Minimize SENSE

SENSE (Space Environmental NanoSatellite Experiment)

Overview   Spacecraft   Launch   Mission Status   Sensor Complement   Ground Segment   References

SENSE is SMC's (Space & Missile Systems Center) premier rapid development effort which will demonstrate the capability of nanosatellites to perform space missions in an affordable and resilient manner. The goal of the USAF (U. S. Air Force) SMC/XR program is to deliver two complete spacecrafts in an 18 month period and to demonstrate the feasibility of using nanosatellites to provide future operational space weather data. The Weather Directorate (SMC/WM) is the primary sponsor of the SENSE mission. 1) 2) 3)

A fundamental objective of SENSE is to present DoD decision makers with an understanding of the benefits and costs of integrating CubeSats into an operational architecture. SENSE is a pathfinder demonstration that will help to establish how this class of vehicle—one radically different in scale from what SMC typically procures—can be acquired, launched and operated to support future SMC missions.

The three primary objectives for the SENSE mission are: 4) 5) 6)

1) to develop best practices for operational CubeSat/NanoSat procurement, development, test, and operations

2) to mature CubeSat bus and sensor component technology readiness levels

3) to demonstrate the operational utility of CubeSat measurements by flowing validated, low-latency data into operational space weather models.


Figure 1: Artist's view of the deployed SENSE nanosatellite (image credit: USAF/SMC)



In April 2011, the USAF SMC awarded a contract to the Boeing Company (Boeing Defense, Space & Security) of Huntington Beach, CA, for the production and delivery of two nanosatellites (3U CubeSats) to help assess the utility of miniaturized satellites for military space operations.

Boeing worked with ONR (Office of Naval Research), SRI International (Menlo Park, CA), The Aerospace Corporation, and ASTRA (Atmospheric & Space Technology Research Associates) of Boulder, CO to build, test and integrate the nanosatellites. The two spacecraft are powered by highly efficient ultra triple-junction solar cells produced by the Boeing subsidiary Spectrolab.

In December 2012, Boeing delivered the two SENSE nanosatellites to the USAF, meeting the 18 month rapid delivery requirement of the USAF. 7)

One satellite has a CTIP (Cubesat Tiny Ionospheric Photometer) monitoring 135.6 nm photons produced by the recombination of O+ ions and electrons. The other satellite has a WINCS (Wind Ion Neutral Composite Suite) to acquire simultaneous co-located, in situ measurements of atmospheric and ionospheric density, composition, temperature and winds/drifts. The mission data will be used to improve current and future space weather models and to demonstrate the utility of data from nanosatellites for operational weather requirements.

The structure of the SENSE nanosatellites is based on the Boeing-built C2B (Colony 2 Bus), originally developed under a contract with NRO (National Reconnaissance Office). The SENSE bus is highly modular and relies heavily on COTS components. This approach to spacecraft design is intended to minimize development time and cost.


Figure 2: Illustration of the SENSE nanosatellite bus, its subsystems and sensor complement (image credit: USAF/SMC)

ADCS (Attitude Determination and Control Subsystem): Both nanosatellites are 3-axis stabilized. The ADCS employs a diverse collection of sensors and actuators to provide inertial three-axis control to 0.5º (3σ) or better. Attitude and position knowledge are measured using star cameras, IMUs (Inertial Measurement Units), magnetometers, and GPS. For control, SENSE has four reaction wheels and three torque coils.

- Attitude sensing is provided with two star cameras, sun sensors, magnetometers, and GPS

- Actuation is provided by a reaction wheel assembly (4 RWs, 3 torque coils)

- A dosimeter is included into the bus design.

EPS (Electrical Power Subsystem): The spacecraft are powered using one bi-fold, one tri-fold, and one body-mounted solar array with ultra-triple junction solar cells on each panel. This arrangement provides a maximum power production capacity of 37 W. Power is stored in six lithium-ion cells capable of providing the vehicle with 10 W average power.

- Generation of 9.5-10.5 W of average on orbit power (orbit dependent)

- Generation of 37 W of peak power when the solar panels are pointed orthogonal to the sun.

- Electrical & data interfaces: 9V-12.6V DC power; RS-422; FTSH-110-01-L-DV-K.


RF communications: The SENSE vehicles are equipped with a USB (Unified S-Band) transceiver designed to operate at 4 kbit/s uplink and 1Mbit/s downlink. SENSE carries a miniaturized encryption module in its full-duplex transceiver that enables 256 bit type II encryption.


Figure 3: Block diagram of the communications subsystem (image credit: USAF/SMC)


Figure 4: Photo of the S-band radio, diplexer and encryption module (image credit: USAF/SMC)

Each SENSE nanosatellite has a mass of ~ 4 kg.


Figure 5: Illustration of the deployed SENSE nanosatellite (image credit: USAF/SMC)


Figure 6: SENSE space vehicle integration (image credit: USAF)


Launch: The two SENSE nanosatellites were launched on Nov. 20, 2013, as secondary payloads on the ORS-3 (Operationally Responsive Space-3) mission of AFRL (Air Force Research Laboratory). The launch site was MARS Mid-Atlantic Regional Spaceport), located at NASA's Wallops Flight Facility,Wallops Island, VA. The launch vehicle was a Minotaur-1 of OSC (Orbital Sciences Corporation). The primary payload on ORS-3 was STPSat-3 (Space Test Program Satellite-3), a minisatellite mission of the USAF-SMC. 8) 9) 10) 11)

Orbit: Near-circular orbit, altitude = 500 km, inclination = 40.5º.

NASA's LSP (Launch Services Program) ELaNa-4 (Educational Launch of Nanosatellite-4) will launch eight more educational CubeSat missions. The ELaNa-4 CubeSats were originally manifest on the Falcon-9 CRS-2 flight. When NASA received word that the P-PODs on CRS-2 needed to be de-manifested, LSP immediately started looking for other opportunities to launch this complement of CubeSats as soon as possible. 12)

Note: The ELaNa-4 CubeSats were originally manifested on the Falcon-9 CRS-2 flight (launch of CRS-2 on March 1, 2013). However, when NASA received word that the P-PODs on CRS-2 needed to be de-manifested, NASA's LSP (Launch Services Program) immediately started looking for other opportunities to launch this complement of CubeSats as soon as possible. 13) 14) 15)

Secondary Payloads: The secondary payloads on this flight consisted of 26 experiments comprised of free-flying systems and non-separating components (2 experiments). ORS-3 employed CubeSat wafers, which enable secondary payloads to take advantage of excess lift capacity unavailable to the primary trial. 16) 17)


ORS-3 mission sponsor

Spacecraft provider

No of CubeSat Units

ORS-1, ORSES (ORS Enabler Satellite)

ORS (US Army)

Miltec Corporation, Huntsville, AL


ORS-2, ORS Tech 1

ORS Office

JHU/APL, Laurel, MD


ORS-3, ORS Tech 2

ORS Office

JHU/APL, Laurel, MD



SOCOM (Special Operations Command)

LANL (Los Alamos National Laboratory)

1 x 3




1 x 3




1 x 3




1 x 3


STP (Space Test Program)

SMC/XR USAF, Boeing Co.




SMC/XR USAF, Boeing Co.




NSF (National Science Foundation)



NRO (National Reconnaissance Office)

Lawrence Livermore National Laboratory


Black Knight-1


US Military Academy, West Point, NY




US Naval Academy, Annapolis, MD




Naval Postgraduate School, Monterey, CA




University of Hawaii, Manoa, HI




St Louis University, St. Louis, MO




University of Alabama, Huntsville


SPA‐1 Trailblazer


COSMIAC, University of New Mexico


Vermont Lunar CubeSat


Vermont Technical College, Burlington, VT




University of Florida, Gainsville, FL




University of Louisiana, Lafayette, LA




Drexel University, Philadelphia, PA




Kentucky Space, University of Kentucky




NASA/ARC, Moffett Field, CA


TJ3Sat (CubeSat)


Thomas Jefferson High School, Alexandria, VA


Table 1: ORS-3 manifested CubeSats & Experiments (Ref. 16)

ORS and CubeStack:

• ORS (Operationally Responsive Space) partnered with NASA/ARC and AFRL to develop & produce the CubeStack

• Multi CubeSat adapter provides "Low Maintenance" tertiary canisterized ride capability

• ORS-3 Mission: Will fly 2 CubeStacks. This represents the largest multi-mission launch using a Minotaur I launch vehicle (26 free flyers, 2 experiments).


Figure 7: Illustration of the CubeStack, (consisting of wafers) configuration (image credit: ORS, Ref. 16)

The CubeStack adapter structure is a design by LoadPath and Moog CSA Engineering. 18)



Status of mission:

• August 2015: A fundamental objective of the SENSE project was to present DoD decision makers with an understanding of the benefits and costs of integrating CubeSats into an operational architecture. SENSE helped establish how this class of vehicle—one radically different in scale from what SMC typically procures—can be acquired, launched and operated to support future SMC missions. Early on, it was decided that the SENSE ground architecture would be based upon assets organic to the Advanced Development and Directorate (SMC/AD). The intent of this decision was to create a ground segment capability that will remain in place to support subsequent national security space CubeSat and Nanosatellite missions. 19)

- Further, SENSE was a pathfinder for exploring the mission capabilities achievable using a 3U form factor while still addressing uniquely military aspects of requirements for spacecraft design. Such aspect features include military grade data encryption, radiation tolerance, and compliance with the SMC mission assurance process. To provide a meaningful demonstration,the SENSE project developed a complete mission architecture that could demonstrate an end-to-end flow of mission data collected by the spacecraft through a processing segment to assess the quality of its ionospheric measurements.

- The principal objective of the SENSE ionospheric science payloads was to provide data that could be usefully ingested by prototypes of the ionospheric prediction models, in particular,GAIM ( Global Assimilation of Ionospheric Measurements), employed by the Air Force Space Weather Agency. To demonstrate this objective, the SENSE system was required to provide SEM (Space Environmental Monitoring) data to the Air Force Research Laboratory at Kirtland Air Force Base in a format that can be directly ingested by GAIM and also be capable of performing frequent downlink contacts to demonstrate that operational SEM data latency requirements are satisfied.

- The ground segment was arguably the most successful component of the SENSE mission. The Air Force operations team, composed primarily of second and first lieutenants, performed extremely well given the complexity of the SENSE mission. A primary goal of SENSE was to provide a training test bed for space acquisition officers. The operations team was also deeply involved in development of the ground architecture and the operational procedures for SENSE.

- Another successful piece of the ground segment was the GSTR (Global Space Telemetry Resource) antenna at Manzano, procured specifically for SENSE but intended to provide a leave behind capability upon conclusion of the SENSE mission. The auto-track feature specifically increased the contact success rate at low elevation passes. In addition to the GSTR antenna, SENSE demonstrated successful vehicle communication using the AFSCN (Air Force Satellite Control Network) for downlink, and another antenna at Blossom Point Tracking Facility with full uplink and downlink capability.

- The foundation of the ground segment was the Neptune CGA (Common Ground Architecture) software, which provided the backbone for all operations for SENSE. Using Neptune CGA, the operations team was able to consistently demonstrate "lights out" operation of SENSE from both Manzano and Blossom Point with data rates near 1 Mbit/s, paving the way for future minimally manned missions. Furthermore, the flexibility of CGA allowed the operations team to customize the graphical user interface (GUI) command screens, and prioritize the most important SENSE-specific telemetry indicators.

- The SENSE ground segment, along with a spacecraft EM (Engineering Model) was used to successfully demonstrate the use of cloud technology for satellite operations. The SENSE satellite operators were able to conduct nominal command and control operations of the SENSE EM, through its radio interface, from a virtualized application hosted on MilCloud, a cloud service offered by DISA (Defense Information Systems Agency) to DoD customers. For this demonstration, Neptune CGA was virtualized and hosted on MilCloud VDCs (Virtual Data Centers) physically located in Oklahoma City and Montgomery, AL. Using MilCloud, a contingency failover scenario between the two VDCs was successfully demonstrated.

- Finally, SENSE was the first CubeSat to use unified S-band frequencies with NTIA(National Telecommunications and Information Administration) frequency assignment and coordination. Altogether, the ground segment for SENSE represents an impressive distributed ground architecture with leave-behind capability to fly the next minimally manned small satellite mission.

- On SV-1, the team's highest priority was to detumble the vehicle and transition it into a two-axis stabilized sun-pointing orientation called sun safe mode. This is the default mode the vehicle was supposed to default to after launch. In reality, it took the team three months on-orbit test to calibrate the sun sensors and magnetometers and integrate these measurements into control outputs that enable the vehicle to maintain sun pointing.

- One significant, though understated accomplishment of the SENSE mission is that SV-1 exceeded the designed mission life of 12 months using primarily COTS components. A secondary goal of the SENSE mission was to investigate the survivability of COTS components for future CubeSat missions. During 16 months on-orbit, the SENSE mission raised the TRL (Technology Readiness Level) and demonstrated the reliability of several science payloads and bus components. The CTECS sensors on both vehicles provided useful radio occultation and navigation data. SV-1 collected CTECS data for varying periods over 127 days and SV-2 collected CTECS data over eight days.

- Another accomplishment with the space segment is the flight software update implemented on SV-1 during the final months of operations. After attempting (unsuccessfully) to resolve on-orbit command and data handling issues through software parameter changes, a full update of the flight software was implemented on the vehicle. This effort was a significant accomplishment in which the SENSE program demonstrated the benefits of the "fly-fix-fly" approach inherent in CubeSat development (Ref. 19).

• May 4, 2015: One of the two experimental U.S. Air Force upper-atmosphere-monitoring satellites (SENSE-1) that launched in 2013 but quickly began having problems, reentered and burned up on March 21, 2015, service officials said. SENSE was designed to collect data to help characterize and forecast changes in the ionosphere. Understanding upper atmospheric conditions is important to the Air Force because they can affect GPS signals from space. 20)

- Air Force officials continue working with the satellite still on orbit, which "experienced a critical solar array deployment failure" upon launch, according to Air Force spokeswoman Peggy Hodge. The satellite has had "severely limited" capability.

• In early spring of 2014, SV-1 remains stable in sun safe mode and is days away from completing testing to transition to its operational attitude. Progress has been made with SV-2, but power limitations with the vehicle have resulted in intermittent losses of communications and slower progress (Ref. 21).

- Once SV-1 reaches its operational attitude, bus checkout will be nearly complete. Completing bus checkout signifies the end of the launch and early orbit period and the beginning of nominal operations. Bus checkout is complete when the functionality of all subsystems and sensors has been verified and the bus has been transitioned to its operational attitude. In the case of SV-1, LVLH (Local Vertical Local Horizontal) is the operational attitude. LVLH keeps the CTIP sensor nadir pointing with high precision. If a positive power balance cannot be maintained indefinitely in LVLH, it may be necessary to temporarily slew toward the sun and/or duty cycle the CTIP/CTECS payloads to maintain positive power. For SV-2, sun safe is the operational attitude the vehicle will spend the majority of its life in. Excursions from sun safe to LVLH will be conducted with SV-2 to characterize WINCS performance, but these excursions are not considered nominal operations at this time.

- Going forward, the completion of bus checkout marks the beginning of the sensor verification and validation study. AFRL/RV is the lead organization for conducting this study. The study is scheduled to last for approximately six months. Pending the quality of the on-obit data, the SENSE mission is working to incorporate mission data into space weather models such as GAIM. In December 2014, the SENSE mission will conclude its thirteen month on-orbit assessment.

• The SENSE mission experienced a number of challenges following launch. First, SENSE operators were unable to obtain an accurate element set for tracking the vehicles for sixteen days following launch. An element set is a standardized method for reporting a space vehicle's orbital parameters and is needed for communications to estimate when and where a satellite will pass over ground antennas. The SENSE program was reliant upon the Air Force's JSpOC (Joint Space Operations Center) for generating element sets. Due to an international launch and a national security space launch that occurred near the same time as the ORS-3 launch, high quality element sets were delayed as Air Force space tracking assets were allocated to higher priorities than the CubeSats on the Enabler mission. 21)

- An additional contributing factor that delayed initial contact was that both vehicles experienced solar array deployment failures. Nominally, the vehicles were designed to autonomously deploy their solar arrays thirty minutes following launch and autonomously point the arrays toward the sun. In reality, SV-1's bi-fold array and SV-2's bi-fold and tri-fold arrays failed to deploy as designed. Both vehicles attempted to orient the arrays toward the sun after launch, but the ADCS was not designed to operate in an off-nominal configuration. The vehicles were essentially crippled with no control and limited capacity to produce power.

- To further compound the vehicle acquisition problem, both vehicles were programmed to activate a beacon if contact was not established within twenty-four hours following launch. The vehicle's beacon was not long enough in duration to enable the SENSE mission's ground antennas to lock on to each vehicle using the antenna's auto track feature. Under nominal conditions, the beacon transmits for ten seconds of every minute. In a low power situation, the beacon transmits for a couple seconds or not at all. These intermittent beacons were the only means of distinguishing the SENSE vehicles from the other twenty-six CubeSats during the first two weeks on-orbit. The sparse beacons combined with the high uncertainty of the element sets made tracking the vehicles extremely challenging.

- Once operators finally obtained an accurate element set for each CubeSat, the real work began of assessing the vehicles' state of health. While the few, brief observations of the beacons were encouraging to operators, the beacons' frequent transmissions were a significant draw on the vehicles' batteries. The first step was to disable the beacons and allow the batteries to recharge. Once enough charge was stored, telemetry was downloaded allowing the team to diagnose the solar array deployment anomaly that occurred following launch. The team also conducted an extensive root cause and corrective action plan to assess why the solar arrays failed to deploy. SV-2, the vehicle with neither array deployed, could not sustain prolonged communications. The team transitioned SV-2 to a low power state for approximately one month in attempt to recharge the vehicle's batteries. During this time, operators pressed ahead with diagnosing problems with SV-1's control algorithms.

• Space vehicle discrimination methods must improve if nanosatellites are to be used in future operational architectures

• Small satellite ≠ low complexity

• Many SENSE space vehicle subsystems are performing exceptionally well:

- Li-Ion batteries

- S-band radio

• Common Ground Architecture (CGA):

- Versatile and stable software platform

- Enables resilient command and control in contingency operations.

Table 2: On-orbit lessons learned 22)



Sensor complement: (WINCS, CTIP, CTECS, Micro Dosimeter)

The two SENSE space vehicles are identical with the exception of two of their payloads. Space Vehicle 1 (SV-1) carries the CTIP (Compact Tiny Ionospheric Photometer), while Space Vehicle 2 (SV-2) is equipped with the WINCS (Winds-Ion-Neutral Composition Suite). Each vehicle is also equipped with a CTECS (Compact Total Electron Content Sensor) and a micro dosimeter.


Figure 8: Illustration of sensor complement on SV-1 and SV-2 (image credit: USAF/SMC)

WINCS (Wind Ion Neutral Composition Suite)

The WINCS instrument was designed and developed jointly by NRL (Naval Research Laboratory) and NASA/GSFC for ionosphere-thermosphere investigations in orbit between 120 and 550 km altitude. The objectives are to acquire simultaneous collocated, in-situ measurements of atmospheric density, composition, temperature and winds (Ref. 1). 23)

The four WINCS instruments are:

- WTS (Wind and Temperature Spectrometer)

- IDTS (Ion Drift and Temperature Spectrometer)

- NMS (Neutral Mass Spectrometer)

- IMS (Ion Mass Spectrometer)

WINCS has a size of 7.6 cm x 7.6 cm x 7.1 cm and a mass of ~ 0.85 kg (including interface electronics) with a total power consumption of 1.3 W.


Figure 9: Schematic view of the WINCS WTS/IDTS devices (image credit: USAF/SMC)


Figure 10: Cut-away view of the WINCS devices (image credit: USAF/SMC)

WINCS theory of operation:

• WTS/IDTS: Ionize incident air stream to measure the angular distribution at many angles simultaneously while scanning energy in time

• IMS/NMS: Time of Flight mass spectroscopy.

SDEA (Small Deflection Energy Analyzer): SDEA is at the heart of the WTS/IDS devices (Figure 11). The SDEA instrument has two identical but mirrored sensors. Ions enter from the upper (lower) left and deflect downward (upward) as they continue to the right toward the exist slit in the lower (upper) right. The exit slit plane is a circular cylinder section with vertical axis passing through the entrance slit — this defines the angle imaging function of SDEA. 24)

The Figure 11 (bottom) considers only one of the two main deflection chambers. The upper exit plates are biased at voltage VSDEA while the entrance and lower plates at at ground potential. The inhomogeneous field produced in the upper left region of the cavity produces the focusing shown in the incident ions, which makes the exit slit smaller than the entrance slit; this is a desirable feature in reducing the unwanted UV photons that may reach the MCP (Micro Channel Plate) detector placed beyond the exit slit.


Figure 11: SIMION trajectories through the multi-chamber WTS/IDS pair (image credit: UMich)

Legend to Figure 11: The top left cavity corresponds to the ion source cavity for the WTS half of the spectrometer. An expanded view of the main deflection chamber with various SIMION trajectories is at the bottom of Figure 11.


Figure 12: Layout of the 4 spectrometers on WINCS (image credit: NASA)

Legend to Figure 12: (2 left side images) WTS and IDS are paired off in two spectrometer modules with mutually perpendicular FOVs as shown by the two pairs of long slits on the right side. The two round apertures on the other module show the entrances for the IMS and IMS.
Right image: Cross-section of the WTS/IDTS on WINCS. This module shows the parallel ion and neutral paths through the instrument. The neutral path is highlighted in yellow. The multi-chamber WTS/IDS pair provides four separate chambers that provide excellent photon rejection capability.

Measuring winds and ion drifts: Neutral wind and ion-drift measurements are obtained from the angular and energy distributions of the particle flux. Two separate analyzers measure the angular-energy distributions in two perpendicular planes, their collinear axes pointing within a few degrees of the ram direction (e.g. WTS1 and WTS2). Each analyzer spans a 30º x 2º FOV in 15 angular pixels, each 2º x 2º pixel scanning the energy distribution in 20 energy steps with the energy-analyzer voltage. Figure 12 (right) shows the 30º FOV of one of the instruments - particles entering from as far away as 15º from ram can enter the instrument (i.e., a cross-track speed of ~ 2km/s). Because the slit half-height is 1º, it is desirable to have the satellite pointed within 1º of ram, but the FOV is 15º, so the satellite can technically be pointed within 15º of ram to accomplish the measurements (Ref. 24).

The IMS/NMS instruments use two GEMS (Gated Electrostatic Mass Spectrometer) devices developed at GSFC. GEMS utilizes a SDEA device, turning the potential on and off. When the potential is turned off, particles streaming into the detector at the entrance slit will continue in a straight line along the length of the SDEA, never reaching the exit slit. After the potential is turned on, the charged particles will be deflected. For the particles that were just about to hit the back wall, the deflection will be minimal, but for the particles that were halfway across the chamber, the deflection will be greater, since they will have more time to accelerate. The acceleration is a function of mass. Therefore, a very short time after the potential is turned on, low mass particles will be deflected enough such that they start to come through the exit slit. Some time later, heavier particles will start to come through. The will result in a temporal distribution that looks like Figure 13, which enables mass resolution of the particles. 25)


Figure 13: An example plot of ion accumulation versus time after the potential is turned on in GEMS (image credit: NASA/GSFC)


Figure 14: Photo of the WINCS instrument (image credit: NASA/GSFC) 26)


CTIP (CubeSat Tiny Ionospheric Photometer)

CTIP is a nadir-oriented UV photometer developed by SRI (Stanford Research Institute) International. CTIP is a miniaturized version of TIPS (Tiny Ionospheric Photometer System) heritage, developed by NRL (Naval Research Laboratory) and flown on the FormoSat-3/COSMIC constellation. The goal of CTIP is to measure ionospheric plasma column density to refine the global ionospheric models. CTIP gathers data to characterize the ionosphere through the natural decay rate as seen in recombination of O+ ions and electrons. CTIP measures UV airglow at 135.6 nm.

• Atomic oxygen ions constitute the primary ionospheric species in the F-region

- 135.6 nm photons are emitted spontaneously

- from the recombination of atomic oxygen ions

- O+ + e- → O (5P) + hν135.6

• O+ and e- are in equal number and 135.6 nm emission is proportional to the path integral of [O+] squared.

The CTIP instrument has a volume of < 1000 cm3, a mass of < 1 kg, and an orbit average power consumption of 2-3 W. The CTIP instrument matches the TIPS performance.


Figure 15: CTIP optics based on heritage COSMIC TIP design (image credit: USAF/SMC)


Figure 16: View of the CTIP instrument and its components (image credit: SRI International)


Figure 17: Photo of the CTIP instrument (image credit: USAF/SMC)


CTECS (Compact Total Electron Density Sensor)

CTECS is a GPS occultation sensor to measure the line-integrated electron densities in the Earth's ionosphere. Tracking capability of L1, L2, L2C signals.

• The primary data product: line-of-sight TEC to all GPS satellites in view for ingest into ionospheric models

• Secondary data product: L-band scintillation observations.

The CTECS instrument consists of the NovAtel OEMV-2 receiver and a custom dual patch antenna.

- 1557 MHz and 1227 MHz

- A LNA (Low-Noise-Amplifier) is placed between antenna and receiver.


Figure 18: Photo of the CTECS sensor, a NovAtel OEMV-2 receiver (image credit: USAF/SMC)

- Receiver: The OEMV-2 has 72 channels available for tracking L1, L2, L2C and GLONASS signal capability (but not used). The receiver can track 14 GPS satellites simultaneously (L1 and L2). The board measures 60 mm x 100 mm x13 mm with a mass of 56 g. It requires a 3.3 VDC input voltage and consumes ~1.2 W. The most important feature of the OEMV-2, for its use as a CubeSat GPSRO sensor, is its ability to accept modified software/firmware through NovAtel's API (Application Programming Interface). Receiver mass = 153 g.

- Antenna: There are three parts to the antenna: a dual patch antenna, an integral 90º hybrid, and a low noise amplifier (LNA). The dual patch antenna is mounted onto the ground plane that includes a stripline 90º hybrid internally. The entire antenna is then mounted into a bracket for placement on the spacecraft. The antenna has a centered ground via connected to both patch conductors to prevent electrical charge buildup. To get better axial-ratio bandwidth, dual probes and a 90º degree hybrid were used (instead of a single probe technique) for RHCP (Right Hand Circular Polarization). The overall dimensions of the antenna are 7.6 cm x 7.6 cm x 1 cm. The gain pattern is hemispherical with a 10 dB gain reduction 90º from the antenna boresight. The bandwidth centered on L1 and L2 is 20 MHz. The gains of the antenna are 6.2 dBic and 6.4 dBic for L1 and L2, respectively.

The CTECS instrument is being provided as a government furnished equipment to the program by The Aerospace Corporation, El Segundo, CA. The PSSCT-2 (Pico Satellite Solar Cell Testbed-2)nanosatellite of The Aerospace Corporation hosted already the CTECS instrument (demonstration mission) in the timeframe June 20 - December 8, 2011 on a reentry trajectory from the ISS (deployed from the STS-135 Space Shuttle mission). 27)

For the SENSE mission, both nanosatellites are equipped with a CTECS (Compact Total Electron Density Sensor) instrument to provide radio occultation measurements of TEC (Total Electron Content) and L-band scintillation.


Micro Dosimeter:

Micro Dosimeter of TMT (Teledyne Microelectronic Technologies): µDOS001 is a compact microcircuit which directly measures TID (Total Ionizing Dose) absorbed by an internal silicon test mass. The test mass simulates silicon die of integrated circuits on-board a host spacecraft in critical mission payloads and subsystems. By accurately measuring the energy absorbed from electrons, protons, and gamma rays, an estimate of the dose absorbed by other electronic devices on the same vehicle can be made. The Micro Dosimeter can operate from a wide range of input voltages. The accumulated dose is presented to three DC linear outputs and one pseudo-logarithmic output giving a dose resolution of 14 µRad and a measurement range up to 40 krad. These outputs are intended to be directly connected to most ADCs (Analog-to-Digital Converters) or spacecraft housekeeping analog inputs (0-5 V range), which makes minimal demands on the host vehicle. The Micro Dosimeter incorporates a test function to allow electrical testing of the hybrid without the need for a radiation source. 28)

The objective of the µDOS001 device on the SENSE nanosatellites is to provide radiation dosage for measurement and to correlate system performance with exposure:

• First compact microcircuit that provides a repeatable measurement of radiation dose and dose rate over a wide range of energies

• Enables routine monitoring of spacecraft radiation environment

• Custom microchip in a small footprint package for low mass and power

• Correlates environmental models and ray-tracing analyses with real in-flight measurements.


Figure 19: Photo of the Teledyne µDOS001device (image credit: USAF/SMC, Ref. 1)

Device dimensions, mass

3.6 cm x 2.5 cm x 0.1 cm, 20 g


10 mA, 13 VDC to 40 VDC, 3 DC linear outputs

Dose resolution
Device survivability

14 µRad
up to 40 kRad

Data updates

Every 30 s


- Class K space qualified
- 1 pseudo log
- 100 kRad total count
- Test input bypasses silicon detector for circuitry detection
- Volatile count retention

Table 3: Specifications of the Teledyne Microdosimeter



Ground segment:

Early on, it was decided that the SENSE ground architecture would be based upon assets organic to the Space Development and Test Directorate (SMC/SD). The intent of this decision was to create a ground segment capability that will remain in place to support subsequent national security space CubeSat and NanoSat missions. The principal objective of the SENSE ionospheric science payloads is to provide data that can be usefully ingested by prototypes of the ionospheric prediction models, e.g. GAIM (Global Assimilation of Ionospheric Measurements), employed by the Air Force Space Weather Agency. To demonstrate this objective, the SENSE system is required to provide SEM (Space Environmental Monitoring) data to the Air Force Research Laboratory at Kirtland Air Force Base in a format that can be directly ingested by GAIM and also be capable of performing frequent downlink contacts to demonstrate that operational SEM data latency requirements are satisfied.

The distributed ground segment architecture uses three stations for the acquisition of the SENSE mission data located at Manzano, NM, Blossom Point, MD, and AFSCN (Air Force Satellite Control Network), Oakhanger, UK. SENSE operations will be conducted from the RSC (RDT&E Support Complex) at Kirtland AFB. The Space Development and Test Directorate (SMC/SD) is the owner-operator of the SENSE ground system, and the Space Vehicles Directorate at AFRL (AFRL/RV) is responsible for data processing, analysis, and distribution.

The SENSE ground system consists of several antenna sites around the globe controlled from the Research, Development, Test and Evaluation (RDT&E) Support Complex (RSC) at Kirtland AFB, NM. As shown in Figure 20, SENSE will utilize antennas located at Kirtland AFB, NM. To link these sites together, the SENSE ground system relies on the CGA (Common Ground Architecture) developed by the NRL (Naval Research Laboratory).


Figure 20: Overview of ground station locations (image credit: USAF/SMC, Ref. 5)


Figure 21: SENSE ground terminal at Kirtland Air Force Base, New Mexico (image credit: USAF)

The requirements call for a "lights-out" ground architecture with leave-behind capability to fly the next minimally-manned satellite mission.

CGA is a command and control software tool that enables highly autonomous operation of the SENSE satellite system. CGA allows operators to manage simultaneous space vehicle passes and schedule "lights-out" operations for times when no operators are present. Further, CGA is a key enabler of SENSE's requirement to achieve data latencies of 90 min or less during high-tempo operations. "Data latency" is the duration of time between taking measurements on orbit to delivering processed data to space weather models for use by the space weather community. Lastly, CGA and the SENSE ground assets provide a leave-behind capability that can be readily leveraged by future space programs.


Figure 22: SENSE data flow and partner organizations (image credit: USAF/SMC, Ref. 22)


Figure 23: SENSE system view and ground segment (image credit: USAF/SMC)


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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 (

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