Minimize DSX

DSX (Demonstration and Science Experiments) in MEO

DSX is a small spacecraft mission of the U.S. AFRL (Air Force Research Laboratory), Albuquerque, NM, USA. The objective is to research technologies needed to significantly advance Department of Defense (DoD) capabilities to operate spacecraft in the harsh radiation environment of MEO (Medium-Earth Orbits). The ability to operate effectively in the MEO environment significantly increases the DoD‘s capability to field space systems that provide persistent global space surveillance and reconnaissance, high-speed satellite-based communication, lower-cost GPS navigation, and protection from space weather and environmental effects on a responsive satellite platform. 1) 2) 3) 4) 5)

The three DSX physics-based research/experiment areas are:

1) WPIx (Wave Particle Interaction Experiment): Researching the physics of VLF (Very-Low-Frequency) electromagnetic wave transmissions through the ionosphere and in the magnetosphere and characterizing the feasibility of natural and man-made VLF waves to reduce and precipitate space radiation.

2) SWx (Space Weather Experiment): Characterizing, mapping, and modeling the space radiation environment in MEO, an orbital regime attractive for future DoD, Civil, and Commercial missions.

3) SFx (Space Environmental Effects Experiment): Researching and characterizing the MEO space weather effects on spacecraft electronics and materials.

The DSX mission will provide an important database from which future designers of space missions intended to operate in the MEO environment can rely. These are:

• New climatology models for satellite design in MEO will be developed as a result of DSX

• Key wave-particle interaction components of global radiation belt space situational awareness and forecast models will be developed and validated from the DSX data

• New methods for ruggedization of electronics for MEO will be tested.

Beyond the inherent value of the scientific experiments, DSX will also serve as a pathfinder for future DoD rapid response missions, by developing and demonstrating both technologies and processes that enable low-cost and rapid integration of one-of-a-kind R&D satellites and operational systems. DSX addresses the rapid integration problem through the use of an innovative network based infrastructure, the Planning System Inc. (PSI) Network Data Acquisition System (NDAS), for implementation of bus and payload box connectivity, as well as distributed remote sensor payloads on large deployable structures. 6)


The mission was originally conceived by AFRL researchers in 2003 to conduct physics-based experiments, and was selected as an AFRL mission in 2004. With primary funding from AFRL, DSX enjoys additional support from DARPA (Defense Advanced Research Projects Agency) and NASA. Note: Prior to 2005 DSX was known as `Deployable Structures Experiment' mission.

The MEO region is a relatively unused and unexplored domain of spaceflight with regard to the radiation environment requiring investigations for model development (energetic particle distributions, VLF injection, propagation, wave particle interaction processes, etc.), and to study the effects of exposure on the spacecraft electronics and materials. 7) 8) 9) 10) 11) 12)

The MEO region is of great interest to DoD because it significantly increases the potential to field future space systems that provide persistent global targeting-grade space surveillance, high-speed, satellite-based communication, lower-cost global positioning system navigation, and protection from space weather on a responsive satellite platform. The MEO orbit with its longer spacecraft contact periods (than in LEO) is also being studied and considered by NASA/NOAA and other agencies for future integrated concepts of environmental (weather) satellite missions. The uninterrupted Earth coverage requirements of many missions are making the MEO range attractive for a number of observation services.

To address the space access aspect of the rapid-response problem, DSX utilizes an ESPA (EELV Secondary Payload Adapter) capability as a platform for highly-capable small and medium free-flying satellites (or ESPASats) that have plentiful and relatively inexpensive launch opportunities on EELV as secondary payloads. Note: The first flight of ESPA took place on the STP-1 mission of DoD with a launch on March 9, 2007 (6 vehicle launch: Orbital Express (OE), MidSTAR-1, STPSat-1, NPSat-1, CFESat, and FalconSat-3). 13) 14)



Figure 1: Artist's rendition of the deployed DSX spacecraft attached to the ESPA ring (image credit: SNC, AFRL)


The DSX spacecraft is flying on the HSB (Host Spacecraft Bus), provided by SNC (Sierra Nevada Corporation), of Louisville, CO [note: in 2008, SNC acquired MSI (MicroSat Systems, Inc.) of Littleton, CO which was awarded a contract to build DSX in 2007]. The DSX flexible design structure (a derivative of MSI's modular TacSat-2 bus) maximizes the compatibility with numerous launch vehicles. 15) 16)

The HSB consists of the AM (Avionics Module), the PM (Payload Module) structure and the harness attached to the ESPA (EELV Secondary Payload Adapter) ring connecting the two modules together electrically. The AM is essentially the spacecraft bus with all the requisite subsystems (communications, power, thermal, ADCS (Attitude Determination and Control Subsystem), C&DH (Command and Data Handling), etc). It also houses the HEPS instrument, the ECS (Experimental Computer System), and the radiometers and photometers. The PM contains all the DSX specific components, except those just mentioned, along with the HSB Magnetometer and IMU (Inertial Measurement Unit).

The modules are mechanically bolted to the ESPA ring via a 39 cm bolt circle on two opposing ports of the DSX four port ESPA ring. DSX will fly the ESPA ring with both AM and PM attached at all times; the AM and PM will not separate. The AM and PM are each mechanically connected to the ESPA ring via a conical adapter, part of SNC’s module design, a SoftRide system developed by CSA Engineering for shock and vibration suppression and attenuation. The DSX in its stowed configuration is shown in Figure 2.

DSX has no propulsion, GPS unit, or star trackers. DSX does have an IMU, three sun sensors, and a magnetometer (in addition to the VMAG unit) for ADCS. The spacecraft is 3-axis stabilized.


Figure 2: Illustration of the DSX stowed spacecraft configuration (image credit: AFRL, SNC)

EPS (Electrical Power Subsystem): EPS consists of two Li-ion batteries, each of 30 Ah capacity, designed and manufactured by Yardney, Inc., a fixed, single wing, three panel solar array (with composite facesheets and Al honeycomb core) with triple junction GaAs cell technology, an efficiency of 28.5% at BOL (Beginning of Life), and the associated switches, diodes, and wiring.

The thermal subsystem consists of closed-loop heaters, white paint (AZ-93 or Alion Z-93) on the Y-boom canisters, and multi-layer insulation (MLI, 12 layers – 1 black Kapton, 10 Mylar, 1 AL Kapton). White coatings (Z-93C55) and black coatings (MH55-ICP) coat the structural surfaces of the AM and PM (black on the outside, white on the inside with an emissivity > 0.8). The ESPA ring is an all aluminum structure that serves as a large heat sink for the DSX spacecraft. DSX also has 63 temperature sensors distributed around the spacecraft with 47 of them as AD590s and 16 of them as PRT (Platinum Resistance Thermometers). Predicted temperatures for the DSX range from a max cold of – 116ºC for the Y and Z antennas to a max hot of 45ºC for the CXS-810C L-3 SGLS (Space Ground Link Subsystem) radio. The duty cycle for the heaters is no greater than 72%.

The structures and mechanisms subsystem of DSX consists of the AM structure, the PM structure, the ESPA ring, solar array panels, and the associated bulkheads and brackets. The margins of safety inherent in the designs are 8.5 g’s in two directions simultaneously which worst case is 12 g’s in each of three directions individually. The ultimate factor of safety is 1.4. The ESPA ring was qualified for flight for the STP-1 Mission and the qualification test data correlated to FE (Finite Element) models. DSX’s ESPA ring has a few modifications (4 port boss ports vs. the original 6 flange ports), so the DSX PMO had CSA engineering perform a Qualification by FE Analysis for the new design that showed a positive margin with a safety factor of 2.0 for the same testing conditions as STP-1’s qualification tests and the basic DSX design. The AM and PM flight structures were workmanship tested by subjecting them to sine burst levels up to 14.4 g (1.2 safety factor) with mass simulators in place of payloads.

C&DH (Command and Data Handling): C&DH and the communications subsystem consists of the SGLS CXS-810C radio, built by L-3 Communications Inc., the IAS (Integrated Avionics System), built by SEAKR Engineering Inc., the AM, PM, and the ESPA harnesses, built by Cicon Engineering Inc., the ECS (Experimental Computer System), built by Planning Systems Inc. (PSI), space-qualified RF switches, couplers, and diplexers, six helical antennas, and two omni patch antennas.

RF communications: The CXS-810C is the SGLS transponder radio with a BAE provided RAD-750 processor, providing SGLS uplink at 2 kbit/s and SGLS downlink up to 2 Mbit/s with ranging capability. The uplink and downlink signals will be encrypted. With an orbit of 6,000 km x 12,000 km, the uplink margin will be 23.2 dB worst case at 2 kbit/s and the worst case downlink margin will be 5.0 dB for safe mode at 500 bit/s and 5.1 dB in nominal operations at 1 Mbit/s. The worst case ranging margin will be 8.1 dB. The L-band command uplink operates at 2 kbit/s and can be received by either of the two RHCP (Right Hand Circular Polarization) patch antennas for approximately 4π steradian coverage. The six helical antennas are high data rate downlink only and are spaced around the AM structure and each has a 64º FOV (Field of View).

ADCS consists of the VMAG (Vector Fluxgate Magnetometer), an IMU ( Northrup Grumman) running at 400 Hz, a Billingsley magnetometer (also known as the HSB magnetometer), three Adcole sun sensors working together as the SSA (Sun Sensor Assembly), three Honeywell 25 Nms reaction wheels, and three 400 Am2 magnetic torque rods (ZARM). The magnetometers, sun sensors, and IMU provide the sensor data to the system for attitude control and command generation. ASI (Advanced Solutions Inc.) designed the ADCS subsystem and wrote the ADCS software. The reaction wheels and torque rods are controlled by this software to orient the spacecraft as required per the CONOPS (Concept of Operations).


Figure 3: Alternate view of the DSX spacecraft and its components (image credit: SNC)


Spacecraft Bus
(Avionics Module=AM)

DSX Spacecraft
Payload Module =PM)


Mission design life

12 months

12 months


Launch vehicle compatibility

Minotaur, Falcon X



Orbit capability


6,000 km to 12,000 km any inclination


Spacecraft control

3-axis stabilized

3-axis stabilized


Payload mass (max)


170 kg

Additional mass available on other ESPA ports

Bus mass

180 kg



Spacecraft mass (Total, AM, PM, ESPA Ring)


600 kg, 180 kg, 180 kg, 240 kg

Includes 240 kg ESPA ring

Stowed dimensions

1.1 m x 0.8 m x 0.8 m

3.6 m x 2 m x 1.1 m

Deployed DSX: 81 m x 17 m x 8 m

Payload power (orbital average, peak)

350 W, 1.1 kW

350 W, 1.1 kW


Array power (average, peak)

650 W, 790 W

650 W, 790 W

For 6,000 km x 12,000 km orbit

Battery capacity

60 Ah

60 Ah

Li-ion battery

S/C pointing (knowledge, control)

1.0º, 1.0º

1.0º, 1.0º


Slew rate

0.25 to 1.0º/s

0.25 to 1.0º/s


RF communications (uplink, downlink)

2 kbit/s, 2 Mbit/s

2 kbit/s, 2 Mbit/s

S-band encrypted

Command & Data Handling

240 MIPS, 128 MB, 128 MB

240 MIPS, 128 MB, 128 MB


Table 1: Overview of DSX spacecraft parameters


Launch: A launch of DSX is scheduled for October 2012. DSX has a launch manifest with the DMSP Flight-19 (DMSP F-19) primary payload. The launch site is VAFB (Vandenberg Air Force Base) and the launch vehicle is the Atlas-V-401.

Orbit: The DSX nominal elliptical MEO (Medium Earth Orbit) has a perigee of 6000 km and an apogee of 12000 km, inclination = 120º (retrograde orbit) period ~ 5.3 hours. An average of 85 minutes contact time per orbit is provided.

This orbit will permit DSX to fly through the outer region of the inner Van Allen radiation belt, the slot region, and the inner region of the outer Van Allen radiation belt. The total ionizing dose (TID) for the mission is estimated to be equivalent to 10 krad/year behind 8 mm of aluminum shielding. The implication of this is that all DSX equipment must be designed to survive the radiation environment, through the use of shielding, selection of radiation and SEU (Single Event Upset) tolerant or hardened components, or a combination of these approaches. 


Sensor complement: (WPIx, SWx, SFx)

DSX has thirteen payloads arranged into three main groups of experiments: WPIx, SWx, and SFx.

WPIx (Wave Particle Interaction Experiment):

The WPIx experiment package is designed to measure critical parameters for space and ground-based wave particle interactions: efficiency of VLF injection, propagation, and efficacy. The mission objectives for WPIx are to:

1) Determine VLF injection efficiency from a space-based antenna.

2) Map VLF fields from geophysical and man-made sources

3) Measure precipitating particles scattered by VLF from geophysical, ground-based transmitter, and DSX VLF transmitter sources.

The WPIx experiment consists of the following payloads: BBR (Broadband Receiver), TATUs (Transmitter and Tuning Units), TCU (Transmitter Control Unit) including the NBR (Narrowband Receiver), TASC (Tri-Axial Search Coil), VMAG (Vector Magnetometer), ECS (Experiment Computer System), LCI (Loss Cone Imager), Y antenna, and Z antenna.

VLF (Very Low Frequency Receiver):

The VLF broadband receiver is comprised of the BBR, TASC, Y antenna, and Z antenna. The VLF receiver has three search coil magnetometers (3 B [magnetic] components via TASC), two linear, orthogonal dipole antennas with 2 E (electric) components. The frequency range is 100 – 50 kHz and the sensitivity is 1.0 e-16 V2 m-2 Hz-1 (E) and 1.0 x e-11 nT2 Hz-1 (B). The VLF receiver is built by Stanford University, NASA/GSFC, Lockheed Martin, and ATK.

The VLF narrowband receiver is comprised of the NBR and the Y antenna, covering the band from 3 kHz to 750 kHz. The VLF transmitter operates in two modes, high power (i.e. Whistler mode waves) at 3 – 50 kHz at up to 500 W (900 W at end of life), and low power mode (i.e. Boomerang mode waves) at 50 – 750 kHz at 1 W, and the local electron density. The transmitter is built by the University of Massachusetts Lowell (UML), SwRI (Southwest Research Institute), and ATK.

The LCI (Loss Cone Imager) features a HST (High Sensitivity Telescope) which will measure 100 – 500 keV electrons with 0.1 cm2 str geometric factor with 6.5º of loss cone. The LCI also features a FSH (Fixed Sensor Head) with 130º x 10º of pitch angle distribution for 50 – 700 keV electrons every 167 ms. Boston University is building the LCI instrument. 17) 18)
Finally, the VMAG (Vector Fluxgate Magnetometer) instrument is capable of 0 – 8 Hz three axis measurements at ± 0.1 nT accuracy of magnetic field line measurements. The VMAG is built by the University of California Los Angeles (UCLA). 19)

MAG is being used to determine the direction of the magnetic field to better than 1º at all points in the orbit. This performance will allow the mapping of locally measured particle distributions to global distributions. VMAG sensors will measure the DC magnetic field over a 100-10000 nT range with ±0.1 nT accuracy at 20 Hz. VMAG will operate continuously to provide DC magnetic field and ULF wave environment data required by the Space Weather Experiment MEO space particle modeling experiment. Hence, it supports the experiment objectives of both the WPIx and the SWx.

The VMAG electronics will be mounted within the payload module and the fluxgate sensor will be deployed on the tip of the -Z boom, opposite TASC.


Figure 4: The VMAG fluxgate sensor (left) and electronics board (right), image credit: UCLA


Figure 5: Conceptual illustration of the VLF wave particle interaction processes (image credit: AFRL)

VLF transmitter:

The VLF transmitter on DSX provides the capability to conduct active experiments to quantify spaceborne VLF wave-injection efficiency and determine the details of the wave-particle interactions. An electron loss cone detector on DSX allows direct correlation of changes in energetic particle distributions with injected wave power. The objective is to develop and verify accurate models of the VLF injection, propagation, and wave particle interaction processes with DSX data.

The VLF antennas are aligned normal to the local magnetic field line when spacecraft is within ±20º of equatorial (i.e., when VLF propagation experiments and measurements will be performed). Pointing requirements are fairly relaxed with a 2º control and a 1º knowledge requirements deemed to be sufficient.

The design of the transmitter was driven by objectives 1 and 2, but also influenced by the desire to maximize space flight heritage to minimize risk, and budget constraints. The DSX Y antenna is 80 m in length (tip-to-tip) and functions as a VLF receive and transmit antenna. The DSX Z antenna is 16 m in length (tip-to-tip) and functions as a VLF receive antenna in a cross-dipole configuration with the Y antenna. The TASC and VMAG instruments are placed at opposite tips of the Z antenna to separate them from the rest of the DSX instruments and their electrical and mechanical “noise” which would interfere with their operation as VMAG measures the local magnetic field and TASC measures the local electric field.

The antenna lengths were based on the general opinion among space physicists that the longer and wider the antenna the better due to minimizing impedance when generating VLF waves in a magnetized plasma.

The transmitter voltage design is based on NASA’s IMAGE (Imager for Magnetopause-to-Aurora Global Exploration) RPI (Radio Plasma Imager) instrument built by UML that operated at 3 kV and was optimized for > 50 kHz. The DSX design optimizes the transmitter impedance dependant on frequency, antenna length, and diameter. DSX is flying the first ever VLF “dynamic tuning” technology to adjust circuit parameters in real time. The voltages are limited to < 10 kV due to critical component limits. The DSX system is nominally designed for 5 kV with the capability to go to 10 kV at the end of life (EOL).


SWx (Space Weather Experiment):

The objective of the Space Weather Experiments (SWx) are to measure, map, and characterize the space weather environment in the slot region for electrons, protons, ions, and plasma and then develop and validate models of this region. The SWx payloads are as follows: 20)

CEASE (Compact Environmental Anomaly Sensor)developed by Amptek Inc.. CEASE will measure radiation dose, dose rate, surface dielectric charging, deep dielectric charging, and single event effects. The device was flown on TSX-5 (launch June 7, 2000), on STRV-1C (launch Nov. 16, 2000), and on DSP-21 [(Defense Support Program-21) satellite, launch Aug. 6, 2001]. CEASE consists of two dosimeters, two particle telescopes, and a SEE (Single Event Effect) detector. The device has the capability to monitor a broad range of space hazards from surface damage and charging (keV electrons) to SEE events resulting from >100 MeV cosmic rays and solar protons. The angular FOV for CEASE is relatively large and will not resolve pitch-angle distribution. CEASE will be mounted on an exterior panel of the payload module. One change for DSX is that CEASE will capture and downlink the full dose spectra from each dosimeter, whereas prior versions only captured six reduced data points (two for low LET data and four for high LET data). 21)

HEPS (High Energy Proton Spectrometer), developed by Amptek Inc. The objective is to measure protons with energies between 15 and 440 MeV and electrons with energies between 1.5 and 20 MeV. These high energy particles are responsible for microelectronics damage, displacement and total dose damage, SEEs, and deep dielectric charging. HEPS will be mounted on the AM battery enable bracket.


Figure 6: HEPS electronics (left) and sensor (right), image credit: AFRL

LIPS (Low Energy Imaging Proton Spectrometer), developed by PSI. The objective is to measure the ring current particles that are important in the energy flow processes in the magnetosphere. This particle population plays an important role in spacecraft charging and surface damage due to sputtering. The instrument uses specially designed combinations of scintillator materials to detect electrons and protons with energies between 30 keV and 2 MeV. Eight 10º x 8º apertures will provide pitch angle resolution. LIPS will be mounted on an exterior panel of the payload module.

LESSA (Low Energy Electrostatic Analyzer), developed by Amptek Inc. The objective is to measure energy fluxes and energy spectra for low-energy electrons and protons (100 eV to 50 keV). These low energy particles are responsible for surface electric charging and damage to thin films such as thin-film photovoltaics, conventional solar cell cover glasses, and coatings. LEESA will be mounted on an exterior panel of the avionics module.

HIPS (High Energy Imaging Proton Spectrometer), developed by PSI (Physical Sciences Inc.). The objective is to measure electrons with energies between 1-10 MeV and protons with energies between 30 and 300 MeV. These high-energy particles are responsible for microelectronics damage, displacement and total dose damage, SEEs, and deep dielectric charging. HIPS will be mounted on an exterior panel of the payload module.


Figure 7: Schematic view of the HIPS device (image credit: PSI)


Figure 8: Space weather energy coverage of DSX (image credit: AFRL)

DSX will be flying the most comprehensive particle energy coverage ever in MEO. The benefits of the SWx experiment are that electron and proton detectors measure both the spectral content and angle-of-arrival of both species over broad energy ranges. Also, the DSX sensor suite will help correct deficiencies in the current standard radiation-belt models of the inner magnetosphere by providing:


- Spectrally resolved, uncontaminated measurements of high energy protons (10-400 MeV) and electrons (1-30 MeV).

- Accurate mid-to-low energy (< 1000 keV) measurements of the energetic particle and plasma environment.

- Transformation of angle-of-arrival measurements into estimates of the flux distribution with respect to the local pitch-angle (via on-board magnetometers).


SFx (Space Environmental Effects Experiments):

SFx consist of the NASA Space Environment Testbeds-1 (SET-1) and the Air Force Research Laboratory; Propulsion Directorate (AFRL/RZ) produced radiometers and photometers. The objectives of SET-1 are to improve engineering approaches to accommodate and/or mitigate the effect of solar variability on spacecraft design and operations, reduce risk for new technologies infused into future space missions, and provide a standard mechanical, electrical, and thermal interface for a collection of small flight investigations. 22) 23)

The SET-1 payload is comprised of two units, CEM (Correlative Environment Monitor) and CCA (Central Carrier Assembly). The carrier provides a single interface for power and data between the DSX spacecraft and the SET-1 microelectronic investigations (inside the CCA) and CEM.

The CCA houses five sub-experiments:

• CREDANCE (Cosmic Radiation Environment Dosimetry and Charging Experiment)

• DIME-1 (Dosimetry Inter-comparison and Miniaturization Experiment Board #1)

• DIME-2 (Dosimetry Inter-comparison and Miniaturization Experiment Board #2)

• ELDRS (Linear Enhanced Low Dose Rate Sensitivity)

• COTS-2 (Commercial Off the Shelf Technology)

The SET-1 project is part of the LWS (Living With a Star) program managed at NASA/GSFC. The LWS program and the SET -1 project are sponsored by the Science Mission Directorate of NASA/HQ.

The objective of the photometers and radiometers (Figure 10) is to measure changes in optical transmission, thermal absorption and emission due to the MEO radiation environment. Coupon level testing of specific developmental coatings intended for thin-film photovoltaics will be performed using these sensors. The radiometers will detect changes in thermal control paint coefficients of heat gain / loss based on exposure and the photometers will measure the effects of erosion from quartz windows and re-deposition of material onto nearby optics.


Figure 9: Accommodation of the SET-1 instrumentation on the payload module (image credit: NASA)


Figure 10: Photos of the photometer (left) and of the radiometer (right), image credit: NASA, AFRL


Figure 11: Equipment packaging on the payload module, (image credit: AFRL)


Figure 12: Equipment packaging on avionics module (image credit: AFRL)


Technology introduction:

Modular spacecraft integration concept: DSX maximizes the benefits of modular integration via an architecture which requires that all devices adhere to a standard interface for control and communications (electrical, power, command & communications protocol). This is enabled by the use of dual-redundant network interface cards in the ECS (Experiment Computer System) chassis that provide for command and data interface between the flight control computer and all payloads. Rather than a single onboard computer responsible for handling all spacecraft functions, DSX unburdens the main spacecraft computer from payload-related operations by providing a dedicated payload-interface computer: the ECS. The highly generalized ECS design, developed by QNA-PSI (Qinetiq-North America-Planning Systems, Inc.), Melbourne, FLA, is capable to handle a wide range of payload types and classes simultaneously, with low data latency. QNA-PCI serves as principal investigator for the attitude control experiment as well as providing the ECS. 24)


Figure 13: Overview of the deployed DSX spacecraft (image credit: AFRL)

In general, the role of the ECS on DSX is required to provide the following functions:

- Isolate the interface complexity associated with 10 distinct payloads from the bus computer both in terms of hardware and command/telemetry

- Provide storage of large amounts (up to 10 GByte) of experiment data and manage the downlink of this data

- Offload the bus computer from intensive computational tasks associated with the ACE (Adaptive Controls Experiment) and payload data processing.

Key operational environmental requirements for the ECS include:

- 1-year mission life

- Tolerance of 100 kRad total ionizing dose in MEO environment

- Emissions and susceptibility per tailored MIL-STD461E

- Operational temperature range of -43ºC to 41ºC.

The ECS is a processor-based command and data handling system consisting of four separate 6U cards on a cPCI backplane. It is capable of 1500 MIPS, while providing 10 GByte of volatile mass storage and 12 distinct serial interfaces. It’s key physical attributes are:

- Total mass: 8 kg

- Peak power: 66W (allocated), 34W (measured @ 30ºC) at 28V.

- Dimensions: 27.5 cm x 20 cm x 17.5 cm.


Figure 14: Photo of the ECS flight unit (image credit: QNA-PSI)


Figure 15: Block diagram of the ECS (image credit: QNA-PSI)

As shown in Figure 15, the four cards within the ECS enclosure consist of a power supply card for conditioning raw 28 V spacecraft power to conditioned supply voltages for ECS components, a single board computer, a serial interface card, and a mass storage card.

The ECS employs the SCS750 single board computer from Maxwell Technologies, running the VxWorks real-time operating system. The SCS750 uses three PowerPC 750FX processors running in a triple-modular redundant configuration to provide 1500 MIPS in the harsh DSX environment with radiation susceptibility equivalent to only 1 uncorrected processor upset every 128 years.

The Network Interface Card (NIC) was purpose built for the ECS application by QNA-PSI. The NIC provides 9 bidirectional and one unidirectional serial ports. All ports use an RS-422 hardware layer. The bidirectional ports use a standard UART (Universal Asynchronous Receiver/Transmitter) protocol with software programmable baud rates ranging from 9600 to 115,200 baud. The unidirectional port uses a clock, data, and enable protocol, with the clock provided by the associated payload to receive data at 40 kbit/s.

The PIC (Payload Interface Card) was designed and built for QNA-PSI by SEAKR Engineering Inc. who also built the power supply card, cPCI backplane, and enclosure. The PIC provides 10 GByte of SDRAM mass storage, available to the processor via dual DMA (Direct Memory Access) channels, as well as an LVDS serial data recording port running at 2 Mbit/s. 25)

The PIC also provides primary and redundant bidirectional RS-422 serial interfaces with the host spacecraft as well as primary and redundant pulse input channels used for synchronizing the ECS timebase with the spacecraft.


CONOPS (Concept of Operations):

There are six separate and distinct mission CONOPS phases: 1) the launch and separation, 2) bus initialization, 3) experiment boom deployment, 4) post-deploy comprehensive checkout, 5) on-orbit experiment operations, and 6) end-of-life (EOL). The potential for residual operations before EOL would be defined as a seventh phase.

Normal operations will be constrained by spacecraft power, data bandwidth and scheduling contacts through the AFSCN (Air Force Satellite Control Network). Therefore payloads will have to be duty cycled. SWx payloads are low impact on resources, basically single modes of operation, and can be turned on early for the duration of the mission with little intervention required. SFx and specifically the SET-1 instrument will have two operating modes that are dependent on solar activity (for S3 storms as detected by CREDANCE) and position in the orbit (triggered by orbit propagator indications of the ascending nodes). Resource demands for SFx are similar to SWx. WPIx is the most demanding on the DSX spacecraft resources, consuming the greatest power (> 650 W peak) and generating the most data (up to 5 GB per orbit).

High power VLF transmission is limited to ±20º latitude about the equatorial plane and to less than 30 minutes. The ACE (Adaptive Controls Experiment) is a secondary experiment and, as its name implies, involves an adaptive controls experiment using the HSB attitude control system interacting with the deployed flexible structures. The only resource requirement for ACE is downlink of data generated from ACE.


DSX ground system:

The DSX ground segment consists of three major elements: mission control, global telemetry, tracking, and control (TT&C) network, and the remote experiment sites. The DSX mission will be operated by the MOC/DC (Mission Operations and Data Center) at AFRL. The the MOC/DC is collocated with RSC [RDT&E (Research, Development, Test and Evaluation) Support Complex]. Global data reception/communication is provided by AFSCN (Air Force Satellite Control Network). From the MOC, the flight operations team interfaces with the users, payload and spacecraft equipment manufacturers, remote experiment support sites (optical and radar imaging ranges and VLF ground installations), and RSC operations personnel required to operate the mission. The RSC provides connectivity to AFSCN.

Commands will be sent to the RSC for uplink via the SGLS (Space-Ground Link System) communication protocol to the DSX satellite. Downlinked information from the space vehicle state-of-health will be validated from the real-time operations system located at the RSC. All space-ground links will be accommodated by AFSCN. The DC will collect all DSX data products, conduct analyses, archive data products, and distribute the data products to PIs (Principal Investigators) and users at AFRL and the NASA/POCC (Payload Operations Control Center).


Figure 16: Overview of DSX ground system (image credit: AFRL)


Figure 17: DSX RF emissions flow diagram (image credit: AFRL)

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6) Dan Cohen, Joseph Wieber, Jan King, Shane Kemper, Shawn Stephens, Larry Davis, Gregory Spanjers, James Winter, Aaron Adler, Shaun Easley, Martin Tolliver, Jason Guarnieri, “The SSTE-4: DSX Flight Experiment: Design of a Low-Cost, R&D Space Mission with Responsive Enabling Technologies,” Responsive Space Conference 2005, Los Angeles, CA, USA, April 25-28, 2005, paper RS3-2005-3004, URL:

7) N. G. Heinsohn, T. Girard, D. Smith, J. Stuart, A. Adler, J. Schoenberg, M. Scherbarth, E. Klaiber, “AFRL's Demonstration and Science Experiments (DSX) Program - Quest for the Common Micro Satellite Bus,” Proceedings of the 21st Annual Conference on Small Satellites, Logan, UT, USA, Aug. 13-16, 2007, SSC07-II-4

8) D. Cohen, J. Schoenberg, G. Ginet, B. Dichter, M. Xapsos, A. Adler, M. Scherbarth, D. Smith, “The Demonstration and Science Experiments (DSX): A Fundamental Science Research Mission Advancing Technologies that enable MEO Spaceflight,” Proceedings of the 4S Symposium: `Small Satellite Systems and Services,' Chia Laguna Sardinia, Italy, Sept. 25-29, 2006, ESA SP-618

9) Greg Spanjers, James Winter, Dan Cohen, Aaron Adler, Jason Guarnieri, Martin Tolliver, Greg Ginet, Bronek Dichter, Jeff Summers, “The AFRL Demonstration and Science Experiments (DSX) for DoD Space Capability in the MEO,” Proceedings of the 2006 IEEE/AIAA Aerospace Conference, Big Sky, MT, USA, March 4-11, 2006

10) Dan Cohen, Ggregory Spanjers, James Winter, Gregory Ginet, Bronik Dichter, Aaron Adler, Martin Tolliver, Jjason Guarnieri, “Design and Systems Engineering of AFRL's Demonstration and Science Experiments,” GATech Space Systems Engineering Conference, Nov. 8-10, 2005, Atlanta, GA, USA, URL:

11) J. Winter, G. Spanjers, D. Cohen, A. Adler, S. Kemper, S. Easley, K. Denoyer, M. Tolliver, L. Davis, J. Guarnieri, R. Glover, “Deployable Structures Mission in Medium Earth Orbit.,” Proceedings of the IEEE Aerospace Conference, Big Sky, MT, USA, March 5-12, 2005

12) Jon Schoenberg, Gregory Ginet , Bronislaw Dichter, Michael Xapsos, Aaron Adler, Mark Scherbarth, Durand Smith, “The Demonstration and Science Experiments (DSX): A Fundamental Science Research Mission Advancing Technologies that enable MEO Spaceflight,” 2006, URL:

13) “ESPA: The EELV Secondary Payload Adapter,” URL:


15) “Demonstration and Science Experiments,” SNC, URL:

16) “Flight Proven SN-200,” SNC brochure, URL: URL:

17) James D. Sullivan, “DSX loss cone imager differential response functions,” Proceedings of SPIE, Vol. 7438, 'Solar Physics and Space Weather Instrumentation III,'. Ed. Silvano Fineschi & Judy A. Fennelly,. San Diego, CA, USA: SPIE, August 4, 2009

18) David L. Voss, Avi Gunda, Doug Carssow, Theodore Fritz, Anton Mavretic, James Sullivan, “Overview of the loss cone imager fixed sensor head instrument,” Proceedings of SPIE, Vol. 7438, 'Solar Physics and Space Weather Instrumentation III,'. Ed. Silvano Fineschi & Judy A. Fennelly,. San Diego, CA, USA: SPIE, August 4, 2009

19) Mark B. Moldwin, “Vector Fluxgate Magnetometer (VMAG) Development for DSX,” UCLA, Final Report, June 3, 2010, URL:

20) J. A. Fennelly, “Demonstration and science experiment (DSX) space weather experiment (SWx),” Proceedings of SPIE, Vol. 7438, 'Solar Physics and Space Weather Instrumentation III,'. Ed. Silvano Fineschi & Judy A. Fennelly,. San Diego, CA, USA: SPIE, August 4, 2009, URL:


22) “SET Payload,” NASA, URL:

23) Mike Xapsos, Reginald Eason, “DSX Mission/SET-1,” URL:

24) James A. King, Robert L. Gillis, Scott W. Greeley, Lawrence D. Davis, Jon Schoenberg, Mark Scherbarth, “A PCI-Based, Multiple-Payload Processing System for the DSX Flight Experiment,” 6th Responsive Space Conference, April 28–May 1, 2008, Los Angeles, CA, USA, paper: AIAA-RS6-2008-3007, URL of paper:, URL of presentation:

25) Ian Troxel, Matthew Fehringer, Michael Chenoweth, Paul Murray, “Achieving High Performance Computing and Application Flexibility within the Spacecraft Payload,” MAPLD (Military and Aerospace Programmable Logic Devices) Conference, Annapolis, MD, Sept. 16-18, 2008, URL:

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.