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MSX (Midcourse Space Experiment)

The MSX satellite is an observatory-class observation platform and global surveillance system of the US DoD; it is funded and managed by BMDO (Ballistic Missile Defense Organization) with JHU/APL (Johns Hopkins University/Applied Physics Laboratory) of Laurel, MD as the prime contractor, spacecraft integrator, and operator of the mission. MSX represents the first system demonstration in space of technology to identify and track ballistic missiles during their midcourse flight phase. The suite of optical sensors cover the spectrum from the far ultraviolet (110 nm) through the very-long-wave infrared (28 µm) spectrum. 1)

The primary objectives of MSX are to detect, acquire and track targets and to discriminate lethal from nonlethal objects (detailed characterization and modeling of target objects and their associated phenomenology of the terrestrial, Earth-limb, and celestial backgrounds). The information gathered by the MSX spacecraft will help fill significant spatial, spectral, and temporal gaps that currently exist in space environment models. Another driving factor in the MSX Program is sensor and spacecraft technology improvement. The latest technology is used to provide the most up-to-date phenomenological information and to serve as a new-technology demonstrator. 2) 3) 4) 5) 6)

In addition to meeting the needs of the BMDO, the MSX spacecraft provides a civilian benefit owing to its multispectral and hyperspectral capabilities. Data collected from BMDO-related experiments can be used to perform environmental studies of the Earth. Special environmental monitoring experiments can also be conceived and performed.


Figure 1: Artist's view of the MSX spacecraft in orbit (image credit: JHU/APL)



The MSX satellite is three-axis stabilized consisting of the structure, five primary instrument systems, and the subsystems needed for instrument support and S/C control. The fundamental configuration was required by the size and thermal requirements of SPIRIT-III (Spatial Infrared Imaging Telescope-III). The spacecraft structure is divided into three discrete elements consisting of (Figure 2): 7) 8)

1) An electronics section, which provides an interface to the Delta II launch vehicle and a mounting platform for instrument and spacecraft electronics.

2) Truss section: A thermally stable graphite/epoxy composite truss structure for mounting the SPIRIT-III instrument and providing an interface between the instrument and electronics sections. A major design driver for the thermal team was to keep the average temperature of the SPIRIT-III outside shell below 250 K, with a goal of 225 K.

3) Instrument section: A temperature-controlled instrument section with embedded heat pipes for mounting delicate sensors and instruments. The instrument section also provides a thermally stable mounting structure for the optical bench and the beacon receiver bench.

The instrument section also contains two ancillary structures that play an important role in the tracking and attitude control of the spacecraft. The optical bench is a precision inertial measurement platform that is attached to the top of the instrument section structure. The G/E (Graphite/Epoxy) composite bench supports the ring laser gyroscopes and a star camera for establishing and maintaining spacecraft attitude control. The G/E composite bench contains four S-band receiving antennas that can acquire and lock onto a target signal beacon at frequencies of 2219.5 and 2229.5 MHz during spacecraft tracking activities.

The instrument section is a four-sided box structure that provides mounting locations for most of the spacecraft sensors. The entire section is blanketed. It contains the SBV instrument telescope, the UVISI instruments, and the contamination experiments. The section is cut away in the middle to allow for the SPIRIT-III telescope.

The MSX spacecraft has a design life of five years, its body measures approximately 160 cm x 160 cm x 520 cm, and its launch mass is about 2812 kg.


Figure 2: Two views of the stowed configuration of the MSX spacecraft (image credit: JHU/APL)


Figure 3: Functional block diagram of the MSX spacecraft (image credit: JHU/APL)

The ADCS (Attitude Determination and Control Subsystem) consists of four reaction wheels and three magnetic torque rods. Any three of the four wheels can provide three-axis control of the satellite. Attitude sensors include two three-axis ring laser gyro systems, a star camera, two horizon sensors, five digital sun-angle detectors, and a three-axis magnetometer. The system achieves a real-time pointing accuracy of better than 0.1º and post-processing knowledge of 9 µrad over instrument integration durations of about 1s.

Coarse attitude sensors

- MAG - a 3-axis magnetometer measured the Earth's magnetic filed
- HSA (Horizon Scanner sensor Assembly); 2 units, 45º scan cone~1º FOV to detect the Earth's horizon at two points/scan, HSAs 180° apart in Y-Z plane
- DSAD (Digital Sun Angle Detector); 5 heads in full sky coverage

Fine attitude sensors

- ST (Star tracker); measured up to five stars
- ID by ADS processor; initialized by a coarse solution

Rate sensors

- RLG (2x) - IRUs (Inertial Reference Units), ring-laser gyroscopes


- RWA (Reaction Wheel Assembly) with 4 wheels
- Two 3-axis magnetic torque coils for the unload of excess momentum

AP (Attitude Processors)

- Dual 1750 computers (2 x)

Tracking processors

- 1750 computer (2 x) - all scenario calculations


- Park & Track - coarse or fine attitude knowledge or propagated RLG data
- Safe - coarse sensors, wheels, (RLGs) - point solar arrays at sun, instruments to zenith if possible

Table 1: Overview of the ADCS


Figure 4: Overview of the ADCS instruments on MSX (image credit: JHU/APL)

EPS (Electric Power Subsystem): An unregulated direct energy transfer topology of 28 VDC.. The solar array (1.2 kW BOL) consists of two wings, each comprising four solar panels supported by a boom and attached to a SAD (Solar Array Drive). The battery consists of a series stack of 22 rechargeable nickel-hydrogen (NiH2) cells with a capacity of 50 Ah. Redundant power management modules (PMMs), which monitor the battery and provide the main EPS command and data handling (C&DH) subsystem interfaces; and the dual shunt control electronics (DSCE), which control the analog and digital shunts and contain the low voltage sense system (LVSS) circuitry. 9)


Figure 5: Block diagram of the MSX electrical power subsystem (image credit: JHU/APL)

C&DHS (Command and Data Handling Subsystem): The command subsystem consists of command processor, data handling, key generator (KG)/ tempest, power switching, and tape recorder components. Commands transmitted from ground stations are acquired via the S-band transponders.

Spacecraft science and housekeeping data are gathered by the DHS and formatted to generate three serial data outputs. Each stream has selectable formats and can be turned on and off independently as required by the mission. These three streams are the following: 10)

1) 25 Mbit/s (or 5 Mbit/s) prime science data stream containing imager, processor, and housekeeping data; transmitted in real time or stored on one of the spacecraft tape recorders

2) 1 Mbit/s snapshot science and housekeeping wideband downlink data stream; transmitted in real time

3) 16 kbit/s narrowband downlink data stream containing spacecraft housekeeping and memory dump data; transmitted in real time.

The C&DHS subsystem includes data encryption and decryption units and two tape recorders of Odetics, each with a storage capacity of 54 Gbit.


Figure 6: Functional block diagram of the C&DHS (image credit: JHU/APL)

Tracking function of targets: The target tracking experiments represent the primary mission of the MSX; they are in fact the main driver of spacecraft design. These experiments last about 30 minutes, with all sensors powered and collecting data and the spacecraft slewing at high rates to follow ballistic missile targets. 11)

The tracking experiments sized the spacecraft power, attitude, and data handling systems. There is also a tracking requirement which called for the capability of the spacecraft of performing a full target tracking experiment without solar input, to permit experiments to be done in eclipse phase of the orbit or with the solar panels in an unfavorable solar attitude position. Hence, a 50 Ah NiH2 battery was chosen to supply the needed power for such a target tracking event.

The MSX incorporates another function not normally seen in spacecraft: closed loop tracking on targets other than stars. As seen in Figure 3, a TP (Tracking Processor) can take sensor information from several sensors to acquire and closed-loop-track a target. These sensors are identified in the following list:

• Beacon receiver. This S-band passive radar tracker tracks telemetry transmitters on target vehicles. It has a much wider FOV (Field of View) than the optical instruments, and it guides spacecraft attitude, through the TP and attitude system, to point the target to well within the smaller field of view of the optical sensors.

• UVISI (Ultraviolet and Visible Imagers and Spectrographic Imagers) narrow FOV visible imager, a 1.3º x 1.6º visible band imager.

• UVISI narrow FOV ultraviolet imager, a 1.3º x 1.6º ultraviolet imager.

• UVISI wide FOV visible imager, a 10.5º x 13.1º visible band imager.

• UVISI wide FOV ultraviolet imager, a 10.5º x 13.1º ultraviolet imager.

• SPIRIT-III radiometer. Band A or D can route images to an experimental target processor called the Onboard Signal and Data Processor, which routes target tracking information to the spacecraft TP.


Figure 7: Block diagram of the tracking processor and interfaces (image credit: JHU/APL)

RF communications: The MSX communicates on three bands: S-band uplink (1827.8 MHz), S-band downlink (2282.5 MHz), and X-band (8475 MHz). The S-band uplink is used exclusively for commanding and memory loads (data encryption). The S-band has two transponders, diplexers, and antenna pairs for the transmission of 16 kbit/s housekeeping data and either ranging or 1 Mbit/s transmission of compressed or sampled science data.

The X-band (two transmitters) transmits 25 (or 5) Mbit/s prime science data from the S/C data recorders (54 Gbit capacity each).

Beacon receiver antenna bench: This device is a 0.4 m2 panel that houses a four-parabolic-dish, phased-array antenna. It is folded down to stow against the side of the instrument section during launch and is released and pinned in place so its antenna beams are pointed in the same direction as the optical instruments. The S-band beacon receiver is a passive radar tracker that can acquire a target with an initial pointing uncertainty of ±5º at a maximum range of 8000 km. It provides calibrated angle tracking with a residual error of 0.1º.


Figure 8: Photo of the MSX spacecraft with the instrument section at the top (image credit: JHU/APL)

Launch: The MSX spacecraft was launched on April 24, 1996 with a Delta-2 7920 vehicle from VAFB (Vandenberg Air Force Base), CA, USA.

Orbit: Near sun-synchronous polar orbit, altitude = 897 km x 907 km, inclination = 99.4º. The daily precession rate is < 0.04º.
This orbit allows the spacecraft to fly in a roughly Earth-oriented attitude while keeping direct sunlight out of the instrument apertures. Controlling the spacecraft's attitude relative to the sun is one requirement that shapes the whole mission.


Mission status:

• In July 2008, after more than 12 years of successful operations and contributions to two diverse defense missions, the MSX spacecraft was retired, having operated well beyond its design life of 4 years. The spacecraft collected vital data for designing missile defense systems. With no fuel onboard, reentry of the spacecraft will take place in a few centuries. 12) 13)

Decommissioning the MSX spacecraft did not consist of a series of deorbit maneuvers, as is the practice for some satellites, because it had no propellant. Nor was it as simple as just powering down the spacecraft components and walking away because the design of the spacecraft included a level of autonomous survival methods that were hard-coded into the flight software. Any decommissioning plan needed to account for the flight software design and to defeat those survival actions. 14)

The decommissioning goal was to place MSX in a state such that it would not generate any RF signature and would minimize any chances of debris. Because there was no propellant or other expendables on the spacecraft, the decommissioning planning revolved around depleting the MSX power subsystem to a level at which there would be insufficient energy to support any subsystem operation and insufficient energy for the spacecraft to autonomously reconfigure itself for survival.

The decommissioning of MSX marked the end of one of the most successful satellite programs ever undertaken by APL. The MSX spacecraft had fulfilled not only an ambitious primary mission, but it had also provided an abundance of useful data for a time well beyond its designed lifetime. The 12 years of active operations that followed years of development and testing at APL were filled with frequent adaptation to changing spacecraft capabilities as well as to sponsor needs. These changes presented challenges to the operations teams throughout the program, including during the decommissioning process (Ref. 14).

The objectives of the primary mission (of 1 year duration ) were: 15)

- Track ballistic missile targets to study phenomenology

- Survey horizon “clutter”

- Track rocket & missile launches

- Infrared star survey for calibration & cataloging

- Sensor performance evaluation

During the program’s primary mission, the first year in orbit, it was determined that the RLGs had a much more limited life than expected due to a surprisingly rapid decline in the intensity of one laser in each RLG. A decision was made quickly to use the RLGs only for those events that required the highest degree of pointing accuracy and stability [i.e. SBV tasking DCEs (Data Collection Events)]. Without gyroscopes, the MSX needed to rely on other sensors for attitude rate knowledge.

The objectives of the secondary mission were: support of the ACTD (Advanced Concept Technology Demonstration) program - to demonstrate space surveillance using the SBV (Space Based Visible) telescope. Since MSX was being used only for space surveillance after the primary mission, a joint effort by APL and MIT/LL developed a method for continuing these DCEs that did not require gyro data for attitude determination and maneuvering.

Sensor loss mitigation strategies during the secondary mission:

- SBV operations team designed maneuvers to keep HSA1 & HSA2 on horizon at all times during surveillance observations (the objective was to maintain coarse attitude)

- HSA1 failure: Operations team designed “roll-over” maneuver at ascending node to keep HSA2 on horizon while maintaining SBV at Geo-Belt (maintain coarse attitude)

- Surveillance performance gradually degraded over time (ST suspected, but analysis did not pin-point problem)

- HSA2 erratic behavior occurred in early 2007: The consequences were to a) rely on ST much more -slues and tracks, b) the loss of fine attitude corrupted SBV observations, and c) prompted an ST in-depth analysis & control system tune-up

- Productivity recovered well but termination decision had been made

- However, very soon after implementing all these changes, MSX suffered near-simultaneous failures of HSA2 and the primary Attitude Processor (AP1). This meant the end of any further mission operations. Since a safe and controlled shutdown was essential, carefully planned and phased procedures were executed in mid-2008 to drain the battery and terminate the mission.

In conclusion: a) The mission turned out to be very successful in spite of unspecified long-term requirements; b) the performance was good through most of the operational phase. c) the operations, instrument, and G&C teams worked together to mitigate problems and failures; and d) limited support for operations and engineering team throughout mission likely had impact on longevity.


Figure 9: MSX spacecraft status as depicted on January 26, 2008 (image credit: JHU/APL)

• On April 24, 2006, the AFSPC (Air Force Space Command) as well as the other MSX partner organizations celebrated the 10th anniversary of the MSX spacecraft launch. At the time, the SBV instrument provided its services by collecting “Space Situational Awareness” data as the only system able to “see space from space.” The satellite had also contributed to scientific pursuits to include various astronomical experiments on global change of atmospheric gases, studies of the chemistry and physics over the poles, gathering data on space contamination and debris and even galaxy phenomena to include the Hale-Bopp comet and Quasars. 16) 17)

• After completing BMDO's mission (4 years), the spacecraft was transferred to AFSPC in October 2000, becoming the Air Force's first operational spaceborne sensor to track and monitor objects in orbit around Earth. AFSPC realized the potential of the space surveillance capabilities inherent with the SBV (Space Based Visible) sensor and assumed ownership from BMDO on Oct. 2, 2000. From here on, AFSPC managed the program while APL continued with spacecraft operations, and MIT/LL continued the operations for the SBV instrument, the only sensor in use onboard the MSX at the time. - The SBV provided full metric and SOI (Space Object Identification) coverage of the geosynchronous belt regardless of weather, day/night or moon light limitations. 18)

• During the Leonid meteor shower on Nov.18 1999, the five UVISI spectrographic imagers onboard the MSX satellite recorded the first complete meteor spectra from 110 to 860 nm. 19)

• The cryogen period included the time from launch through the lifetime of the SPIRIT-III cryogenic telescope. MSX was launched with the SPIRIT-III telescope already cold. The cryogen period lasted for approximately 10 months and ended in early 1997 when the dewar, containing solid hydrogen, warmed up to a temperature above 12 K.
The UVISI suite of instruments, not requiring such cooling, continued its observation services. Between April 1996 and March 2000, the UVISI instrument suite observed roughly 200 stellar occultations using a combined extinctive and refractive technique developed at JHU/APL. 20) 21)


Figure 10: Orbital deployed configuration of the MSX spacecraft (image credit: JHU/APL)


Sensor complement: (SPIRIT-III, UVISI, SBV, Contamination Experiments, OSDP)

The suite of state-of-the-art sensor systems include: SPIRIT-III (a cryogenic scanning radiometer and FTIR spectrometer), the UVISI system, SBV, and a suite of instruments to monitor contamination on and around the S/C. MSX is a fully steerable spacecraft providing a range of pointing capabilities (0.1º pointing accuracy): limb scans, nadir scans, or limb point and stare, to cross-track scans. Of importance to the science community are stellar occultation and daytime limb observations for the retrieval of trace gases: ozone, aerosol, pressure, temperature and NO2 in the stratosphere. 22) 23)

Some of the sensor data (mostly concerning targets and satellites) will be classified, but the majority are unclassified. These latter data will eventually be released to the general science community after the normal period (~2 years) reserved for exclusive exploitation by the principal investigators.


SPIRIT-III (Spatial Infrared Imaging Telescope):

SPIRIT-III was developed and built at SDL (Space Dynamics Laboratory) of Utah State University. The instrument package is the primary sensor system on MSX with the objective to perform midcourse surveillance functions, collect target and background phenomenology data, and to demonstrate advanced cryogenic sensor technologies. The requirements are to characterize Earth-limb and celestial backgrounds and measure the spectral, spatial, temporal, and intensity parameters of stars and upper atmospheric phenomena, including airglow and aurora. The goal is also to characterize the signatures of selected targets operating against natural backgrounds and to assist investigators in identification and evaluation of the chemical and physical properties of the upper atmosphere on a global scale.

The instrument consists of an off-axis re-imaging telescope with a 35 cm diameter unobscured aperture, a six-channel high-spectral-resolution Michelson interferometer, a five-band scanning radiometer, and a cryogenic dewar/heat exchanger (the dewar cools the telescope, radiometer and spectrometer, the cryogen is expected to last for 15 months, cryogen volume of 944 liter). Total instrument mass = 967 kg; size = 107 cm diameter x 360 cm in length. 24)

Note: SPIRIT-I and -II were instruments (FTIR spectrometers only) flown on sounding rockets in the late eighties and in 1991 - as part of the SDIO (Strategic Defense Initiative Organization) heritage.


Figure 11: Photo of the SPIRIT-III instrument package prior to spacecraft installation (image credit: SDL, JHU/APL)


Figure 12: Mechanical configuration of the SPIRIT-III instrument (image credit: SDL, JHU/APL)

The sensor components are conductively cooled and maintained at operating temperatures via thermal links, connecting them to the solid-hydrogen tank of the cryostat subsystem. Power consumption ranges from 50 to 410 W, depending on the mode of operation. The optical components of SPIRIT-III telescope operate at temperatures ranging from 10 to 20 K. The cryosat subsystem was built by Lockheed Martin of Palo Alto, CA under contract to SDL.

Optics (Figure 13): The telescope has three sections: the afocal foreoptics (M1, M2, M3), the radiometer re-imaging optics, and the spectrometer collimating optics. The foreoptics use an off-axis all-reflective design. An auto-collimator measures telescope alignment with respect to the S/C attitude system optical bench to an accuracy of 5 µrad. The Michelson interferometer (Fourier-transform spectrometer) has six Si:As detectors operating at 10.5 to 11.0 K. It collects double-sided interferograms in six spectral bands with programmable resolution of 2, 3.9, or 20 cm-1 over sample times of 4.2 s, 2.2 s, and 0.55 s.

Radiometer: The radiometer has five Si:As focal plane detector arrays of 8 x 192 pixels each, operating between 11 and 12 K. It collects data in six passbands with a spatial resolution of 90 µrad. The scan mirror can remain fixed or can operate at a constant scan rate of 0.46º/s with programmable scan fields of regard of 1º x 0.75º, 1º x 1.5º, or 1º x 3º. The radiometer focal plane assembly employs a combination of dichroic and bandpass filters to allow simultaneous measurements in bands A, D and E and or in bands B and C. Table 3 lists the four operational modes and provides actual values for the radiometer field of regard (FOR). - Earthlimb observations of SPIRIT-III include an experiment to measure chlorofluorocarbons and nitric acid. Channel 5 (10.6-13 µm) of the FT spectrometer collects data in the altitude range from 10 to 40 km. Dominant emissions in this channel include nitric acid, CFC compounds, ozone, and the thermal signature of stratospheric aerosols.


Figure 13: Optical layout of the SPIRIT-III instrument (image credit: SDL, JHU/APL)

The SPIRIT-III interferometer-spectrometer focal plane consists of six detector elements (Figure ) physically separated on individual substrates with no overlapping fields of view. The detectors are similar to the radiometer detectors in composition but are significantly larger. They are arsenic-doped silicon, blocked impurity-band conductor detectors with spectral responsivity ranging from 2.5 to 28.0 µm. The focal plane arrays were provided by Rockwell International under contract to SDL.


Figure 14: SPIRIT-III interferometer-spectrometer detector layout (image credit: SDL, JHU/APL)


Figure 15: Optical layout and band separation of the SPIRIT-III radiometer(image credit: SDL, JHU/APL)


Figure 16: SPIRIT-III interferometer-spectrometer layout (image credit: SDL, JHU/APL)

FT Spectrometer Passbands

Radiometer Passbands


Passband (µm)


Passband FWHM (µm)

Active Columns
(array size)

Noise equivalent radiance for
MS1 mode (Wcm-2 sr-1)


17.2 - 28.0
2.6 - 4.9
5.8 - 8.9
4.0 - 28.0
10.6 - 13.0
2.5 - 24.0


6.03 - 10.91
4.22 - 4.36
4.24 - 4.46
11.10 - 13.24
13.50 - 15.90
18.30 - 25.00





Table 2: SPIRIT-III instrument parameters



Mirror-scan 1

Mirror-scan 2

Mirror-scan 3 (3.0º)

Scan rate

(orbit drift rate)




Scan period (one full, double scan cycle)


3.66 s

7.32 s

14.65 s

5-color scan FOR

drift position limited

0.426º x 0.99º

1.268º x 0.99º

2.953º x 0.99º

Color-set scan FOR

drift position limited

0.842º x 0.99º

1.685º x 0.99º

3.369º x 0.99º

Sampling rate





Table 3: Operational scan modes of the SPIRIT-III radiometer


Figure 17: Block diagram of the SPIRIT-III instrument (image credit: JHU/APL)

Legend to Figure 17: Heavy lines indicate paths for spacecraft control data bus. (TGA = thermogravimetric analysis, QCM = quartz crystal microbalance, BRDF = bidirectional reflectance distribution function, and OSDP = Onboard Signal and Data Processor).


UVISI (Ultraviolet/Visible Imaging and Spectrographic Imaging):

UVISI is a hyperspectral suite of nine instruments, designed by APL. The primary objective of UVISI is to collect data on celestial and atmospheric backgrounds; secondary objectives include target characterization and contamination observations in conjunction with the contamination instruments.

The UVISI sensor system consists of five SPIM (Spectrographic Imager) instruments, four imagers, and a set of instrument electronics that provide control functions and image processing capabilities (Figure 19). The imagers include two WFOV (Wide-Field-of-View) and two NFOV (Narrow-Field-of-View) sensors in the UV and VIS spectrum, respectively. Together the 5 SPIMs as well as the 4 imagers cover an overlapping spectral range from 110 nm (UV) to 900 nm (VNIR). See Figure 20 and Tables 4 and 5. 25)


Figure 18: MSX spacecraft layout showing the accommodation of the UVISI instruments (image credit: JHU/APL)

The nine instruments share a common boresight with each other, with the SPIRIT-III IR radiometer, and with the SBV instrument.

Each UVISI sensor has its own radiation-hardened 80C85-based local controller or SEU (Sensor Electronics Unit). The functions of the SEU include CCD and analog-to-digital timing control; image intensifier gain control; filter wheel, scan mirror, and slit wheel mechanism control; command and telemetry communications with the DCS (Data Control System); data collection from the FPU (Focal Plane Unit); and secondary power generation and distribution to the FPUs and mechanism controllers.

The DCS is the single interface for the operation of all nine UVISI sensors via the spacecraft command and data handling systems, and the UVISI imagers interface via the image processor. It is one of two redundant packages in the instrument. The DCS commands the sensor's operational mode and its on and off sequence. It is designed to process the sensor data to match the bandwidth of the sensors to the available spacecraft telemetry bandwidth on the 16 kbit/s, 1Mbit/s, and prime science 5 or 25 Mbit/s telemetry links. This subsystem synchronizes all sensor data collection in either the 2 or 4 Hz frame rates.

The UVISI instrument contains an image processor that selects and processes images from one of the four imaging sensors. Its purpose is to provide the spacecraft tracking processor with information needed to complete closed-loop tracking of targets of interest. From an analysis of the contents of each image, the image processor generates a list of potential targets. This list is sent to the MSX tracking processor, where its data are input to a tracking loop.


Figure 19: Block diagram of the UVISI instrument (image credit: JHU/APL)


Figure 20: Wavelength coverage (in µm) of the UVISI instrument suite (image credit: JHU/APL)

UVISI SPIMs: All five SPIMs feature an all-reflective off-axis parabolic design (external sunshade, scanning mirror, slit/filter, collimating mirror, dispersive grating, and an intensified CCD focal plane) in which selectable slits provide spectral resolutions between 0.5 nm to 4.3 nm. The SPIM image planes have programmable spectral dimensions with 68, 136, or 272 pixels and programmable spatial dimensions with 5, 10, 20, or 40 pixels. A scan mirror sweeps the slit through a second spatial dimension and generates a spectrographic image once every 5, 10, or 20 seconds.


Figure 21: Optical layout of a spectrographic imager (image credit: JHU/APL)

The SPIM CCD detector plane collects an image of the slit length in one direction and spectral information in the other (Figure 23). One spatial dimension is resolved along the slit and the other spatial dimension information is obtained by moving the FOV either 0.05º or 0.1º before recording another observation. It takes about 10 s to complete a 1º x 1º image in either 10 or 20 steps. The five-position slit/filter mechanism provides two slit sizes (1.0º x 0.10º and 1.0º x 0.05º) and various blocking filters that eliminate extraneous spectral orders and long-wavelength contaminants.

Note: The spatial resolution of the SPIMs is driven by the point-spread function in one direction (along the slit) and by the point-spread function and the mirror step size in the other direction. For the 0.05º mirror steps one can assume that it is driven by the point-spread function in both directions, and is about 0.85 mrad. The spatial resolution is diminished by using the 0.1º steps or by reducing the number of bins in the readout, by co-adding 2, 4, or 8 adjacent pixels. This is to reduce the bandwidth requirement by trading spatial resolution, spectral resolution and frame rate. The nadir resolution is 0.85 mrad x 900 km -770 m. Nadir FOV is 17 mrad (1º) x 900 km -15 km x 15 km.


Figure 22: Schematic of spectrographic imaging of UVISI SPIMs


Passband Range (nm)

Resolution Δλ (0.10º slit)

Resolution Δλ (0.05º slit)

No. of Bins



0.8 nm
1.2 nm
1.8 nm
2.8 nm
4.3 nm

0.5 nm
0.9 nm
1.5 nm
2.1 nm
2.9 nm

68, 136, or 272
68, 136, or 272
68, 136, or 272
68, 136, or 272
68, 136, or 272

Table 4: UVISI spectrographic-imager (SPIM) characteristics


Figure 23: Mounting arrangement of the UVISI sensors on MSX

The SPIM electronics (Figure 19) perform pixel-summing operations (programmable) on the 40 x 272 element SPIM focal plane array, yielding 40, 20, 10, or 5 bins in the spatial direction, and 272, 136, or 68 bins in the wavelength direction. The pixel resolution is 14 bit.

Note: The bins are formed in the SPIM electronics by co-adding 1, 2, or 4 adjacent pixels; this is done to reduce the data bandwidth requirement in cases where UVISI is not the principal instrument, or higher frame rates are needed which can be traded off against resolution. For the case of 136 and 272 bins, the bins overlap; for the case of 68 bins, the bins are noncontiguous.

With a dynamic range of 1011, UVISI has the capability to obtain data ranging from noontime surface-reflected visible radiation to observations of a diffuse UV celestial background. Objective and applications: generation of spatial/spectral maps of various objects and scenes; prime SPIM observations include the atmospheric dayglow and nightglow, aurora, stars, zodiacal light, and plume contrails. When suitably inverted, SPIM radiance measurements can reveal atmospheric properties such as species concentrations, temperatures, and altitude profiles. SPIM investigations of auroral radiances can provide estimates of fluxes and energies of precipitating particles that cause auroral emissions.

UVISI Imagers: The four imager suite consists of: external sunshade, imaging optics, filter wheel and drive motor, and intensified CCD focal planes. Three of the four imagers employ all-reflective optics. Both NFOV imagers use a Cassegrain design; the UV WFOV imager uses an off-axis three-mirror design.

The commandable filter wheel in each imager houses three bandpass filters on a neutral-density (ND) filter. In case of the visible WFOV imager, a lens is substituted for one of the filters. This lens is used to focus the near-field backscattered emissions from the xenon flashlamp as part of the MSX contamination experiment. The filter bands were chosen for specific objectives. For example, the UV-NFOV 200-230 nm filter is used to observe the NOy bands in airglow measurements. The UV-WFOV 117-127 nm filter is used for viewing Lyman-α emission from hydrogen in the geo-corona, from auroral hydrogen precipitation, or from outgassing of the SPIRIT-III cryogen. Each imager has a focal plane CCD array of 256 x 244 pixels with 12 bit resolution. The imager nadir resolutions are:

• NFOV: 0.09 mrad x 900 km ¿80 m

• WFOV: 0.82 mrad x 900 km ¿750 m

The limb resolutions can be calculated from geometry, given a tangent height.







1.28º x 1.59º

10.5º x 13.1º

1.28º x 1.59º

10.5º x 13.1º

Resolution (IFOV)

90 µrad

820 µrad

90 µrad

820 µrad

Passband (nm)





ND filter
WB1 filter
WB2 filter
WB3 filter





Table 5: UVISI imager characteristics


Figure 24: Schematic diagram of the UVISI focal plane unit

The UVISI image processing system can acquire and isolate likely targets in an image FOV and communicate their positions to the MSX flight processor, which in turn points the satellite in a “closed-loop” fashion. The term “target” may refer to either a point source such as a star, to another satellite, or to an extended source such as an auroral or cloud-top feature. The processor software performs initialization and tracking functions that include filtering, smoothing, thresholding, and centroiding. The image processor seeks targets based on an a priori target description file containing weights for numerous target features such as size, brightness, and location. The image processor then transforms target locations from UVISI pixel coordinates to S/C coordinates and passes them on the MSX flight processor, which performs Kalman filtering of other targeting inputs to select “true” targets for observation.


Figure 25: Optical diagram of the UVISI spectral imagers (SPIMs)


SBV (Space Based Visible) Sensor:

The SBV optical camera was designed and built by the MIT Lincoln Laboratory (MIT/LL) under the ACTD (Advanced Concept Technology Demonstration) program. The objective is to collect data on celestial target signatures (stars) in the VIS range and to perform above-the horizon surveillance demonstrations (cataloging of resident space objects). It uses a broadband visible wavelength detector and signal processor to automatically detect space objects, such as intercontinental ballistic missile targets and satellites, from their reflected sunlight. The goal of the SBV instrument is to provide high-performance search, detection, tracking, and data gathering on space objects to support both midcourse surveillance and space surveillance missions.

The instrument incorporates a 15 cm aperture off-axis, re-imaging, all-reflective telescope, a thermo-electrically cooled (front-illuminated) bare CCD focal plane, and electronic subsystems. The focal plane consists of four CCDs with 420 x 420 (27 µm) pixels each, Use of MIT/LL CCID7 frame-transfer visible-light CCDs, which were specifically developed for space surveillance - where CCID7 (Charge-Coupled Imaging Device 7). A primary reason for selecting the CCID7 imager is its unique combination of high charge transfer efficiency and low readout noise. 26) 27)

Spectral range

0.3 - 0.9 µm

Spatial resolution (IFOV)

60 µrad


1.4º x 6.6º

Aperture, focal ratio

15 cm, f/3

Focal Plane Array (FPA, four CCDs)

420 x 1680 pixels

Frame times

0.4 s, 0.5 s, 0.625 s, 1.0 s, 1.6 s, 3.125 s

Instrument mass, power

78 kg, 68 W

Table 6: SBV instrument parameters


Figure 26: Optical layout of the SBV telescope (image credit: MIT/LL, JHU/APL)


Figure 27: Block diagram of the CCID7 CCD imager (P1, P2, and P3 are three-phase polysilicon clocks), image credit: MIT/LL


Figure 28: SBV telescope schematic and focal plane projection (image credit: MIT/LL, JHU/APL)


Figure 29: Schematic view of the fields of view (FOVs) of the MSX instruments (image credit: JHU/APL)

Legend: The SPIRIT-III interferometer FOV is completely outside of the field of regard (FOR) of the SPIRIT-III radiometer. The two are placed in context by the UVISI wide field of view (WFOV) imagers, IUW and IVW. The UVISI spectrographic imagers (SPIMs) and the SPIRIT-III radiometer have instantaneous FOVs that are perpendicular to each other. Both the UVISI SPIMs and the SPIRIT-III radiometer use a scan mirror to achieve the desired FOR (Ref. 5).


Contamination experiments:

MSX carries in addition a suite of contamination instruments with the objective to characterize the molecular and particulate environment onboard and around the spacecraft (all APL sensors). This suite includes the following instruments: 28) 29)

NMS (Neutral Mass Spectrometer). A quadrupole RF instrument with two independently programmable filaments which can be operated at either high or low emission currents. It measures species (H2O, H2, O, hydrocarbons) ranging in mass from 1-150 amu (atomic mass units).

IMS (Ion Mass Spectrometer). A Bennett RF positive ion mass analyzer with sampling port oriented towards the ram velocity vector (+X axis) during the park mode of the MSX S/C. It measures ions (H+, O+, OH+, CH4+) ranging in mass from 1-60 amu. Gas densities as low as 106 molecules/cm3 and 103 ions/cm3 can be measured with NMS and IMS, respectively.

CQCM (Cryogenically-cooled Quartz Crystal Microbalance). Five sensors for measuring depositions on the SPIRIT-III primary mirror.

KF/KR (Krypton Flashlamp/Radiometer). KF/KR works in conjunction with the UVISI SPIM3 to measure water vapor concentrations.

TQCM (Temperature-controlled Quartz Crystal Microbalance). Four sensors are distributed about the MSX instrument section with the objective to view the mass flux arriving from a particular direction of interest.

TPS (Total Pressure Sensor). Measurement of ambient pressures ranging from 1 x 10-10 to 1 x 10-5 Torr.

XF (Xenon Flashlamp). Produces an intense VIS photon beam for detection of backscattered radiation by the UVISI wide field visible imager (IVW). The photon beam is pulsed at 1/2 s, the beam intersects the imager line-of-sight at 200 cm. The imager collects a series of frames of the backscattered emissions, from which the size and velocity of particles as small as 0.5 µm can be detected.


OSDP (Onboard Signal and Data Processor)

OSDP is a pathfinder event processor, a “third generation” signal processor, manufactured by Hughes Aircraft Company. The objective is to demonstrate real-time detection and tracking of targets in space, using long-wave infrared data from the SPIRIT-III (Spatial Infrared Imaging Telescope-III) sensor. The goal is to support a number of technology demonstrations, both on-orbit and off-line, to show that current state-of-the-art onboard signal and data processing technology can meet the requirements of the next generation of spaceborne optical sensors. Some of the key objectives of these demonstrations are: 30)

• Tracking of up to 100 objects

• Background-adaptive thresholding

• Scan-to-scan correlation of objects in track

• Acquisition and tracking of a designated object

• Jitter correction

• Object-velocity-corrected signal processing

• Birth-to-death tracking of object clusters.

OSDP (Figures 30 and 31) consists of two main components, the TDP (Time Dependent Processor) and the ODP (Object Dependent Processor). The TDP is a combination of very large scale integration (VLSI), application-specific integrated circuit (ASIC) chips and field-programmable gate arrays. Its processing is called time dependent because the TDP operates upon all of the digitized detector samples that are output from the color A and color D focal plane assemblies to detect the presence of target objects.

The ODP is implemented in Honeywell's radiation-hardened, generic, very-high-speed, spaceborne computer (GVSC), which hosts software programmed in Ada. Its processing is called object dependent because the ODP operates only upon data sets (strip reports) surrounding the target objects identified by the TDP. The OSDP hardware is fully redundant. The unit consumes 28 W, weighs 18 kg, and occupies a volume of 0.024 m3.


Figure 30: Schematic layout of the OSDP architecture (image credit: Hughes Aircraft Company)


Figure 31: Photo of the OSDP instrument (image credit: Hughes Aircraft Company)

The TDP parameters are fed back from the object-dependent processor (ODP) to improve the signal processing performance. The TDP performs gamma (spike) event circumvention, detector responsivity correction, background subtraction, time delay and integration (TDI), matched filtering, and thresholding. The gamma pulses (spike events) are caused by such phenomena as cosmic ray and proton collisions with the FPA. The process of TDI involves boosting object signal-to-noise ratio (SNR) by coherently adding the outputs of detectors aligned in the scan direction of the sensor. Using the known scan rate, and assuming that object inertial motion is negligible, the data from each detector are stored in a TDI chain for the time delay required to cause each detector to observe the same spatial position.

The ODP is capable of initiating and maintaining tracks on objects in the sensor’s field of view. The availability of these tracks gives rise to a number of spin-off functional enhancements, such as (1) feedback of object velocity estimates to correct for velocity distortions in array correlation and matched filtering; (2) windowing, a convenient and computationally efficient form of scan-to-scan correlation; and (3) designated object tracking, where the MSX tracking processor provides (in the broadcast message) a start-up state vector for an object to be tracked.

The TDP-ODP interface and the manner in which the TDP passes thresholded data to the ODP are key to the OSDP design. The TDP hands off matched filter threshold exceedances to the ODP in groups called strip reports, constructed in a manner that maintains the integrity of clumps extending across many pixels. This function is called clump processing.


MSX ground system:

The MSX spacecraft is operated by APL; its S-band is also monitored by AFSCN (Air Force Satellite Control Network). APL also receives the prime science data and provides processing and archiving for all data.

Communications associated with ground operations encompassed a variety of networks that connected the spacecraft to the PCC (Payload Control Center) at a specific test site; to the Command Center at the Air Materiel Command's Detachment 2/Space and Missile Center, Onizuka AFB, CA; or to the MCC (Mission Control Center) at APL. 31)


Figure 32: MSX communications networks for ground operations (image credit: JHU/APL)


Figure 33: Organizational interfaces of the MSX Flight Operations Team (image credit: JHU/APL) 32)

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The information compiled and edited in this article was provided by Herbert J. Kramer from his documentation of: ”Observation of the Earth and Its Environment: Survey of Missions and Sensors” (Springer Verlag) as well as many other sources after the publication of the 4th edition in 2002. - Comments and corrections to this article are always welcome for further updates.