was inclusion of a deployable radiator cover and
continuous dry nitrogen purging (see Sections 3.8 and 6.2).  A heater is also
included for the cold stage, for use in annealing the detectors in the unlikely
event of radiation damage (see Section 3.3).

The radiator is mounted at 62.5 deg to the telescope optical axis, in the plane
of rotation of the scan platform.  Its full field-of -view is circular, with an
angular radius of 71.5 deg.  The total mass is 1.9 kg.  The thermal mass of the
cold stage is about 250 J deg^-1 (at 90 kelvins,temperature dependent), giving
a cool-down time from 300 to 65 kelvins of 30 hours.  The initial cooling rate
is very rapid, but slows at lower temperatures, requiring 24 hours to cool from
125 kelvins to 65 kelvins.  Additional in-flight performance values are given
in Section 6.1. 

3.5 Mechanisms 

There are three mechanisms in the NIMS instrument: an optical chopper and two
nearly identical, mirror and grating drives.  In addition, the telescope and
cooler had protective covers, since deployed in flight. 

The chopper serves to modulate the detected radiation, allowing the dark
current level of the detector to be subtracted on a pixel-by-pixel basis.  It
is a tuning fork with a resonant frequency slightly above the spacecraft 63 Hz
timing signal, and contains two moving blades, each mounted on opposing tines,
which chop the light.  The chopper is located immediately in front of the
telescope field stop (the spectrometer entrance slit) and modulates the
radiation with an approximate 50 % duty cycle.  The back surfaces of the blades
are coated with a black finish, while the front is coated with reflective gold.
When the instrument is off, the blades are in the open position, and behind a
pre-slit. This protects the low-thermal-mass blades from damage by inadvertent
sun pointing. 

The chopper is driven at the spacecraft 63 Hz timing rate.  This rate is
slightly below the natural resonant frequency; in the event of a drive circuit
failure the fork can be operated at its natural frequency, with overall
synchronization with spacecraft timing being maintained by NIMS instrument
software. Motion sensing and drive power are provided by magnetic pickup coils
and a phase comparator feedback loop.  The amplitude is kept at a constant
level through an amplitude comparison circuit.  The tuning fork and electronic
design were provided by the American Time Company. 

Mirror and grating motion is accomplished with direct-current magnetic motors,
similar to a loudspeaker drive, but adapted for angular rather than linear
motion.  The rotor, which undergoes very small, stepwise, angular excursions,
is a fork which pivots on low-loss flexures and contains a samarium-cobalt
magnet.  An enclosing drive solenoid is attached to the stator.  Motion is
sensed through a pair of linear voltage differential transformers (LVDTs), one
for position/velocity feedback control, the second as an independent,
telemetered measure of the grating and mirror positions.  Each LVDT consists of
a permeable core, mounted on the rotor, and three stator-mounted coils - two
for the drive current and one for position sensing.  Motion of the core within
the coils varies their coupling, and the induced EMF is linearly proportional
to displacement. 

In operation, the desired position (mirror or grating) is generated in the
microprocessor and converted to an analog voltage.  Error signals generate
motion drive, which is further controlled by rate information. The mechanism
thus steps from one position to the next, settling to a stable position within
a short interval (less than the dark half of a chopper cycle for the mirror,
somewhat more for the grating). 
 
Two ejectable covers are used to protect the NIMS optics and radiator (see
Section 3.8). Each covers consists of a lightweight aluminum frame with a
multi-layer thermal blanket closure.  They are mounted onto the telescope and
cooler apertures using uncaptured hinges, and locked into place with removable
latch pins.  Conical dowels ensure against launch-induced translation.  Both
the 25 cm optics cover and the 41 cm cooler cover are ejected at the same time
by a common release device.  The cover eject command is performed by an
electric signal that fires redundant pyrotechnic squibs.  These "bellows
actuators" remain hermetically sealed after firing to prevent any contamination
to the instrument, and through a lever-and- piston mechanism cause a pair of
steel cables to be pulled.  Each cable simultaneously unlatches a cover by
pulling the constraining latch pin. Cover ejection is accomplished using
torsion springs mounted at the hinge, causing a cover to rotate open when
unlocked, and to slide free under its own momentum once it has opened
approximately 100 deg. 

3.6 Electronics Design 
 
The electronics assembly is mounted on the spacecraft scan platform near the
optical assembly and contains the following circuits:  analog, digital, scan
mechanism and chopper drivers, and power supplies.  A block diagram of the
electronics showing the interaction of the various sub-modules and their
external connections to the Command and Data Subsystem (CDS) and the Power/Pyro
Subsystem (PPS) is presented in Fig. 7. 
 
The analog subassembly consists of 17 signal processing circuits (one for each
detector), a multiplexer, analog-to-digital converter and miscellaneous
circuitry such as engineering telemetry and a calibration lamp driver. 
 
To maintain detector bias voltage and dark current stability with temperature
and time, each InSb detector amplifier (channels 3-17) incorporates a bias
correction servo loop which samples the amplifier output signal during the
dark-signal portion of the chopper cycle and holds the amplifier output, and
thus the detector bias voltage, at a predetermined level. 

There are two types of signal processing circuits employed in the NIMS (see
Fig. 9).  For most of the NIMS wavelength interval, the signals are determined
by surface or atmospheric albedo and their range can be accurately predicted.
Thus one can use ground commands to accomplish the infrequently required gain
changes.  On the other hand, there are transient, localized hot spots in the
Jovian atmosphere which are due to unpredictable cloud clearings.  These
features allow one to probe the deep atmosphere, showing variable and often
intense thermal emission in the 5 micron region, and require a much larger
dynamic range than those required for the lower wavelength channels. 

Commandable gain state switching is used in channels 1 through 14 (albedo
channels).  These commandable gain states are achieved by switching resistors
at the input to the gated integrators.  The instrument may be commanded to one
of four possible gain states which then selects the appropriate resistors for
each of the 14 albedo channels.  The gain of each channel has been set for that
detector's predicted response at Jupiter.  In particular, the nominal gain state
(gain state 2) corresponds at full scale (1023 data numbers, DN) to an albedo
of ~ 1.2 at Jupiter's solar distance.  Gain states 3 and 4 are each more
sensitive by a factors of two and four, respectively.  Gain state 1 is similar
to gain state 2, except channels 10-14 are each reduced in order to obtain
measurements from the spacecraft Radiometric Calibration Target (RCT-NIMS, see
Section 4.5). 

Channels 15 through 17 (the thermal channels) use a dual-gain amplifier which
automatically switches gain with input level and achieves a dynamic range of
10,000.  The two possible gains (low signal level = high gain, high signal
level = low gain) for each of the thermal channels has also been optimized for
Jovian system measurements, with brightness temperatures of 230 to 325 kelvins
measurable in the low gain configuration. 

A commandable electronic calibration signal can be introduced to verify the
gain (in gain state 2) for each of the seventeen signal chains. 

The digital subassembly contains the central controller and mode sequencing
functions necessary to operate the mechanisms in response to external commands,
to acquire and format science and engineering data, and to output engineering
telemetry and science data via the spacecraft data bus.  Science data is
transferred at a rate of 11.52 kilobits/sec and engineering telemetry is
transferred at a rate of 36 bits/sec.  Commands to the mechanisms are initiated
at the closing of the chopper blades.  Since the chopper is phase-locked to the
instrument timing chain, which in turn is phase-locked to the spacecraft
real-time interrupt (RTI), all mechanism and other instrument operations are
accurately synchronized to the spacecraft clock. 
 
The digital subassembly is based on the RCA 1802 microprocessor.  Its operating
software is interrupt driven and permits a flexible selection of instrument
operating sequences.  Mode defining parameter tables are used to control the
operation of each sequence.  Ground commands permit modification of the
parameter tables which allows the instrument operation to be uniquely tailored
to a specific science opportunity (see Section 5.3). 
 
The chopper driver circuitry operates on commands from the digital subassembly
and phase-synchronizes the chopper blades to the instrument and spacecraft
clocks.  The analog signal chain demodulator signal can be derived directly
from the chopper (Chopper Reference mode) or from the 63 Hz drive signal (63 Hz
mode).  There are slight phase differences between the two; the former is the
preferred mode, and is slightly more sensitive.  The relative timing of the
mechanisms is shown in Fig. 10.  The circuitry also regulates the physical
amplitude of the chopper vane movement.  The chopper's resonant frequency is
slightly greater than the 63 Hz drive frequency. This allows a third chopper
mode, wherein the chopper oscillates at its own natural frequency, with
synchronism maintained by software control, occasionally slipping chopper
cycles to maintain lock. 

Both the scan mirror and the grating are actuated by torque motors.  Each
torque motor is driven by a closed-loop servo circuit which uses the output of
a linear voltage differential transducer (LVDT) as the position sensor.  The
LVDT output is continuously compared against the position requested by the
microprocessor.  An error signal is generated which is then applied to the
torque motor for position correction.  To minimize instantaneous power demand,
the software sequences commands to the mirror and grating drivers such that the
two will not be in motion simultaneously. 
 
The power supply consists of current-limited power converters and regulators
necessary to supply all voltages needed by the various subassemblies.  It also
generates a power-on reset signal which controls the initialization of the
instrument.  At power-on, the instrument is placed in a quiescent mode, with
neither the mirror nor grating scanning; the signal chain is set to the nominal
gain state 2 and the instrument is synchronized to the 63 Hz clock instead of
the chopper. 

3.7 Thermal Design 

The low temperatures required for the FPA and optics posed many unique problems
in both instrument design and thermal loads from external sources; the latter
is discussed in Section 6.1.  There are two general categories of design
problems, the first is in achieving the required low temperatures, the second
is ensuring that alignment shifts and mechanical stresses are reduced to
tolerable levels. 

The optical assembly is passively cooled by radiation, attaining ~ 150 kelvins
during both laboratory measurements as well as inflight.  This is accomplished
by having an unblanketed instrument and using narrow stainless steel mounting
struts with low thermal conduction.  The three struts form a kinematic mount,
consisting of a monopod, bipod, and tripod.  A prestress was applied during
mounting to allow for thermal contraction of the instrument relative to the
warm scan platform.  Thermal conduction by the electrical cabling was reduced
using long, minimum diameter wiring.  Radiative coupling to the scan platform
and nearby instruments is minimized by their thermal blanketing. The optical
assembly was painted with a low solar absorptance, high emissivity white paint
(zinc orthotitanate). This material is a moderate electrical conductor,
providing protection against electrostatic discharge. 

Optical alignment was carried out at room temperature and it is important that
it be maintained at the much lower operating temperature.  Most critical are
the primary-secondary mirror spacings, which are athermalized using invar rods.
Thermal gradient distortion is avoided by directly mounting the optics, except
for the grating, to a central optical "bench" of high-thermal-conductivity
material (aluminum).  The mirror mount design is crucial to the optical
quality, since differential thermal contraction can distort, or even break, the
optical elements.  NIMS uses an inherently athermal mounting scheme.  The back
surface of each mirror, in the central mounting region, is ground flat, while
the front has a conical shape, with the apex of the cone in the plane defined
by the back (c.f. Fig. 3).  Flat and conical aluminum retainers were machined
with the same geometry. Since there is no change in shape with temperature,
clearances and clamping forces remain constant. Only a relative sliding motion
results, facilitated by 0.001 inch thick mylar sandwiched between the surfaces.

3.8 Instrument Contamination Protection 

There are a multitude of contamination sources which can degrade, or destroy,
NIMS instrument performance.  Ground-based sources include water-vapor
absorption and deposition of particulate matter on the optics and radiator.
These sources are especially pronounced during the launch phase, where
vibration frees many particles, and outgassing of organics and other molecules
also occurs. During the interplanetary injection IUS burn, the entire
spacecraft is subject to impingement by motor products.  During cruise, the
spacecraft is a source of outgassing water vapor and organics, which can
condense on the cold NIMS instrument.  Thruster byproducts from trajectory and
attitude correction maneuvers are continuing sources of contamination.  For all
of these reasons, NIMS has adopted a dual approach for contamination
protection:  first, to incorporate within the instrument as many protective
measures as possible (discussed below), and second, to minimize external
contamination sources insofar as possible (see Section 6.2). 

The first NIMS protective measure was to use only low-outgassing materials and
to further subject them to high temperature bakeout, prior to assembly.  Upon
final assembly, the instrument was kept in a dry environment, either in a
dessicated container, or it was purged with dry nitrogen.  Purge protection was
nearly continuously maintained, even during Shuttle operations and launch,
finally terminating two minutes prior to release from the Shuttle bay. 

There were two separate purge paths within the instrument, dictated by relative
contamination sensitivities.  The low-emissivity surfaces within the cooler are
most sensitive to contamination, consequently the cooler has its own purge
path, separate from the optics purge. 

Protective covers were installed over the telescope and cooler apertures, and
remained in place, except during calibration and thermal-vacuum testing.  They
were deployed 77 days after launch, hopefully after most spacecraft outgassing
had occurred. 

During launch and cruise, NIMS tends to be the coldest object on the
spacecraft, and is subject to condensation of water and other volatiles onto
sensitive surfaces.  Heaters were included in order to minimize any immediate
condensation, and to subsequently drive off condensates that might be deposited
during unheated periods.  The optics assembly has two 40 Watt heaters, operated
simultaneously, which produce temperatures of 240 - 250 kelvins.  This suffices
to remove water, but thruster products may be less easily removed, particularly
if they have remained on the surfaces and have reacted to form less-volatile
species.  A heater was also incorporated in the radiator shield (see Section
3.4).  It serves the same purpose, in this case protecting the most
contamination-sensitive surface on NIMS - the inner surface of the radiator
shield. 
 



                    4. Instrument Calibration 
 
4.1 Introduction 

For the NIMS instrument, there are three broad calibration categories:
spectral, radiometric, and spatial.  The majority of the calibration
measurements were performed in the laboratory, however there are several
important calibration verification activities that will be performed in flight,
and discussions of the relevant spacecraft hardware is included below. 

Unless otherwise noted, the measurements reported here were obtained in the
NIMS thermal-vacuum facility, which is a large stainless steel vacuum chamber,
evacuated with a liquid-nitrogen-baffled diffusion pump. The NIMS instrument is
mounted on an internal table which simulates thermal properties of the
spacecraft scan platform.  A liquid-nitrogen-cooled shroud surrounds the
instrument, allowing the optical assembly to radiatively cool to its flight
temperature - roughly 130 kelvins.  An additional space background simulator,
cooled with liquid nitrogen or liquid neon, was used to cool the radiator.  A
large area blackbody source was installed in the vacuum chamber and can be
rotated into the field-of-view of the telescope. External optical access is
provided by two interchangeable window assemblies, one being a single, large
diameter quartz window, the second consisting of a mosaic of small diameter
calcium fluoride windows. 

Calibration measurements were performed for a variety of operating conditions
and instrument modes.  The focal plane temperature, optics temperature,
electronics temperatures, and input power voltages were all varied over
expected operating ranges.  In addition, all appropriate instrument gain states
and modes were investigated. 

4.2 Spectral Calibration 

The goal of the spectral calibration is to establish the wavelengths sensed by
a detector for each of the 32 possible grating positions over the range of
conditions expected in flight.  This involves calibrating the NIMS spectrometer
itself, prior to launch, and also characterizing an internal spectral light
source, to be used in flight to detect any spectral shifts.  In addition, when
investigating highly detailed atmospheric spectra, it is necessary to know
spectral bandpass profiles in order to convolve theoretical spectra to the NIMS
resolution. 

The grating equation, when expressed for the NIMS optical geometry, reads

      m(lambda/d) = sin(gamma + phi) - sin(gamma - phi + chi), 

where m is the order of diffraction, lambda the wavelength, d the grating
constant, gamma is half of the angle separating the collimator and camera
optical axes, phi is the grating rotation angle, and chi is the angular
displacement of each detector from the optical axis, 

      chi-sub-i =arctan(x-sub-i/f). 

Here, x-sub-i is the linear displacement of detector i and f
is the effective focal length, and includes refraction effects by the cooler and
FPA windows. The grating rotation angle 

    phi = phi-sub-zero + eps(p - delta p)

with p being the grating position ranging from 0 to 31, eps is the grating
rotation increment, and phi-sub-zero is the grating offset for grating position
zero.  delta p is included to account for any launch induced shifts; it is
defined as zero for the laboratory calibration. Determination of these
constants, and any variation in them due to thermal and electrical effects,
constitutes one portion of the spectral calibration.  Several of these
quantities, in particular chi and x-sub-i were well determined during
fabrication and assembly, leaving only eps, f, and phi0 to be determined. 

In order to accomplish these measurements, an auxiliary monochromator was used,
which itself was calibrated using the HgI lambda 5461A green line in various
orders.  This monochromator, used in conjunction with an incandescent source, a
diffuser, and a collimator, was used to illuminate the NIMS telescope, forming
a diffuse, spectrally narrow image at the entrance slit of the NIMS
spectrometer.  For a given setting of the external monochromator, a spectrum
was obtained by the NIMS instrument. The wavelength setting of the auxiliary
monochromator was then changed by an interval small compared to the NIMS
spectral bandpass and the process repeated.  Thus, for each detector and
grating position, one can determine the spectral position and bandpass profile;
an example is shown in Fig. 11.  Least squares fitting of such data for a large
number of grating positions and detectors allows an accurate determination of
the calibration parameters, with an observed variance in wavelength of 0.001
microns, or less than one tenth of a grating step. Wavelength checks were also
performed using indene and polystyrene absorption features.

Temperature variations of the wavelength positions were determined using Hg
lines and by varying the optics assembly temperature using the instrument
heaters.  Comparing the relative displacement of the positions of atomic lines
shows that the grating step value, eps, varies by 0.03 % per degree centigrade,
an insignificant amount considering the observed constancy of inflight
temperature values. 

Spectral shifts could arise from launch induced vibration or thermal effects.
In order to quantify any such deviations, a spectral lamp is contained within
the instrument, mounted on the telescope spider, which can be exercised by
ground command.  The lamp is an InGaAs light emitting diode, which emits a
relatively narrow band of radiation, about 0.025 microns wide, centered at
0.8500 microns (at 130 K).  The center wavelength is slightly temperature
dependant, and there is a platinum resistance thermometer located nearby on the
spider which can be used to accurately determine the temperature and
wavelength.  Using this internal source, any spectral shifts can be measured to
an accuracy of better than 0.05 of a grating step. 
 
4.3 Radiometric Calibration 

The purpose of the preflight radiometric measurements is to determine the
parameters which relate the signal S received from the NIMS instrument (in data
numbers, DN, ranging from 0 to 1023) to the radiance I of the target.  The
response of the instrument was found to be linear, with S = S-sub-0 + sigma I
and where the sensitivity sigma depends upon several parameters, including
detector, wavelength, detector temperature, instrument gain, chopper mode, and
polarization of the source.  The dark value offset, S-sub-0, depends upon
detector, gain state, and other variables.  These dark values were
simultaneously determined from laboratory measurements but will also be
determined in-flight, before and after an encounter sequence. 

In order to cover the NIMS wavelength range of 0.7 to 5.2 microns, we used two
types of light sources, the first being an incandescent tungsten-filament
spectral irradiance standard which allows calibration to 2.5 microns. The
second source is the aforementioned extended blackbody source mounted within
the NIMS vacuum chamber, and provides useful spectral radiance for wavelengths
longer than 2 microns. Some details of the two different sources are given
below. 

The shorter wavelength measurements used a 1 kW filament lamp with a quartz
envelope containing halogen gas.  The spectral radiance of this source was
calibrated by EG&G Inc. and this calibration is directly traceable to the
National Institute of Standards and Technology (NIST). The lamp was powered by
regulated direct current at the prescribed amperage, measured using
NIST-traceable instruments.  In order to produce an extended source of known
radiance, a large-area Halon target was constructed according to the
prescription of Weidner and Hsia (1981), illuminated by the standard lamp.  We
use Weidner and Hsia's (1981) measured directional/hemispheric spectral
reflectance values and their bidirectional reflectance data to find the
reflectance for our particular geometry: normal incidence and ~ 20 deg emission
angle. This target was viewed by the NIMS instrument through the quartz window,
whose transmission was independently measured.  The lamp was placed at various
distances from the target, allowing a test of the linearity of response and
yielding a precise determination of the instrument sensitivity. The precision
of the sensitivity determination was found to be a fraction of a percent.  The
accuracy of the derived values is estimated to be about 10 percent, partly due
to uncertainties in the original lamp calibration (< 5%) and partly due to the
bidirectional reflectance of the Halon target. This target is currently being
compared to a Labsphere Spectralon standard, itself calibrated with
traceability to NIST. 

Measurement of the instrument sensitivity for the longer wavelength region was
performed using an extended blackbody source, with a diameter larger than the
NIMS telescope aperture.  This source is a V-groove radiator, electrically
heated and regulated to maintain a constant, preset temperature.  The physical
temperature of the radiating surface, which exhibits an emissivity of 0.99, is
measured by two copper-constantan thermocouples, each relative to the ice point
established by a distilled water ice bath. The thermocouples and potentiometer
used were all calibrated to NIST-traceable standards.  The heated target is
loosely thermally-coupled to a liquid nitrogen-cooled heat sink, allowing a
controllable temperature range of 200 to 350 kelvins to be achieved.   As with
the previous lamp, measurements were obtained for a variety of source settings,
in this case different temperatures, and the data fit in the sense of least
squares to find sensitivity values.  The precision in this determination is
generally a fraction of one percent.  Thermal gradients of ~ 0.5 kelvins  occur
over the surface, limiting the accuracy of the derived sensitivity to 5
percent. In the overlap region, 2.3 to 2.5 microns, the difference in
sensitivities found using the two different sources is 12.5 percent. 

The instrument uses a diffraction grating, which are known to show efficiency
differences for different polarizations. The polarization sensitivity was
checked throughout the entire operating range using dichroic and wire grid
polarizers. The maximum difference in sensitivity for two orthogonal
polarizations was found to be only 5 %, and this occurs only in the vicinity of
the grating blaze wavelengths, as predicted by scalar diffraction theory of
gratings (Strong, 1958). 
  
4.4 Spatial Calibration 

In this section we discuss aspects of the spatial calibration, which includes
the pointing geometry, the angular resolution and angular sensitivity profile,
and scattered light rejection.  The NIMS spatial scan pattern is twenty pixels
aligned along the cross-cone direction of the scan platform and formed by
stepping the telescope secondary mirror through twenty positions.  Measurement
of the angular location of each pixel was accomplished using an illuminated
slit, mounted in the focal plane of an external collimator.  By translating
this horizontally oriented slit in the vertical (cross-cone) direction one can
find the point of maximum response, and thus the angular location for each of
the twenty mirror positions.  It was found that the angular locations are well
represented as a linear function of the mirror position number, with an angular
step size of 0.528 mrad. 

Using a very narrow slit, either horizontally or vertically oriented, and
translating it in the orthogonal direction, one can obtain a measure of the
angular sensitivity profile in the spacecraft scan platform cone and cross-cone
directions, respectively.  An example is shown in Fig. 12.  It can be seen that
the response in cone angle approximates a rectangle, as would be expected since
this response is determined by the entrance slit of the NIMS spectrometer which
is in the telescope focal plane.  The response in the cross-cone direction is
more triangular, and is the natural result of spectrometer astigmatism, which
gives a slightly defocused vertical image of the detectors at the telescope
focal plane. 

The co-alignment of the Galileo scan platform experiments (NIMS, SSI, UVS, PPR)
was measured prior to launch.  The co-alignment of the NIMS and SSI experiment 
was found to be within 0.3 mrad, well within specifications.  Simultaneous
tests using stellar sources will be used to verify post-launch alignment.

For a variety of NIMS measurements, the instrumentally scattered light is of
concern.  Examples of observations which could be affected by scattered light
include limb scans, dark feature measurements, terminator scans, and Jupiter
dark side imaging. Although the best time to measure the scattered light
properties of the instrument is during flight, at which time any launch and
contamination influences will be included, prelaunch measurements of stray
light effects were also performed.  These tests, reported here, are only upper
limits to the actual performance, since additional contributions from the
collimator and both surfaces of the quartz window are included. Two different
tests were performed to evaluate the scattered light, the first being a "knife
edge" test and the second a "black hole" test. 

In the "knife edge" test, the mirror motion scans across a sharp boundary in
the focal plane of the external collimator, the upper half-plane being a
brightly illuminated surface while the lower unilluminated portion is made as
dark as possible using a black surface oriented to form a cavity.  For this
geometry, the near-field (1 to 5 mrad from the boundary) scattered light
contribution is detector dependent, varying from 0.1 to 0.4 percent.  The
disadvantage of this test is not knowing just how dark the lower boundary
really is, and is partially solved using the following test. 

In the "black hole" tests, the instrument and collimator are focused at the
center of a hole contained within a large, diffusely reflecting plate.  An
absorbing conical blackbody cavity is placed behind the plate, and the entire
assembly is illuminated by a source placed in front and displaced from the axis
of the conical absorber. Radiation reflected by the plate, which does not enter
the collimator, is absorbed by optically black baffles.  By using plates with
the same hole size but with varying albedos, radiation emanating from the
center of the hole is constant, while the instrumentally produced stray light
varies with the surface albedo. Using plates with various albedos and hole
sizes, the near-field scattered light response was found to be 30 % less than
that for the "knife edge" tests.  Again, all of these measurements include
extraneous contributions from the three intervening optical elements.
Definitive tests will be performed inflight, however preliminary results based
upon the Venus and Earth limb scans indicate excellent scattered light
rejection. 

4.5 Spacecraft Radiometric Calibration Target 
 
Calibration targets are provided on the Galileo spacecraft for inflight
calibration verification of the remote sensing instruments, and to monitor the
relative response throughout the mission by performing periodic calibration
observations.  All of the remote sensing instruments can use the Photometric
Calibration Target, discussed in the next Section.  In addition, there is a
Radiometric Calibration Target (RCT-NIMS) which is intended for verification of
NIMS performance in the long wavelength region, and is described in the
following paragraphs. 

The NIMS Radiometric Calibration Target is a near-field, extended, blackbody
source, mounted on the scan platform sunshade and in front of the NIMS
telescope when the scan platform is in the zero degree cone angle position. 
When used as a calibration source, the target is heated with 25 Watts of
electrical power, elevating the target surface to ~315 kelvins.  This provides
a known radiance which can be used for radiometric calibration for wavelengths
longer than 2.5 microns. The emitting surface, slightly larger than the
telescope aperture, consists of a mosaic of hexagonal honeycomb cavities, each
0.25 inch in width and 0.50 inch in depth.  An infrared black paint (Chemglaze
Z004) was applied, which, with the cavity geometry, gives a normal emissivity
of greater than 0.98 (Sparrow and Heinisch, 1970).  Thermally insulating rods
are used to mount the target, and a strip heater is employed to provide uniform
heating and temperatures. Temperature differences across the surface were
measured using thermocouples, with a root-mean-square deviation of 0.7 kelvins
being found. Upon heating, the target reaches equilibrium temperatures, within
0.5 kelvins, in a time period of 6 hours. 

The physical temperature of the target is measured using the spacecraft Command
and Data System (CDS) engineering telemetry.  Two temperature sensors are
employed, consisting of platinum and nickel resistance thermometers, and both
were calibrated with NIST traceability.  At the target operating temperature,
both sensors exhibit nearly the same resistance, roughly 550 ohms. In order to
calibrate the CDS circuitry itself, a temperature insensitive resistance of 562
ohms is also measured.  All three measurements use the same current source and
measurement circuitry.  Consequently, temperatures can be measured to an
accuracy limited only by the digitization interval, ~ +- 0.3 kelvins.  If we
assume a combined temperature error from all sources of +- 1 K, then
radiometric accuracies of 7 and 4 percent will be achieved at 2.5 and 5
microns, respectively.  Comparison of flight RCT data with pre-launch
measurements indicates stable instrument performance. 

4.6 Spacecraft Photometric Calibration Target 

The Photometric Calibration Target (PCT) and an associated optical element, the
Photometric Calibration Mirror (PCM) together form a source of diffusely
reflected solar radiation which can be used by the remote sensing experiment for
intra- and inter-instrument comparisons.  These two elements are mounted on the
Science Boom; the mirror reflects solar radiation onto the diffusing target
surface which is placed outboard from the mirror and in a position that can be
viewed from the scan platform.  In order that the target be illuminated over
the nominal range of solar cone angles, the mirror is convex, with a radius of
curvature of 46 cm  and located 53 cm from the target.  The reflecting surface
is vacuum deposited aluminum, with a protective overcoat formed by its natural
oxide.  The target surface is similar to the Voyager diffuser plate,
consisting of sand blasted aluminum. The combination produces a spectrally gray
diffuse surface, with an effective albedo of roughly 0.05.  They were
calibrated over the spectral range of 0.3 to 5.2 microns and for a variety of
incidence angles and azimuths.

 



               5. Operating Modes and Data Acquisition 

5.1 Instrument Modes and Operation 
 
For most Galileo observations, the time available is limited, and one must
tailor each observation for specific scientific goals.  For the NIMS
experiment, this translates into optimization of the spatial and spectral
sampling aspects. For example, atmospheric measurements usually require the
best available spectral resolution, whereas surface reflectance spectra are
generally broader, allowing coarser spectral sampling.  In addition, the
spatial coverage and resolution demands are quite different for Jupiter and
satellite measurements, the latter requiring much more rapid spatial sampling
during the short amount of time available.  Not only are there internal
spatial/spectral tradeoffs to be considered. but, in addition, it has been a
longstanding goal among the Galileo remote sensing experiments to perform
coordinated and compatible observations through simultaneous use of the scan
platform. With these considerations in mind, we have developed a flexible set
of instrument modes, described in this Section. 

The relevant NIMS instrument parameters that can be adjusted for differing
observations are mainly spectral, determining the number of wavelengths to be
sampled and their relative placement.  In one extreme, the entire spectrum is
obtained at full resolution, at the other extreme, the grating is fixed and
only one wavelength band is sampled for each of the seventeen detectors.
Intermediate combinations are possible, each with differing times required to
complete a spectrum.  Throughout this time, mapping is being accomplished by
scan platform motion, with spatial and spectral sampling occurring
simultaneously.  During the time required to form a complete grating motion
cycle, the scan platform will have moved some fraction of a NIMS spatial
resolution element (0.5 mrad). In the following, contiguous spatial sampling
corresponds to a motion of one spectral sample per spatial resolution element,
Nyquist sampling is twice as frequent. 

NIMS modes are implemented in the instrument software using parameter tables
(PTABs). There are two such tables, each describing a specific spectral
measurement sequence. Use of two PTAB tables allows for hybrid combinations,
giving flexibility to instrument sequencing.   The assignment of an individual
parameter table (PTAB) is shown below (Fig. 13) and described as follows:  N is
the number of grating positions per cycle, ranging from 1 to 24, D is the
grating angle step size, unity corresponding to a single step of
one-half of a spectral resolution element, and S is the grating start position.
The number of times to repeat a given spectral sequence is given by the
parameter R.  Additional parameters include M, mirror scanning (on or off), and
A, autobias (on or off). 

There are a total of sixteen modes available to NIMS, twelve of which are
pre-defined in the instrument read-only-memory (ROM).  An additional four
modes, yet undefined, can be placed in the instrument's random-access-memory
(RAM) via uplink commands.  All mode parameters can be changed by ground
commands.  The standard ROM modes are summarized in Table 2. 

The NIMS instrument modes were designed to be synchronous with spacecraft
timing and its various time units.  The largest unit is a RIM (or MAJOR FRAME)
which is 60 2/3 seconds, subdivided into 91 MINOR FRAMES, having whole
fractions of 1/91 (2/3 seconds), 1/26 (2 1/3 seconds), 1/14 (4 1/3 seconds),
1/7 (8 2/3 seconds), and 1/2 (30 1/3 seconds).  There are NIMS modes
synchronous with all of these fractions except 1/2, which is that used for a
compressed imaging mode for the SSI instrument. 




 
                    Table 2.  NIMS Standard ROM Modes. 

           The execution time, given in sec, is the time required
           to complete a grating cycle.  The Nyquist slew rate 
           is the scan platform cone angle rate necessary to move 
           through one-half of a NIMS pixel (1/2 of 0.5 mrad) in
           the execution time.  It is given in microradians/sec.
     
 Mode                                                Execution   Nyquist
Number           Name             N    D      M         Time    Slew Rate
====== ========================  ===  ===  =======   ========== =========

   0   SAFE                        0   0     0ff         1/60      15000

   1   FULL MAP                   12   2  Scanning      4 1/3         60

   2   FULL SPECTROMETER          12   2     Off        4 1/3         60

   3   LONG MAP                   24   1  Scanning      8 2/3         30

   4   LONG SPECTROMETER          24   1     Off        8 2/3         30

   5   SHORT MAP                   6   4  Scanning      2 1/3        110

   6   SHORT SPECTROMETER          6   4     Off        2 1/3        110

   7   FIXED MAP                   1   0  Scanning        1/3        750

   8   BAND EDGE MAP               2   -  Scanning      1 1/3        190

   9   BAND EDGE SPECTROMETER      2   -     Off        1 1/3        190

  10   STOP&SLIDE MAP           24,6  1,4 Scanning     Variable   Variable

  11   STOP&SLIDE SPECTROMETER  24,6  1,4    Off       Variable   Variable

 


The instrument steps through a grating cycle (N positions of the grating) then
repeats R times (Mode Repeat Count).  When the repeat count is consumed, then
the other PTAB controls the instrument mode.  This process continues as long as
the instrument is powered on. 

The instrument software has specific branching instructions which depend
mainly on the value of N, the number of grating positions.  These instructions
results in somewhat different timing for the scan mirror, as well as the number
of steps required to complete a cycle.  N = 12 results in 13 spatial scan
cycles (1/3 second each) and a total grating cycle time of 13/3 = 4 1/3
seconds. The scan mirror rests at the starting point during the 13th interval,
during which time the grating returns to its starting position.  If N = 24, then
this number of grating positions are positioned, followed by two scan mirror
rest times, during which the grating is reset to the origin,
giving a total cycle time of 26/3 = 8 2/3 seconds.
 
The NIMS instrument modes in ROM are named according the the following
conventions. The MAP mode indicates the spatial mirror is scanning,
SPECTROMETER indicates the spatial mirror is off.  LONG, FULL, SHORT (which is
sometimes called PARTIAL) and FIXED refer to the number of grating positions in
a single grating cycle (24, 12, 6, and 1 respectively).  The FIXED mode has
been embedded in spacecraft documentation as FIXED GRATING; however the
systematic name consistent with the definition in ROM would be FIXED MAP.  The
BANDEDGE mode has one grating position per PTAB, with a different grating
position in each PTAB, alternating between PTABS each 2/3 second.  The
STOP&SLIDE mode is a hybrid mode, combining the FULL mode in one PTAB with a
FIXED mode in the other. 
 
In the MAP modes the instrument has the spatial mirror turned on.  This yields
an effective field of view of 20 pixels, arranged in a linear stripe of 10 mrad
x 0.5 mrad in the cross-cone direction. These 20 pixels are measured
sequentially, and all within 1/3 sec. The initial scan motion, at the beginning
of a RIM, is downward (toward the scan platform), then up for the next grating
step. 

The SPECTROMETER mode, for which there is no internal spatial scanning,  is
used for certain observations, for example atmospheric limb scans, when only
one-dimensional image scanning using the scan platform motion is required, or
when redundant sampling of a given location can be used to advantage.  In
general, the scan platform rates are the same as those for the MAP modes, since
the same amount of time is required to complete a grating cycle.  The only
exception is between the FIXED MAP and SAFE modes. 

Among the various grating modes, the LONG mode provides the best spectral
resolution, with two spectral samples per spectral resolution element (spectral
Nyquist sampling).  This mode requires 8 2/3 seconds to complete, with seven
spectra contained in a RIM. For this mode, Nyquist spatial sampling requires
~30 microrad/sec slew rates, yielding two complete spectra while crossing
through a NIMS field of view (0.5 mrad).  It may be  difficult to match this
mode with simultaneous imaging (SSI) observations when time is the limiting
factor, since the slew rate for a nominal SSI frame is about 120 microrad/sec. 
One strategy is to use alternate imaging filters in a sliding mosaic. 
 
The FULL mode gives lower spectral resolution, with only one spectral
measurement per spectral resolution element, and is useful for diffuse solid
surface reflection features. This mode permits a coordinated observation with
other scan platform instruments using a slew rate of about 110 microrad/sec,
which is twice the preferred spatial Nyquist sampling rate for NIMS. 
 
The SHORT mode under-samples the NIMS spectral capability by a factor of 4,
using only 6 grating positions.  This mode is a compromise which was included
to work with the SSI camera's compressed mode and to permit obtaining some
spectral information when the time to complete a mosaic is extremely limited. 
Some optimization is possible by matching the starting grating position for a
cycle (and thus the spectral "comb") to spectral bands expected on the observed
surface. 
 
The FIXED mode reduces the number of grating positions to 1.  This mode is spectrally
minimal, but ensures excellent spectral registration for the wavelengths 
that are measured.  It proved to be very useful in Venus darkside observations.
 
The BANDEDGE mode alternates between two grating positions, providing two
stripes (20 samples at one grating position over the 17 detectors which are
approximately evenly spaced over the spectral range of the instrument) at
different grating positions.  This mode, for instance, alternates sampling
between the continuum and the maximum of a spectral feature.  It is useful for
spatial mapping a particular spectral feature at a high spatial scan rate. 
This mode samples one stripe every 2/3 second - 1/3 second for the stripe and
1/3 second preparing for the next grating position (defined by the next PTAB). 

The STOP&SLIDE mode is a combination mode - normally a combination of the FULL
mode and the LONG mode.  Its purpose is to provide a mode compatible with
multi-color sequences for the SSI framing camera.  The sequence works as
follows:  The scan platform is stopped and NIMS enters its highest spectral
resolution mode (LONG) while the SSI instrument is acquiring frames in several
filters. The scan platform then slews slowly over to the next picture position
(one overlapped SSI FOV) with NIMS in a mode compatible with the slew rate
(nominally FULL mode), in a time period which corresponds to the readout time
for the framing camera.  The Mode Repeat Count parameter is set to accomplish
this compatible sequence.  The result for SSI is a multi-color image. The
result for NIMS is a spatially complete spectral map of the target at modest
spectral resolution, and spatially sparse high spectral resolution samples with
an best spectral resolution.  This optimizes the use of scarce scan
platform time, characteristic of close encounter geometries, when multi-color
framing mosaics are being performed.
 
In addition to the ROM modes discussed above, four of the 16 instrument modes
may be loaded from RAM.  The PTABs for these modes are defined by ground
commands to the instrument and are stored in RAM. The RAM mode may then be
loaded into the active area with a single command. The RAM mode remains valid
as long as the instrument is powered on and the instrument is not reset. 
 
After the two PTABS which define a mode are loaded into the active area, the
active area may be modified to change the characteristics of the mode.  The
appropriate command modifies one of the four PTAB values in each loaded
PTAB (independently) with each invocation of the command.  This capability is
particularly important for the BANDEDGE and STOP&SLIDE modes.  The BANDEDGE
grating start positions and the STOP&SLIDE duration in the high spectral
resolution mode often need to be modified from those defined in the instrument
Read-Only Memory (ROM). 
 
5.2 Instrument Commands

In addition to the various modes that the NIMS instrument is capable of
executing (described above), there are a variety of instrument states which can
be used for measurement or calibration.  Of prime concern is the instrument
gain state, for which there are four (see Section 3.6).  In addition, two
different calibration sequences can be executed: an OPCAL command, for which an
internal electroluminescent diode provides a wavelength reference, and an ECAL
command, which injects a known signal into the electronic amplification chain. 
 
The commands are labeled by instrument number (NIMS=37) and a mnemonic, which
indicates the type of command sent.  One NIMS command will be executed in a
given spacecraft RIM; in the event of multiple commands, the last command
loaded will be executed.  The command must be loaded into the instrument
command buffer by minor frame 89 of the previous RIM.  The following is a brief
description of available commands. 
 
37IOP - Instrument Operation:  This command loads a NIMS mode (set of PTABS)
from ROM or RAM to the active area.  It also permits specification of the
grating start position (loaded into both active PTABS). 
 
37IST - Instrument Status:  This command modifies the gain state (4 available),
the chopper state (63 Hz, chopper off, chopper reference, and free run), and
can invoke the electronics calibration or the optical calibration. The
electronics calibration should be carried out in gain state 2. The calibration
lamp (a light emitting diode) measurement should be done in gain state 4 and
LONG SPECTROMETER mode. 
 
37MPT - Modify Parameter Table:  This command modifies one of the 4 parameters
in the PTAB for both of the PTABS in the active area.  The value for each PTAB
is specified independently in the command (but the same parameter is modified
in each PTAB).  It is possible to turn off the thermal channel autobias with
this command - a capability which is intended for use only where the focal
plane is approximately at room temperature. 
 
37GO - Grating Offset: This command sets the grating offset.  Acceptable 
values are 0 - 7; the default is 4 which is entered whenever the instrument
is turned on.  Note that this grating offset is different than the PTAB
grating start position parameter.

37SS - Special Sequence: This command programs the RAM modes.  It must 
be invoked twice, once for each PTAB in a mode.
 
37IRT - Instrument Reset:  This command permits resetting the instrument
without cycling instrument power.
 
37MN - Memory Normal:  This sets the instrument CPU to use the normal ROM
address space. 
 
37MRL - Memory Reallocate:  This sets the instrument CPU to branch to the RAM
address space. 
 
37PL - Program Load:  This permits loading programs into the spare RAM. 



5.3 Typical Encounter Operations
 
During each of the eleven Jupiter orbits, there will be satellite flyby
opportunities which occur on time scales of hours, demanding efficient usage of
the various instrument capabilities in order to maximize the scientific return.
Observations of the Jupiter atmosphere occur over longer time scales, measured
in days, but the number of important features to measured, their angular sizes,
and temporal measurement frequency also demands efficient instrument mode
usage.  The following discussion illustrates, in a simplified fashion, the NIMS
use of modes during a typical orbit at Jupiter. General sequencing priorities
are 1) keep NIMS in LONG mode as much as possible to optimize spectral
resolution, 2) maximize spatial coverage and resolution, and 3) include as many
scan platform instruments as possible in an observation - to maximize
synergistic science return. 
 
On a typical Jovian orbit, Jupiter becomes visible over the spacecraft sunshade
at about 25-40 Jovian radii (RJ).  At this distance it is possible to mosaic the
entire planet - at full spectral resolution - in a reasonably short time.  At
periapsis, occurring at about 10 RJ, the angular size of Jupiter is quite large,
and it is not possible to mosaic the entire planetary disc at full spectral
resolution. Similarly, for a close Galilean satellite encounter, it is possible
to fully map the satellite at highest spectral resolution at about 4 hours out,
but at closest approach the angular size and surface-relative smear rates
become very high, forcing a choice between spatial and spectral mode coverage. 
 
On the inbound portion of a Jupiter orbit, the NIMS instrument would be in LONG
MAP mode to mosaic the Jupiter day and night side, occasionally executing the
STOP&SLIDE mode for compatibility with SSI multi-color images of the planet.  
Many of the mosaics will be oversampled vertically, which is required for
overlapping fields-of-view for the thermal instrument (PPR).  As perijove
approaches, the LONG MAP mode is maintained for atmospheric measurements, but
the areal coverage is reduced to include only a few specific features.  These 
features will be consistently measured throughout the orbital pass, yielding
their temporal and phase-angle variations.  It may be useful to develop some
RAM modes for these atmospheric measurements, particularly if one is interested
in only a portion of the complete spectrum available to a detector.  It is
possible to maintain high resolution, i.e. using a grating delta parameter of
D = 1, but use a lesser number of grating positions N and an appropriate 
grating start position S to choose the spectral region of interest.  In doing 
so, the observations will encompass less time, allowing more features to be
examined.

For the satellite encounters, a full-disk mosaic will be acquired each time the
spatial resolution changes by a factor of ~2.  Thus, a LONG MAP mosaic would be
followed by a FULL MAP mosaic, which would be followed by a SHORT MAP mosaic,
as spectral resolution is traded for spatial resolution.  The STOP&SLIDE mode
would be invoked as the framing camera acquires it's multi-color mosaics.  At
closest approach, the FIXED mode would be utilized to compensate for the high
apparent rate of the satellite surface motion. 

 



                      6. Spacecraft Interactions 

6.1 Thermal Control 

Integration of the NIMS instrument into the Galileo spacecraft design involved
a number of new and unique problems for a remote sensing instrument.  The
greatest problem was minimizing spacecraft thermal loads on the instrument and
its radiator, so that it would be able to cool passively to the very low
operating temperatures desired.  Also, because the Galileo configuration
changed several times to accommodate Shuttle and upper-stage delays and
different launch opportunities, it was necessary to pay constant attention to
changes which might have negative impacts on the instrument's performance. 
 
One problem involved the location of NIMS on the despun scan platform, relative
to the other remote sensing instruments.  The desired location for NIMS was at
the right edge of the platform so that the side-mounted radiator had a clear
view towards space.  In addition, the end position prevented NIMS from being
sandwiched between two warm, thermally blanketed instruments, and allowed the
instrument body to radiate to space both above and to the side of it.  However,
this desire conflicted with the Photopolarimeter-Radiometer (PPR) which also
wanted the outboard platform position so that it would be able to view around
spacecraft structure down to relatively small angles from the Sun. The solution
was to mount the PPR on a downward extension of the scan platform so that it
was to the right of, and below NIMS, satisfying the requirements of both
instruments. 
 
During spacecraft assembly it was discovered that the PPR telescope barrel
extended into the field-of-view of the NIMS radiator.  Located at the end of
the Photopolarimeter/Radiometer (PPR) barrel was a thermally actuated hinge
mechanism which would have provided a significant thermal source for the NIMS
focal plane.  The PPR cover was rotated so that its actuating mechanism was out
of the FOV of the NIMS radiator.  The small fraction of the blanketed telescope
barrel still viewed by the radiator was estimated to cause an increase in focal
plane temperature by approximately 0.5 kelvins. 
 
The greatest thermal problem for NIMS came from the two radio-isotope
thermoelectric generators (RTGs) used to produce the spacecraft's electricity.
Mounted on booms on the spun side of the spacecraft, each RTG radiates more
than 4,000 Watts of heat.  The NIMS instrument body and radiator had clear
views of the RTGs.  Thermal modeling showed that when the radiator viewed the
RTG's spinning through its field of view, it would be heated to at least 120
kelvins, well beyond the operating range of InSb detectors.  The solution was
to implement RTG shades mounted on the booms, blocking direct views to NIMS.
The shades had to be carefully designed because the RTG's depended on a clear
view to space for cooling; if the shades reflected back too much thermal energy
the electrical power of the spacecraft would be degraded. 
 
Solar loads on NIMS at Jupiter were not expected to be a significant thermal
source.  Flight rules required that the NIMS radiator not be illuminated by the
Sun prior to observation sequences.  The scan platform sunshade provided
protection when the platform was stowed at 0 deg cone angle.  However, after
construction it was discovered that a small slice of the NIMS radiator was not
shadowed by either the radiator shield or the sunshade, when the platform was
at 0 deg cone.  The solution of this problem would nominally have been not to
leave the platform parked at low cone angles.  However, the spacecraft bus
sunshade added for the 1989 VEEGA mission provided shading of the radiator. 
 
Early flight experience with NIMS showed that the attention given to minimizing
spacecraft thermal loads was successful.  However, Venus operations revealed a
new problem.  Venus darkside observations by the PPR and NIMS required the
instruments to "shoot through the booms" of the spinning section, and allowed
the NIMS radiator to view a significant amount of warm spacecraft structure.
Because of the small heliocentric distance, some spacecraft structures were
much warmer than expected.  The first PPR observation, 16 hours before Venus
closest approach, left the scan platform at a cone angle of 27 deg, giving the
radiator a view of the back of the bus sunshade.  Apparently, the back of the
shade was warmer than predicted and the NIMS focal plane was heated from 64
kelvins to about 95 kelvins, beyond its desired operating temperature.
Real-time commands were sent to the spacecraft, which moved the platform to a
safe position, enabling the radiator to view deep space.  The focal plane
cooled to 86 kelvins by the time of the first NIMS observation.  Because the
NIMS observations of Venus were at larger cone angles than the initial PPR
observation, the radiator no longer viewed the bus sunshade, and the focal
plane did not heat beyond 88 kelvins during the critical nightside observations
of Venus. 
 
Flight rules were subsequently changed to require that the scan platform not be
left at low cone angles where the radiator could view spacecraft structure,
unless science observations were actually underway.  Tests were also planned
for calibrating the thermal load from the various spacecraft structures, using
NIMS itself to measure the temperature of each spacecraft element. 
 
6.2 Contamination Control 
 
A major concern of the NIMS experimenters was possible contamination of the
instrument from various spacecraft sources.  Because it was unblanketed and
because it operated at such low temperatures, it was feared that NIMS would
serve as a cold trap for volatiles outgassed from the spacecraft and from
thruster plume byproducts.  In addition, the performance of the NIMS radiator
shield was closely tied to maintaining the emissivity and reflectivity of the
shield surface -- contamination could seriously threaten the ability to cool
the NIMS focal plane. 
 
An examination of past experience with Viking and Voyager showed several
unexpected problems of this type.  For example, outgassing of the Viking lander
capsule was so severe that it resulted in nongravitational accelerations on the
spacecraft orbit for the first six months after launch.  On Voyager, thruster
plume impingement on spacecraft structures resulted in a 20% inefficiency in
spacecraft maneuvers. 
 
Most alarming to the NIMS experimenters was data that came from the Infrared
Thermal Mapper (IRTM) and Mars Atmospheric Water Detector (MAWD) on Viking.
These instruments had a common calibration target which consisted of a small
sandblasted aluminum plate, illuminated by the Sun.  It was not possible for
either instrument to view the target prior to deployment of the Viking landers.
When they were finally used, it was found that each target had decreased in
albedo by several percent, and continued to decrease over the life of the
mission.  Spectral data from the two instruments allowed the investigators to
determine that the contaminating material was reddish in color. 
 
Numerous sources for the contamination were considered, such as dust in Mars
orbit, but were rejected.  The studies concluded that the most likely
contaminant was byproducts from the Viking main propulsion engine.  This engine
used a bipropellant combination of monomethyl-hydrazine and nitrogen tetroxide,
and one of the byproducts was monomethyl-hydrazine nitrate, a dark reddish
material with extremely low volatility, particularly after exposure to
sunlight.  Surprisingly, the IRTM/MAWD cal target did not have a direct view of
the Viking main engine, and was located about 160 degrees from the engine
thrust centerline.  However, plume expansion in vacuum was known to carry
contaminants to locations behind the main engine, and even able to expand
around spacecraft structures. 
 
Galileo used the same type bipropellant system as Viking, not only for the main
engine but also for the attitude control thrusters; Viking used a cold gas
attitude control system.  Extensive studies were undertaken by the Galileo
spacecraft team to quantify the expected contamination from thrusters and the
main engine.  These studies showed that significant contamination could occur
on the scan platform instruments and on the calibration targets.  As a result,
a number of protective measures were instituted. 
 
The first protection was shields around the thruster clusters to prevent a
direct line-of-sight to the scan platform, and a similar shield between the
platform and the 400 newton main engine.  Secondly, minimal contamination
positions were determined for the platform for either thruster or main engine
firings.  The platform was commanded to the respective safe positions prior to
any thruster use or maneuver.  Third, the thrusters on the spinning section
were limited to firing during passage through relatively narrow arcs centered
approximately 90 deg from the despun scan platform position.  Fourth, the NIMS
and calibration target heaters were kept on whenever possible to prevent
condensation of volatiles on the surfaces, and to sublimate any materials that
had condensed.  This latter measure also provided protection from spacecraft
outgassing early in the mission.  The primary outgassing product is water from
thermal blankets. 
 
Deployable covers were incorporated into the NIMS design for both the telescope
and the radiator, fired by a single, redundant pyro device.  The covers were
retained until 77 days post-launch, providing protection during the initial
period of spacecraft outgassing, the first two trajectory correction maneuvers,
and during many other early pyro events on the spacecraft.  In addition a dry
nitrogen purge was provided for the instrument during ground testing and
spacecraft assembly.  This purge was maintained during the Shuttle launch and
in Earth orbit, until a few minutes before deployment of the Galileo/IUS stack
from the Shuttle bay. 
 
6.3  Instrument Pointing
 
NIMS is a whisk-broom imager with one spatial dimension being provided by its
scanning secondary mirror, and the other dimension provided by either
spacecraft motion or movement of the scan platform.  The ability to move the
scan platform very smoothly at a predetermined rate relative to a target was
thus very important for the success of the experiment.  This was a somewhat new
concept to the attitude control engineers on Galileo whose previous experience
was with imaging systems that typically worked in a "stop-and-shoot" mode. 
 
Numerous discussions were held with spacecraft engineers to perfect the
pointing requirements for NIMS and the other remote sensing instruments.  In
addition, discussions were held with the other remote sensing science teams to
determine compatible instrument modes where all instruments could take data
simultaneously.  Both the UVS and the PPR were also interested in taking data
in a continuous slew mode, whereas the Imaging system still preferred a
"stop-and-shoot" system.  Scan platform capabilities and instrument modes were
developed to accommodate both situations. 

The nominal FULL MAP operating mode for NIMS called for a continuous Nyquist
slew rate of 60 micro rad/sec.  Other desired rates were 30 micro rad/sec for
LONG MAP, 110 micro rad/sec for SHORT MAP, and 750 micro rad/sec for FIXED MAP.
These are all relatively slow rates compared to the maximum scan platform
capability of 17500 micro rad/sec (i.e. a degree per second).  Thus, the scan
platform performance was optimized for slew rates less than 3000 micro rad/sec.
Maximum allowed deviation from a desired slew path was set at 125 micro rad,
or one-quarter of a NIMS pixel. 
 
Since some satellite flybys could result in fairly large target smear rates,
target motion compensation was added to the pointing system.  This allowed the
spacecraft attitude control system to take out relative target motion without
the need for extensive ground design of sequences. 
 
6.4 Spacecraft Obscuration 

The dual spinner design of Galileo led to a situation where much of the sunward
hemisphere of the sky is often obscured by spacecraft structure.  The
obscuration of NIMS by the spacecraft is shown in Fig. 14.  It is not possible
to synchronize instrument operation with the spinning booms, so observations
that view through the spinning structures must be repeated to fill in gaps
caused by boom obscuration.  Although NIMS can view down to less than 30 deg
cone angle, limited by the bus shade, obscuration by other structures is
considerable at that angle. 
 
Considerable effort was devoted to keeping the anti-sunward hemisphere
completely free of obscuring structure.  The only spacecraft part that extended
into this hemisphere was the ends of the Plasma Wave Spectrometer (PWS)
antenna, located at the end of the magnetometer boom.  However, the switch to
the 1989 VEEGA mission required a second low gain antenna on the aft end of the
spacecraft.  This was accomplished by hanging a deployable boom off one of the
RTG booms, and extending down to about 114 degree cone angle.  This additional
obscuration was accepted because LGA-2 would not be needed after the second
Earth gravity assist flyby, and the boom would be folded out of the way and not
used again. However, Venus and Earth observations near 90 degree cone angle
were degraded by this additional obscuration. 
 
Careful attention was paid to minimizing sources of glint from spacecraft
structures, to minimize scattered light during observations.  Since most of the
spacecraft was expected to be covered in black thermal blankets, this was not a
very difficult problem.  However, the switch to the 1989 VEEGA mission required
the addition of gold foil blankets to many sun-facing surfaces to minimize
heating at the expected post-Venus perihelion distance of 0.69 AU.  Integration
of these changes with science concerns was not as optimal as it had been
earlier in the mission, though science suggested changes were accommodated in
many cases. During the Venus flyby,  glint was observed from the edges of some
of the spacecraft booms, however it was weak and from very narrow regions, and
did not degrade any of the data. 
 



                    7. NIMS Mission Design Aspects 

Design of the NIMS experiment required the opportunity to view targets at
optimum geometry for obtaining spectral images.  For solid surfaces, this meant
viewing at low phase angles, preferably less than 30 deg phase, where shadowing
of target surfaces would be minimal.  This conflicted sharply with Imaging Team
desires to view surfaces at phase angles near 75 deg where shadowing would help
to highlight surface topography.  Another conflict involved the NIMS team
desire to emphasize global mapping of satellite surfaces at resolutions of 25
km/nimsel or better, whereas the Imaging Team was interested in obtaining the
highest possible resolution of surface features, at better than 1 km/line-pair.
 
The approach and departure hyperbolae to satellite flybys afforded good
opportunities for global mapping by both investigations.  However, close
flybys, desired by both Imaging and by the fields & particles investigations,
made it impossible to view the entire sub-spacecraft hemisphere.  In addition,
there was not sufficient time to image the sub-spacecraft hemisphere at closest
approach because of the high flyby velocity and high smear rates. 
 
The solution to many of these problems involved "non-targeted" or "accidental"
flybys of satellites.  Because of the resonant motion of the inner three
Galilean satellites, the spacecraft would often fly close, within about 105 km,
to one satellite when being targeted to a gravity assist encounter with
another.  If these passes occurred on the sunlit side of the satellite they
provided good opportunities for global mapping at resolutions of better than 50
km/nimsel, or 2 km/line-pair for the SSI.  In some ways these non-targeted
flybys were even better than global mapping on the approach hyperbolae to
targeted encounters, because the range to the target changed more slowly, and
there was a greater possibility for viewing at relatively low phase angles. 

Again, interaction with the engineers designing the satellite tour at Jupiter
was an important factor in optimizing the mission for NIMS.  Analysis programs
were written to estimate mapping coverage for each satellite prospective tour,
and these results were compared and suggestions for improvements made.  This
was a difficult exercise because of the limits of the trajectory changes
possible using gravity assists, and because of the large number of science
requirements from the different investigation teams. 
 
Jupiter observation geometries were largely dictated by the satellite tour
selected, though they were also considered in the tour design.  Because the
planet covered such a large solid angle near spacecraft periapsis, special
sequences needed to be worked out to allow all the remote sensing instruments
to cover a maximum area of the planet simultaneously. 
 
Flyby trajectories at Venus and the Earth were optimized for gravity assist and
had to be used by the science instruments without modification.  Even with
these restrictions, it was possible to construct valuable science sequences.
These sequences proved to be very useful in conducting complete tests of the
Galileo sequencing, commanding, downlink, and data reduction systems.  Many
problems were discovered and solved which might have not been detected until
much later in the nominal mission. The value of exercising the spacecraft and
instruments on real targets, prior to arrival at Jupiter, cannot be overstated.

 
 


ACKNOWLEDGEMENTS

Development of the NIMS instrument, along with incorporation of necessary
spacecraft accomodations, involved the work of many people, whose unique
contributions ranged from project management to craftsmanship in soldering. All
of these skills are necessary to NIMS' success, and it is unfortunate that we
cannot cite each individual for their own contributions.  We would, however,
like to specifically recognize and thank members of the initial Galileo Project
management team who provided encouragement and advice during the development
stages of NIMS: John Casani, Ron Draper, Jesse Moore, and Bill Fawcett.  This
work was supported by NASA Contract NAS 7-100 to the California Institute of
Technology, Jet Propulsion Laboratory. 





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