Minimize PSSCT-2

PSSCT-2 (Pico Satellite Solar Cell Testbed-2)

PSSCT-2 is a nanosatellite mission of The Aerospace Corporation (El Segundo, CA) with a size of 12.7 cm x 12.7 cm x 25.4 cm and a mass of 3.7 kg. PSSCT-2 was flown on the final U.S. Space Shuttle mission STS-135 (ISS ULF7 flight of Atlantis, July 8-21, 2011, 13 day mission). The PSSCT-2 nanosatellite was ejected on June 20, 2011 from the Space Shuttle Atlantis, shortly before the Shuttle reentry into a low (~ 360 km altitude) orbit.

The PSSCT-2 nanosatellite is a follow-up mission to PSSCT-1 which was flown as a secondary payload to the ISS on STS-126 mission of the Space Shuttle Endeavour to ISS (Nov. 14 - 30, 2008). PSSCT-1 was deployed from the Shuttle cargo bay on Nov. 29, 2008 using the SSPL (Space Shuttle Picosatellite Launcher) device. - PSSCT-2 was developed by The Aerospace Corporation with support from the SMC/XR (Space and Missile Systems Center's Development Planning Directorate). 1) 2) 3)


Figure 1: Photo of the PSSCT-2 nanosatellite as it is released from the launcher inside the Atlantis Space Shuttle bay (image credit: NASA)

Legend to Figure 1: The launcher for PSSCT-2 was developed at The Aerospace Corporation and has been used to deploy several nanosatellites over the years from the Space Shuttles.

Background: The Aerospace Corporation was contacted by the DoD/STP (Space Test Program) in Sept. 2010. The PSSCT project was asked if there was interest in a reprise of the PSSCT-1 mission to be deployed by NASA’s “Launch on Need” flight (STS-135). PSSCT-1 was developed as a rapid turn-around platform for testing new solar cell technologies in different radiation environments, but it abruptly fell silent after 110 days of nominal operation in March 2009.

The PSSCT project accepted the challenge of building a “reflight” within nine months for a flight that might not occur. To make it even more challenging, we decided to redesign the avionics to increase radiation tolerance, improve on-orbit measurements of solar cell IV-curves, add a second communications transceiver, add a space weather instrument, add a GPS receiver for on-orbit position determination, and add three-axis attitude control (PSSCT-1 was a spinner).

The PSSCT-2 was also an unprecedented opportunity to be the last satellite deployed by a U.S. Space Shuttle, so the project increased the image resolution of the visible cameras from 0.3 to 2 million pixels in order to provide the last quality images of Atlantis in space. PSSCT-2 was ejected by the Shuttle Atlantis into a 365 km x 380 km orbit on July 20, 2011 just before its final reentry.


Figure 2: Photo of Space Shuttle Atlantis taken by the PSSCT-2 nanosatellite (image credit, NASA, The Aerospace Corporation)

Legend to Figure 2: A photograph of Atlantis taken using a 180º FOV (Field of View) “fisheye” lens, mounted on the aft end of the PSSCT-2 nanosatellite satellite, within 10 seconds after ejection. The launch tube is near the center of this image on front starboard (right) side of the Shuttle bay.

PSSCT-2 will serve as low cost risk reduction for the upcoming SMC SENSE (Space Environmental Monitoring Nanosat Experiment) mission, because it contains the Aerospace Corporation's CTE CS (Compact Total Electron Content Sensor) that characterizes the ionosphere by measurement of the occultation of GPS signals - a precursor of an instrument with the same function on SENSE.



The PSSCT-2 nanosatellite was a considerable improvement over the original PSSCT-1. The power system was redesigned to be less flexible, but more efficient, and four lithium-ion 18650 batteries instead of two provided 40 Whr of stored energy. On-orbit average power was 2-5 W, depending on spacecraft orientation.


Figure 3: Schematic view of the PSSCT-2 nanosatellite (image credit: The Aerospace Corporation)

Active attitude control with better than 10º pointing accuracy was needed for taking meaningful solar cell performance data, for firing thrusters to raise orbit, and for pointing the CTECS (Compact Total Electron Content Sensor), the GPS radio occultation space weather instrument, in the anti-flight direction. The project succeeded as evidenced by the radio occultation measurements.

ACS: Absolute attitude sensors on-board PSSCT-2 included Aerospace-developed sun and Earth nadir sensors, plus COTS (Commercial-of-the-Shelf) magnetic field sensors from Honeywell. A COTS MEMS IMU (Inertial Measurement Unit) from Analog Devices provided angular rotation rates, inertial angular position, and an extra set of magnetic field sensors. Figure 3 shows a schematic drawing of the spacecraft. The +Z face is the nadir-pointing face.

Two-axis sun sensors: PSSCT-2 had two 2-axis sun sensors to enable pointing of two opposite faces at the sun for measuring solar cell current-voltage curves under near-normal solar incidence. The project designed this sensor using an aperture plate and QPD (Quad Photodiode) geometry ( Figure 4). A square aperture allows directed sunlight to impinge on a QPD located below the aperture. The QPD has four square photodetectors, and the photo currents collected by each electrode uniquely specify an X-Y location for the centroid of illumination. This ultimately yields the angular position of the sun in two directions with respect to the surface normal of the sun sensor. The QPD was an SD 085-23-21-021 from Advanced Photonix Inc., in Camarillo, CA.



Figure 4: Schematic diagram of the basic two-axis sun sensor using a position-sensitive detector (image credit: The Aerospace Corporation)

The aperture plate was fabricated from a blackened photo-etched metal foil, bonded to the photodetector can. Interference from Earthshine was minimized by adding a tubular baffle, mounted onto the TO-5 can, to allow a 70º FOV for the sun (Figure 5).


Figure 5: Schematic cross section of the sun sensor (image credit: The Aerospace Corporation)

Each quad photodiode was connected to ground through a 300 ohm resistor, and the voltage across each resistor was measured by a 16 bit A/D converter with an internal 8X gain preamplifier. Measured voltage was proportional to quad photodiode current, with a maximum possible voltage of 0.21-V.

The project measured the angular response of the sun sensors in a solar simulator as the solar incidence was varied from -45º to +45º along each axis. These sensors are fairly linear over a ± 34º range with varying angular offsets of up to a few degrees. Different offsets resulted from variations in how the photodiode chip was bonded into the TO-5 can by the manufacturer, and how well we aligned the aperture plate onto the TO-5 can. Figure 5 shows the error in measured angle after implementing offset and proportionality corrections for a single axis in one of our sensors. A maximum error of ± 0.5º was measured over a ±34º input angle range, which could have been further reduced to a ± 0.1º error by including second-order and higher fitting terms. Most of the error in Fig. 5 is due to refraction of light rays through the glass cover on the quad photodiode.


Figure 6: Output angle errors for one axis of one of the sun sensors using first-order (offset and gain) fitting (image credit: The Aerospace Corporation)

Earth nadir sensor: PSSCT-2 utilized a novel array of nine COTS Melexis MLX90615 infrared (IR) thermometers to form a patent-pending Earth nadir sensor. Each IR thermometer contains a MEMS thermopile sensor with a 5.5 - 14 µm spectral response directly connected to a digital signal processor within a small (4.7 mm diameter by 2.7mm high) TO-46 transistor can. Each sensor is factory-calibrated to within 0.5ºC over a 0 - 50ºC range and outputs digital temperature data on an SMBus. The initial testing revealed that some of these sensors could leak under vacuum, thus reducing conductive and convective heat transfer within the can, leading to inaccurate temperature readings. The project selected detectors that didn’t appear to leak after a week under vacuum. Unfortunately, the flight experience showed that there were still had leaks in at least two detectors; now, holes are drilled in each detector to vent them and force uniformity in internal pressure, and hence, thermal response.

These sensors have a wide, ~100º FWHM angular response to infrared radiation. In order to determine Earth nadir direction, the project mounted pairs of sensors such that during nadir-pointing, one sensor would see only the ~300ºK Earth, and the companion sensor would see most of Earth in this direction plus some 3ºK empty space. The reduced temperature reading of the “Earth plus space” sensor, relative to the temperature of the “Earth only” sensor, was used to determine the relative proportion of warm Earth in the total FOV of the “Earth plus space” sensor; thus establishing the angular offset in that direction. The project arranged pairs of sensors to cover +X, -X, +Y, and –Y angular offsets from the nadir-pointing +Z direction.

Figure 7 shows a schematic drawing of the Earth nadir sensor as viewed from the nadir direction. The +X, -X, +Y, and –Y “Earth Plus Space” sensors are mounted so that their surface normals are 34º from the array surface normal (+Z direction), while the “Earth” sensor normals are 20º from the +Z direction. When nadir-pointing, the temperature differences between the “Earth” and “Earth plus space” sensor pairs will all be the same for the +X, -X, +Y, and -Y directions. Figure 8 shows a photograph of the 2.5 cm x 2.5 cm Earth nadir sensor array.


Figure 7: Schematic drawing of the Earth nadir sensor array as seen from the nadir direction (image credit: The Aerospace Corporation)


Figure 8: Photo of an Earth nadir sensor array used on PSSCT-2, the ruler is in inches (image credit: The Aerospace Corporation)

Ground testing of the Earth nadir sensor array determined the angular offset errors and proportionality constants used in the flight attitude control software. This sensor was capable of providing the Earth nadir direction within 0.5º along two axes.

Magnetic field sensors: A Honeywell HMC6042 two-axis magnetic sensor was used plus an HMC1041Z one-axis sensor to form a three-axis magnetometer with a ~0.15 mGauss sensitivity. These analog-output sensors were coupled to a 16 bit A/D converter read by a Microchip “PIC” processor that calculated the X, Y, and Z magnetic field values plus their time derivatives. These now-obsolete sensors were not factory calibrated, but their high sensitivity was needed for our B-dot control algorithms.

IMU (Inertial Measurement Unit): The inertial navigation sensor was an Analog Devices ADIS16405BLM module that included a MEMS three-axis accelerometer, a MEMS three-axis rate gyro, and a three-axis magnetometer in volume of ~16 cm3. The rate gyros were accurate to 0.006º/s for up to 200 seconds of operation after a bias error calibration, and the magnetometers had a sensitivity of 0.5-mGauss. The ADIS sensor package was factory calibrated and was used to provide absolute magnetic field measurements for orientation determination.

Attitude actuators: The attitude actuators on-board PSSCT-2 included three magnetic torque coils and three magnetically-shielded miniature reaction wheels. The torque coils were used to detumble the spacecraft to an essentially non-rotating state before any tracking exercise, while the reaction wheels were used for closed-loop attitude control.

Torque coils: The torque coils were designed to provide the maximum torque possible while staying within the power and size constraints of the PSSCT-2 design. Maximum coil dimensions were set by the interior dimensions of the spacecraft, and the internal placement of various systems. The Z-axis coil could be as large as 11.2 cm x 11.2 cm, but the actual coil is 9.5 cm x 10.1 cm due to physical interference from other spacecraft systems and assembly issues. The X- and Y-axis coils could have been 11.2-cm x 23.9 cm , but similar restraints limited both coils to 10.1 cm x 8.5 cm. The winding cross section determines the final coil mass, the project chose to use a 0.8 cm wide by 0.32 cm deep cross section to provide a coil mass of < 80 g. The mass of three coils is about 5% of the total PSSCT-2 mass. Given the dimensions of the torque coil, a wire gage of 30 was determined by choosing the gage that produced the maximum torque possible while consuming no more than 1 W per coil with a driving voltage of 5 V.

PSSCT-2 was ejected by the Shuttle after leaving the International Space Station into an initial altitude of approximately 375 km with an orbit inclination of 52º. In this orbit, the Earth’s magnetic field strength varies between 0.23 Gauss near the equator to 0.50 Gauss at the maximum magnetic latitude of 52º. Using a 0.3 Gauss as a conservative average magnetic field strength, the torque coils were capable of producing ~1 x 10-5 Nm, for coil power levels below 1 W. Figure 9 shows a photo of the X- and Y-axis torque coils mounted into the PSSCT-2 body.


Figure 9: Photo showing the X- and Y-axis torque coils mounted in the PSSCT-2 body; the mounted reaction wheel assembly and IMU is also shown (image credit: The Aerospace Corporation)

Reaction wheels: The PSSCT-2 reaction wheel assembly used miniature brushless DC motors (Faulhaber 1226 S 006 B K179) with a high vacuum lubricant. The motors were bonded into mu-metal cases to provide overall magnetic shielding, plus heat sinking for the motor’s bearings. Mu-metal wheels were bonded onto the motor shafts to provide additional magnetic shielding. The center of gravity of the wheel was located at the center of the forward bearing to minimize vibration-induced bending of the shaft. Figure 10 shows a cross section of a reaction wheel.


Figure 10: Cross section of a PSSCT-2 reaction wheel (image credit: The Aerospace Corporation)

Each reaction wheel has an axial moment of inertia of 1.5 x 10-6 kgm2, a maximum rotation rate of 60,000 rpm, a momentum storage capability of 9.4 x 10-3 Nms, a maximum energy storage of 29.6 J, and an average torque of 2.2 x 10-3 Nm. Since PSSCT-2 has a total mass of 3.7 kg and moments of inertia of ~2.9 x 10-2 kgm2 along the X- and Y-axes, and ~1.3 x 10-2 kgm2 along the Z-axis, at 50,000 rpm, the wheels can generate spacecraft rotation rates of ±15º/s around the X- and Y-axes, and ±34º/s around the Z-axis.

Three orthogonally-mounted reaction wheels, each with a mass of 39 g, were assembled into an aluminum mounting block, which was attached to the wall of the satellite to provide a heat path for thermal management. The entire reaction wheel assembly had a mass of 225 g, or 6% of the total PSSCT-2 mass. Figure 11 shows the reaction wheel assembly, while Figure 9 shows the assembly mounted to the PSSCT-2 body.


Figure 11: Triaxial reaction wheel assembly of PSSCT-2 (image credit: The Aerospace Corporation)

Attitude control: The Aerospace Corporation small satellites generally use a distributed computing architecture. The project used 24 microprocessors in PSSCT-2 with each primary function having its own microchip “PIC” processor and dedicated memory. A central flight computer was used to interface with the radios and delegate commands to the individual processors. The attitude control system itself used 5 PICs (Figure 12) to handle the Earth and sun sensors, magnetometers, reaction wheels (RWs), torque coils, and IMU. One PIC is used to control the speed of each reaction wheel, one PIC is used for reading the Earth, sun, and magnetic sensors, and one main ACB (Attitude Control Board) PIC reads the high precision IMU and executes attitude control algorithms.


Figure 12: Attitude control processor and sensor tree (image credit: The Aerospace Corporation)

Motor controller: Each reaction wheel uses a three-phase brushless DC motor controlled by a three-phase bridge. Each drive phase consists of one motor terminal driven high, one motor terminal driven low, and one motor terminal left floating. Each of the bridge elements are controlled by a PIC18F1330 that has six built-in PWMs (Pulse Width Modulators), whose duty cycles and phase relationships control the speed and direction of the reaction wheel. Rotor position is read by Hall sensors in each motor.

Magnetic detumbling: The first attitude control mode was magnetic detumbling. Magnetic detumbling is the initial attitude control mode used after spacecraft ejection into orbit, and the first mode to be used before any pointing exercise. It is performed with all reaction wheels turned off, and produces an essentially non-rotating spacecraft.

Simple detumbling is typically accomplished by measuring the time rate-of-change of magnetic field on each orthogonal axis, and by generating a magnetic field proportional to the negative time rate-of-change on each corresponding axis. Each axis is independent and no trigonometric calculations are required. This “B-dot” mode was therefore relatively simple to implement in analog or digital control loops once the proportionality constants were known. PSSCT-2 had an on-board 3-axis magnetometer as a magnetic rate sensor, and digital control based on pulse-width modulation of individual coils. Detumbling was critical due to limited gain control resulting from the reaction wheel controllers, approximately 3º/s for the high moment of inertia axes (X, Y) and up to 6º/s for the low moment of inertia axis (Z). The goal of magnetic detumbling was to reduce the initial spin rates well below these numbers to have enough control of the satellite to perform a maneuver.

The B-dot routine nominally ran on a 1.14 s loop cycle consisting of a 0.14 s magnetometer measurement and a 1.00 s torque coil PWM cycle (Figure 13). It was performed prior to any maneuver and often during a maneuver to keep spin rates low enough for the reaction wheel controllers.


Figure 13: Block diagram of the magnetic detumble algorithm (image credit: The Aerospace Corporation)

Every cycle, the Honeywell magnetometer measured the magnetic field, and the measurement was compared to previous measurement to obtain a rate of change (Bdot). Based on the B-dot measurement, the X, Y, and Z torque coils were independently turned on for a fraction of 1 s to achieve proportional control. On-orbit angular rate measurements using the ADIS IMU gyros showed that the satellite could be actively detumbled from 6º/s about any axis to less than 1º/se after 100 s of operation.

Sun pointing: PSSCT-2 had one sun sensor on each of the +Y and -Y faces (Figure 3) to orient a particular face towards the sun to take IV curves of the solar cells on that face.

Using a single sun sensor, the initial sun lock occurred in three steps: find sun, center sun, and track sun (Figure 14). In the first step, the satellite rotated about the Z-axis for a full revolution while periodically taking sun sensor measurements. If the summed QPD output passed a certain threshold, the sun was coarsely found and the satellite went on to the next step. However, if the sun was not found, the satellite was rotated about the X-axis approximately 30º, and Z-axis rotation was then repeated to continue looking for the sun. This process repeated until the X-axis rotated 180º, covering all possible locations of the sun relative to the spacecraft.

After the sun was coarsely found, the next step was to center the sun on the sensor in one rotation of the X and Z-axes. A sun sensor measurement was taken to give the angle offsets, and the X and Z wheels were turned on for a proportional amount of time to center the sun.

The last step was to track the sun using proportional gain control of the X and Z wheels. Every cycle, a sun sensor measurement was made to determine the independent proportional speed and on-time of the X and Z wheels. The X and Z wheels were then synchronously turned on and separately ran for their allotted time and speed. Once the wheels were fully stopped, the cycle repeated for a user-defined number of cycles. We typically used cycle-limited algorithms to prevent getting stuck in infinite “do” loops.

While tracking the sun with one of the sensors, the satellite was commanded from the ground to measure the IV curves of the solar cells on that face. The satellite was then commanded to break out of sun-lock and to manually rotate 90º about the Z axis so that a different side faced towards the sun. IV curves were again taken for a total of two IV curves for two orthogonal faces. The process was repeated over subsequent passes and the other sun sensor to get all four IV curves for the four faces with solar panels.


Figure 14: Overview of sun pointing algorithm (image credit: The Aerospace Corporation)

Nadir pointing: PSSCT-2 had two reasons for pointing the satellite’s +Z face towards the Earth. First, the +Z face had three cameras for taking pictures of the Earth with its fish eye, medium FOV, and narrow FOV cameras. Second, nadir pointing was integral to the pointing algorithm needed for the CTECS payload, which required that the +X face of the spacecraft be oriented in the anti-flight direction in order to follow a GPS satellite as it headed down toward the Earth’s horizon. Our solution was to fly in a LVLH (Local Vertical/Local Horizontal) attitude with the Z-axis being “vertical” and the X-axis along the anti-flight direction.

Only the Earth sensor pointing algorithm was used and was broken down into two steps as shown in Figure 14. The first step was to find the Earth, similar to how the sun algorithm found the sun, using a single IR thermometer in the center of the Earth nadir sensor array. Since the Earth had a large solid angle in orbit, after a complete X-axis rotation, the Y axis needed only one rotation to cover all possible angles for the Earth’s location. While the X-axis was rotating, the optical thermometer was read once per second. If the temperatures failed to meet a threshold temperature representative of a partial Earth view after a complete revolution, the process was repeated for the Y axis.

The second step used non-proportional gain control of the X and Y axes to center and track the Earth. The original plan was to use a differencing algorithm using all 8 of the outer infrared temperature sensors. However, because two sensors were found to be faulty once on-orbit, the algorithm was changed to only use the difference of two orthogonal anti-pairs (a ± X pair and a ±Y pair). For each cycle, temperatures were measured and the ACB PIC then calculated the temperature differences for each pair. The X and Y reaction wheels were then respectively turned on at a constant speed and time to drive the differences towards zero. Each IR sensor for each pair should see the same amount of Earth plus space when nadir-pointing.


Figure 15: Schematic of the nadir-pointing algorithm (image credit: The Aerospace Corporation)


CTECS (Compact Total Electron Content Sensor) anti-flight:

One of the primary missions of the satellite was to take GPS occultation data to measure line-integrated electron densities in the Earth’s ionosphere. This mission required that the GPS occultation (CTECS) antenna on the +X face point in the anti-flight direction to track descending GPS satellites. The CTECS anti-flight LVLH (Local Vertical Local Horizontal) mode used the Earth nadir sensor to point the +Z face towards nadir, and the Analog Devices magnetometer to rotate about the +Z axis such that the +X face pointed towards the anti-flight direction.

The CTECS anti-flight LVLH mode started with magnetic detumbling followed by nadir-pointing of the +Z face (Figure 16). While nadir pointing the +Z face, the X and Y magnetic fields were measured and compared to a time-tagged lookup table that was previously uploaded to the spacecraft. The Z-axis reaction wheel was then proportionally controlled in both time and speed to minimize the error between the lookup table value and the magnetic field measurement.

The lookup table was generated using The Aerospace Corporation’s SOAP (Satellite Orbit Analysis Program) to obtain the satellite’s trajectory as a function of time. The latitude, longitude, and altitude were imported into an Excel program that interpolated magnetic field components based on grid values calculated using the IGRF (International Geomagnetic Reference Field, 2011). Lookup tables were generated to coincide with desired CTECS run times. Care was taken to schedule runs for orbits where the satellite did not pass very close to the magnetic poles. Near these poles, the X and Y fields as seen by a nadir-pointing spacecraft are near zero, resulting in potentially large attitude errors.


Figure 16: CTECS anti-flight pointing algorithm (image credit: The Aerospace Corporation)

Background: The first GPSRO (GPS Radio Occultation) sensor on a CubeSat was flown on the CanX-2 mission launched in 2008. CanX-2 utilized the NovAtel OEM4-2GL receiver and a COTS antenna. It had a dual purpose of providing position/navigation/time data to the spacecraft as well as observing occulting GPS satellites. Kahr et al. showed that the GPSRO sensor was able to track occulting GPS satellites. 4) However, they have been unable to extract TEC from the measurements due to issues related to power and antenna pointing.

Since CanX-2, a number of different groups are developing GPSRO sensors that can be flown on CubeSats. One of those efforts is on-going at The Aerospace Corporation. The CTECS Total Electron Content Sensor) utilizes the next generation NovAtel receiver, OEMV-2, and a custom antenna to obtain ionospheric TEC data. 5)


CTECS instrument:

CTECS consists of the NovAtel OEMV-2 receiver, a custom antenna, connecting RF cable, and a mounting bracket.

- 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 and 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.


Figure 17: Photo of the CTECS sensor (the antenna is mounted in the bracket facing into the page), image credit: The Aerospace Corporation

- 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 experiment required that the host satellite oriented the GPSRO antenna into the anti-ram direction throughout the data collection period. Furthermore, orientation knowledge was required to assist the processing of the CTECS data.


Figure 18: Topside of CTECS sensor assembly showing the antenna (image credit: The Aerospace Corporation, Ref. 5)


Antenna pointing: The satellite had two mildly directional (4 dBi gain) antennas on the +Y and -Y faces for communication to the ground station in El Segundo, CA. By taking advantage of the CTECS anti-flight LVLH pointing algorithm and the 52º inclination of the orbit, either face could be roughly pointed towards the ground station for increased data throughput. For example, if the satellite was on a northbound pass west of El Segundo, the satellite could point the +X face in the anti-flight direction, thus pointing the -Y antenna towards Los Angeles. If the +Y antenna was desired, the satellite could point the +X face in the flight direction.


Figure 19: Photo of the PSSCT-2 nanosatellite (image credit: The Aerospace Corporation)



On-orbit performance:

Ground operations: Ground operations for the satellite were run through a single ground station located at the El Segundo campus of The Aerospace Corporation. Our 5 m parabolic antenna provides +30 dB gain at 915 MHz and has been used for previous small satellite missions to provide an average per-pass data download of 250 kB. The PSSCT-2 orbit provided three to four ground station passes per day, each with a usable communications duration between five and ten minutes.

On-orbit checkout: Initial checkout of the PSSCT-2 spacecraft included downloading the satellite’s state-of-health (SOH) telemetry followed by a functionality checkout of all subsystems. Once routine tracking and communication were established, downlinking the photographs of the Space Shuttle Atlantis, automatically snapped by a pair of rear-facing 2 Mpixel cameras (one 185º and the other 55º FOV, became the priority due to the historical nature of these images. A minor communication issue specific to this camera board slowed this process and consumed many overflights and therefore many mission days. Downloading the images and troubleshooting RF-interference problems with the satellite’s secondary radio system consumed the first month of the mission.

Nadir and anti-flight pointing: The checkout period was followed by uploading various attitude control algorithms for testing. The reaction wheels, torque coils, and attitude sensors were originally intended for another spacecraft with a different geometric configuration and coordinate system, and coordinate translations were different for each sensor and actuator. Due to the short delivery timeline, few of the algorithms had been tested on the flight hardware and even the polarities of the reaction wheels and torque coils had to be re-verified on-orbit.

The major challenge of the mission quickly became evident: it was time-consuming to test attitude control algorithms on-orbit with a single ground station. An additional source of confusion came from the fact that two of the IR thermometers used in the Earth nadir sensor array had compromised hermetic seals, resulting in significantly different response characteristics which confused the control algorithm. Once we confirmed and corrected this anomaly, the project was able to achieve Bdot detumbling, followed by Earth-acquisition, nadir-tracking, and magnetic-pointing of the CTECS antenna in the anti-flight direction. However, the satellite was only able to maintain proper orientation for 15-25 minutes before the attitude control system would lose its control authority and attitude errors would escalate, as shown in Figure 20.


Figure 20: Attitude angle errors while nadir-tracking and magnetic-pointing showing the loss of attitude control after 15-20 minutes (image credit: The Aerospace Corporation)

By measuring satellite rotation rates entering and exiting from the control algorithm, the project observed a buildup of angular momentum. The increase in angular momentum could have been due to current loops in the reaction wheel circuitry, but was most likely caused by Eddy currents generated in our high-speed electrically-conducting reaction wheels. As seen by the wheels, the Earth’s magnetic field rotates about them at up to 60,000 rpm. The rate-of-change of magnetic field in this reference frame generates current loops within the wheel that resistively decay. The Earth’s magnetic field causes a weak drag on the wheel, which results in a transfer of angular momentum. To make matters worse, these wheels were machined out of mu-metal that causes the local magnetic field lines to locally concentrate, thus increasing Eddy currents and rotational drag by one or more orders-of-magnitude.

Although small, typically less than 5o per second over 30 minutes of pointing, this rotation rate was larger than the control authority of our reaction wheels. In order to meet CTECS mission pointing requirements, the project had to maintain pointing for four hours. This problem was solved by adding intermittent B-dot magnetic detumbling in the control algorithm, allowing the spacecraft to magnetically unload inertia for 10 s out of every 120. By detumbling so often, the angular momentum never grew to a detrimental amount, and the satellite did not drift more than 10º during this uncontrolled time-out from active pointing. After detumbling for 10 s, the control algorithm would re-orient the spacecraft and continue to point the spacecraft. This new algorithm allowed the project to maintain the ±15º pointing accuracy required for the CTECS mission over four hours.

Pointing error data from a two-hour CTECS run are shown in Figure 21, demonstrating the capability to maintain ±15º pointing for nearly the entire duration. Small periodic perturbations can be seen in the X-axis and Y-axis pointing angles where the satellite is performing intermittent magnetic detumbling. Periods where the Z-axis pointing errors are large correspond to periods in the orbit where the Earth’s magnetic field predominantly goes through the Z-axis of the spacecraft; regions near the Earth’s magnetic poles.


Figure 21: Attitude angle errors while nadir-tracking and magnetic-pointing in LVLH mode during a two-hour CTECS experiment (image credit: The Aerospace Corporation)

The development of nadir and magnetic pointing had significant additional benefits to the mission. During spacecraft and attitude control algorithm checkout periods, the satellite was typically in an initial uncontrolled tumble. This limited the average data downlink to < 250 kB/pass and occasionally resulted in wasted passes where no data was received. The CTECS instrument generated 8 MB of data during a four-hour test, requiring more than ten days to download the results from a tumbling spacecraft. By scheduling the spacecraft to roughly point the communications antenna at the ground station as it passed over the Los Angeles region, the project was able to more than double our average downlink rate and eliminate completely-wasted passes.

Additionally, nadir and anti-flight pointing allowed the spacecraft to take multiple, aligned, medium-resolution photos of a ground-track that could be joined together to form a larger effective image. Finally, this capability allowed us to orient the spacecraft in a proper attitude for propulsive orbit-raising.

Orbit-raising motor firing: The PSSCT-2 nanosatellite had four small solid rocket motors, with propellant grains from an Aerotech E28T motor, to demonstrate orbit-raising capability. In order to increase altitude, the motors had to be fired along the trajectory of the satellite, and in order to meet NASA safety guidelines, the rocket firing command had to be sent manually while in contact with the ground station. In order to achieve orbit-raising, the satellite was placed in a nadir-pointing LVLH orientation prior to the selected rocket-firing pass and then commanded to pitch-up to the desired orientation during the pass before the rocket-firing command was sent. Figure 22 shows a photo taken from one the satellite’s cameras prior to sending the firing command to verify proper orientation. This fisheye image from a camera on the +Z face shows the highest point of Earth’s horizon near the middle of the image, thus indicating that the spacecraft +Z axis is roughly horizontal to the Earth. The azimuth angle was determined by identifying the visible landmasses on the left side of the image and comparing them to what the +Z axis should have seen using SOAP. In this case, the satellite orientation was correct in local elevation and azimuth to within 10º. The thruster was fired.


Figure 22: Image taken by PSSCT-2 to verify proper orientation prior to sending the rocket-firing command (image credit: The Aerospace Corporation)

The first solid motor was fired on November 4, 2011, successfully raising apogee of the spacecraft by 10 km. The apogee increase should have been 4 times larger, but alignment of the rocket’s thrust vector through the satellite’s center-of-gravity was off by several millimeters even though the project carefully aligned it on the ground. The resulting torque caused an almost complete rotation during the 1.5 s long burn, and a 360º/s spin. This rotation rate was well beyond the initial capability of the B-dot algorithm to detumble the spacecraft. Control authority was regained by shortening the algorithm’s loop time, and the satellite resumed its mission.

Over the next month and a half, the remaining three solid motors were commanded to fire, however while telemetry indicated that each of the igniters fired, none of the remaining solid motors ignited. It is unknown whether the failure was due to the prolonged exposure of the solid propellant to vacuum, or whether the initial firing had an adverse effect on the remaining motors.

Solar cell performance data: PSSCT-2’s sun-tracking capability was developed with lessons learned from nadir pointing. While the nadir-pointing algorithm of the spacecraft resulted in an oscillation of ± 5º of nadir-pointing, which was sufficient for the CTECS experiment, high-fidelity IV curves required the satellite to maintain steady pointing, orienting a specific face normal to the sun but for a much shorter period of time. An improved control algorithm was written using a semi-proportional control scheme instead of the previous limit cycle (bang-bang) approach. By empirically generating an accurate equation for the commanded motor duration and the resulting satellite motion, a more accurate pointing algorithm was achieved resulting in the ability to point one of the spacecraft’s faces normal to the sun and hold the orientation for a few minutes, which was long enough to take solar cell IV data.

Typical results are shown in Figure 23. Due to a limited number of possible scheduled commands, however, IV data had to be obtained while in communication with the ground station in El Segundo, and because of the time of year and the inclination of the orbit, even at local noon there was some contribution of Earth's albedo in the solar cell performance measurements, making direct comparison of the solar cell performance on-orbit difficult. Being able to schedule an IV measurement at an appropriate point in the orbit would eliminate albedo interference.


Figure 23: Sun sensor readings while tracking the sun (image credit: The Aerospace Corporation)


Lessons learned:

A key feature of the PSSCT-2 and all future Aerospace miniature satellites is on-orbit reprogrammability. This reduces the pressure on the development schedule as long as the software code structure for reprogramming is in place and as long as the basic functions or drivers have been tested. Once on orbit, new software can be uploaded and tested. This all worked as planned for the PSSCT-2 attitude control processor with one major exception – time. The first lesson learned is the time needed to fix attitude control algorithm software on-orbit. Ground testing for those functions that can be tested in a 1-G environment is much faster than on-orbit testing due to the ability to use visible sensors, like your eyes, and not having to wait for data downloads from the next pass. The PICOSAT team was notified in September 2010 that if they could deliver a satellite to the space shuttle program by June of 2011, then a flight was possible. No attitude control test fixtures existed and no time was available to make them – the satellite sun and Earth sensors, reaction wheels and torque coils were tested individually but not in concert.

Space can be the perfect place to develop and test attitude control algorithms, but it turned out to be a time consuming effort. The PSSCT-2 overflew the El Segundo, CA ground station about three times per day. Table 1 lists the steps for one iteration of algorithm refinement. It requires one day minimum to run one iteration under ideal circumstances and often this is doubled due to poor communication passes produced by poor satellite orientation. If engineers work a normal 5-day week, there can only be 2-3 iterations on the ACS algorithm per week.

The project experience with PSSCT-2 was that it required two months to achieve sufficient attitude control to run the primary experiment. The lesson learned was two-fold regarding certification and repair of an attitude control algorithm. First, it is quicker to test the ACS algorithm in the lab than in space. We now use a simple single degree-of-freedom hanging string to test the ACS for our new AeroCube-4 satellites. Anomalies are still time consuming to resolve, but the resolution is quicker than experienced with PSSCT-2. Second, if another ground station can be added that is far enough away to not share the same satellite footprint as the primary, then the ACS algorithm test-replace cycle can be almost twice as fast.

Notional pass time


12:00 pm

Upload code; upload schedule

1:30 pm

Continue upload code; upload schedule (if needed)

1:30 pm – 12:00 am

Test runs automatically

12:00 am

Download results

1:30 am

Download results

8:00 am – 2:00 pm

Interpret results

2:00 pm – 11:30 am

Receive code

Table 1: PSSCT-2 ACS algorithm refinement cycle timeline

A second lesson learned is the benefit of having visible cameras on a satellite. The adage that, “A picture is worth a thousand words” is an understatement when trying to understand attitude control sensor data from hundred’s of kilometers away. The CTECS experiment required one end of the PSSCT-2 to face nadir for four hours while the satellite circled the Earth. Reaction wheels continuously modified the nanosatellite orientation based on feedback from the staring Earth nadir sensor. Prior to the actual 4 hour run, many 30 minute nadir orientation experiments were done to prove out the attitude control system before significant experiment data were collected. During test runs, the Earth nadir sensor and rate gyro sensor data were collected. Also, pictures were snapped periodically from a camera on the +Z face. The pictures made understanding the collected data easy, thereby shortening the debugging process of the attitude control algorithm.

A third lesson learned is about testing using a realistic communication link. The PSSCT-2 satellite was reprogrammable on-orbit. This feature was tested prior to shipping but the test conditions did not exactly mimic the flight conditions. During bench testing, a short range communication link was established with the satellite and various satellite processors were reprogrammed to verify this capability. During flight, the communication link was degraded by terrestrial radio frequency interference and non-optimum satellite orientation. If the entire software upload for the CTECS payload code was not transmitted in one uninterrupted transfer, the upload would fail because the communication link timed out. This greatly limited the opportunities for such an upload and wasted useful mission life. If ground testing had included variable radio frequency attenuation with added dropouts, new programming tools like restarts and partial file uploads would have been added. The lesson is that real communication links should be simulated with variable attenuators to observe how communication software will really work.

A fourth lesson learned has to do with moments of inertia. The project assumed that if our satellite had reaction wheels aligned along the geometric satellite axes, the spacecraft rotation about one axis would be decoupled from the other axes. However, the principal and geometric axes were not perfectly aligned, resulting in some cross-talk between axes as each reaction wheel spun up. The PSSCT-2 used sun sensors on two opposing sides to prove that the sun was normal to the solar cells on those faces during a measurement of cell performance. However, the solar cells on adjacent faces were indexed towards the sun using a simple open loop 90º rotation. This was not the case; turning 90º caused a significant compound off-normal angle that degraded solar cell performance measurements on those faces. Future solar cell monitoring experiments should have a sun sensor on each relevant surface, or at least use rate gyro data to provide closed-loop feedback.



The PSSCT-2 mission consisted of three main objectives:

• The CTECS instrument measured GPS radio occultations for the first time in a nanosatellite platform. CTECS collected 13.5 hours of data. Utilizing modified software and a custom designed antenna, CTECS has successfully collected occultation data. It demonstrated the ability to obtain relative TEC profiles showing geophysically meaningful features. For the first time TEC and electron density profiles have been obtained from a GPSRO sensor hosted on a nanosatellite (Ref. 5).

• The solar cell characterization task measured the true in-space AM0 performance of vendor-provided coverglass-interconnect-cells.

• The thruster experiment successfully changed orbit apogee by 10 km, but induced a high spacecraft spin rate.

The sun and Earth nadir sensors enabled 1º attitude knowledge when the appropriate celestial objects were visible, and the attitude control algorithms enabled pointing with at least ±15º accuracy over four hours of operation. This performance was achieved after 3 months of on-orbit testing and algorithm refinement. - Unfortunately, the orbital lifetime was only 4.5 months - when the PSSCT-2 nanosatellite reentered Earth's atmosphere on Dec. 8, 2011.


Figure 24: Trajectory of the PSSCT-2 nanosatellite after release from the Space Shuttle (image credit: The Aerospace Corporation)

PSST-2 mission release events:

• Released into a circular orbit at initial altitude of 380 km.

• Orbit decayed in 4.5 months and re-entry occurred 8 December 2011.

• On-board rockets were fired, temporarily raising orbit.

• Attitude control provided 3-axis stability (1°) and magnetic alignment

- Sun and earth sensors

- Magnetometer

- Torque coils

- Reaction Wheels

- Angular rate sensor

• Downlink data rate: 115.2 kbit/s.


Figure 25: Data example of relative TEC calculated from tracking GPS satellite from zenith to below the horizon (image credit: The Aerospace Corporation)

Legend to Figure 25:

• The yellow region indicates the occultation portion of the track

• In this example, entire track is 30.6 minutes long (occultation takes 6 minutes)

• Because of low altitude of orbit, CTECS was able to observe large a amount of density above the satellite (Appleton anomaly observed).

1) Siegfried W. Janson, Brian S. Hardy, Andrew Y. Chin, Daniel L. Rumsey, Daniel A. Ehrlich, David A. Hinkley, “Attitude Control on the Pico Satellite Solar Cell Testbed-2,” Proceedings of the 26th Annual AIAA/USU Conference on Small Satellites, Logan, Utah, USA, August 13-16, 2012, paper: SSC12-II-1

2) Laura Johnson, “Aerospace Plays Big Role with Small Satellites,” The Aerospace Corporation, March 15, 2012, URL:

3) “Final Nanosatellite Launched from Space Shuttle Atlantis,” Space Daily, July 20, 2011, URL:

4) E. Kahr, O. Montenbruck, K. O’Keefe, S. Skone, J. Urbanek, L. Bradbury, P. Fenton, “GPS Tracking on a Nanosatellite – The CanX-2 Flight Experience”, Proceedings of the 85h International ESA Conference on Guidance, Navigation & Control Systems, Karlovy Vary, Czech Republic, June 2011

5) Rebecca L. Bishop, David A. Hinkley, Daniel R. Stoffel, David E. Ping, Paul R. Straus, Timothy R. Brubaker, “First Results From the GPS Compact Total Electron Content Sensor (CTECS) on the PSSCT-2 Nanosat,” Proceedings of the 26th Annual AIAA/USU Conference on Small Satellites, Logan, Utah, USA, August 13-16, 2012, paper: SSC12-XI-2, URL of presentation:

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.