The technologies developed during this project will conform to the cube-sat standards, such that they can easily be used on future missions as well. Figure 1 is a product tree for the nano-satellite to be developed within the PEASSS project. The boxes with the bold blue lines are the parts that are developed within this project, the boxes with the black lines are parts that are already available from other project. These parts are combined with the new developed parts to form the PEASSS Nano-Satellite.
The main objective of PEASSS can be translated into four sub-goals:
Demonstration and qualification in space environment of a piezoelectric actuated "small structure" as a means of pointing an optical instrument, while correcting for thermal deformation and laying the groundwork for cancelling acoustic noise from the satellite, to achieve better accuracy than current technology, with lower mass, power use, and/or reaction time.
Demonstration and qualification in space environment of a piezoelectric actuated "smart structure" as a means of power generation from the pyroelectric effect, capable of generating >1Wlm2.
Demonstration and qualification in space environment of Fiber Bragg Gratings (FBG's) combined with a miniaturized interrogator, in order to measure composite structure strain for structure actuation control and temperature with 0.3" accuracy.
Demonstration and qualification in space environment of next generation power conditioning and data acquisition components for Nano-satellites by integrating new energy scavenging methods and accommodating distributed sensor networks and novel data gathering techniques.
The first two objectives use the piezoelectric technology to demonstrate precision actuation and power generation. in Space, while the last two objectives are aimed at measuring and exploring their performance. The work plan will follow the proven flow for development of space technology, to research and develop specific functional components for the proposed satellite, in order to maximize the advancement of the Technology Readiness Level (TRL) of the technologies. The technical part of the work plan is divided into 6 technical work packages: WP1 up to WP6.
WP1 consists of Systems Research & Engineering, which will develop and monitor specifications and interfaces such that the developed components come together as a fully functional system.
WP2 consists of component Design Development. Under this work package, the requirements specified under WP1 will be met by designs for the system components, where necessary preceded by a trade-off study.
WP3 Consists of Breadboard Development and Testing. ln this Work Package, the designs that were specified for each of the components in WP3 will be produced in breadboard format for functional development and testing. Again, this WP will focus on breadboard testing the nanosat electronics, power generation, piezo-actuated structure and FBG sensors and interrogator.
WP4 is the component Manufacturing, Assembly, integration and Testing (MAIT). Here the components will evolve from breadboard models to flight-test ready hardware and related software. By the end of WPs, the nanosat electronics, power generation, smart structure, and FBG sensor system will be fully developed and tested, and ready to be integrated into a working satellite.
WP5 consists of satellite integration and functional ground testing.
WP6 is space environment testing of the satellite delivered after WP6. This includes space environment testing analysis and design, space environment operations and test readiness, and space environment testing.
PEASSS Bread Boarding
The bread boarding of the hardware includes nanosat electronics, power generation, piezo-actuated structure and FBG sensors and interrogator. In the bread boarding phase the individual components is functionally tested. Also some Thermal Vacuum testing on very temperature sensitive components has been performed successfully.
thermal vacuum test on interrogator, including the model for thermal simulation.
The CubeSat platform has been chosen due to its relatively low-cost simplicity, usage of COTS, standardized parts. For typical CubeSats missions, the space qualified parts or qualification campaigns for space applications are not affordable due to budget issues resulting in increased risks. Therefore design rules similar to those for “bigger” space projects have been applied, as qualification/acceptance test campaigns, trade-offs, mission scenarios analysis, thermal vacuum campaigns, including thermal modelling and simulation.
The thermal control subsystem, as one of the basic spacecraft subsystems, is responsible for providing appropriate thermal conditions within the satellite for the components to operate. These conditions are defined by the requirements of each component, e.g. typical electronics boards are able to operate between -20 and +70 °C. From a thermal point of view, the most demanding elements of PEASSS mission are the payload components, as these elements are mostly designed for on-ground applications and using COTS, without previous space heritage.
The Breadboard phase has come to a successful conclusion with the vibration test for qualification of the PEASSS Nano-Satellite EQM Model. Following this successful conclusion the consortium started manufacturing the components for the Flight Model (FM). On the pictures below the components of both the satellite and the payloads are shown.
Figure 1, Flight Model of the pyroelectric power generator consisting of two sets of piezo elements covered with black foil for optimal heat transfer
Figure 2, frontside of the FM Optical Bench with the two sun sensors visible. The right sun sensor mounted on the gimbaled ring-disc structure that can be tilted by the two piezo bimorph actuators that are located at the backside.
Figure 3, backside of the FM Optical Bench with the two piezo bimorph actuators visible that can tilt the sun sensor on the gimbaled ring-disc structure. On the backside are also the optical fibers with the Fiber Bragg Gratings (FBG’s) located.
Figure 4, Flight Model of the interrogator mounted on the electronics plate between high voltage amplifiers. The components are mounted in a test frame.
Figure 5, solar panels of the PEASSS satellite
All these Flight Model components have been individually functional tested and have past all success criteria. In this phase the development of the software both for the satellite platform and for the payloads was also started. After the manufacturing phase, the assembly of the components was started. The development of the software continued, followed by extensive testing.
Figure 6, 'flat sat' testing before mechanical integration
Then after successful testing of the software the integration was performed. Finally after integration the environmental testing of the Nano-Satellite was started.
Figure 7, integration of the payload into the PEASSS satellite structure.
The top panel of the satellite is the Smart Structure where a sun sensor can be pointed by means of two piezo beams. Both the tilting sun sensor and a second sun sensor which is used for reference are protected by a black cover with PEASSS inscription.
Figure 8, The fully integrated PEASSS satellite in a thermal chamber for the “thermal bake-out” as required by the launcher
Figure 9, The PEASSS satellite mounted in a test pod on the shaker for vibration testing
The testing of the Flight Model consisted of functional testing of the integrated satellite and environmental tests. The functional testing of the satellite is performed before and after every environmental test to ensure that nothing has changed during the test. The environmental tests are a vibration test at proto-flight level, a shock test at prototype level and a thermal bake-out as required by the launcher
Figure 10, The Flight Model of the PEASSS satellite before integration in the launch-pod
On Wednesday 15 February 2017 at 3:58 UTC the PEASSS nano satellite or Cube Sat. was launched successful into space. The launch was performed in India by the PSLV (Polar Satellite Launch Vehicle) mission C37.
Pictures: Department of Space Indian Space Research Organization (ISRO)
During this mission a record number of 104 (!) nano-sats have been brought to an orbital altitude of 550 km; once the launcher was in orbit, a nano-sat was released every 10 seconds.
The launch was very successful: the PEASSS satellite is in the right orbit, very slowly tumbling and contact with the satellite was already established a few hours after launch during the first orbit over the ground station in Delft (the Netherlands).
PEASSS is a relatively small satellite (10 x 10 x 30 cm) and only weighs 4 kg. The satellite is developed and manufactured in the EU FP7 program led by TNO and in collaboration with five partners: ISIS (Netherlands), NSL (Israel), Sonaca Space Group (Germany), Sonaca (Belgium) and The Technion (Israel). Actually this is the first time that a satellite was launched as part of an EU program.
The ground station for the satellite is located at ISIS in Delft and communication with the satellite to send commands and to receive data from the on-board experiments is possible 3 times per day, during a period of 10 minutes.
The original planning for the project was to start on the 1st of January 2013 and to finish on the 31st of December 2015. It was intended to have a successful launch within this period of time or to perform tests on a fully integrated satellite on Earth in a space representative environment.
The timing of the launch was an issue that needed to be solved during the project. Originally the project was scheduled to finish on the 31st of December 2015. However it was not possible for the consortium to find a suitable launch window before this date. This issue is extensively discussed with the EC project officer and together a very workable solution was found: suspension of the project until a suitable launch window was selected. This was combined with an extension of the project duration of 5 months to enable the consortium to deliver a fully functional, fully tested satellite to the launch.
On the 1st of January 2016 the project was suspended according to plan to cover the time until the launch of the satellite that would carry the PEASSS satellite into space. On the 1st of December 2016 the project was restarted again in order to have time to deliver the PEASSS satellite to the launch provider. During this suspension period it was not possible to get labor funded by the EU, but nevertheless the project partners kept on working on the PEASSS satellite at their own expenses.
The PEASSS satellite was launched on the 15th of February 2017 and from that time until the end of the project experiments were performed in space and data was downloaded and analyzed.
Now the EU project ends in April 2017, while the expected life time of the Cube Sat is about 5 to 7 years. After this period it will fall back to Earth and burnup in the atmosphere. The consortium will follow the Cube Sat after the ending of the project in April 2017 and will keep on conducting experiments with both the platform and the payloads. The goal of this activity goes well beyond the goal of the project.
PEASSS Mission Patch
In Orbit Testing
The PEASSS satellite was successfully launched into space on Wednesday 15 February 2017 at 3:58 UTC. Contact was already established a few hours after launch during the first orbit over the ground station located in Delft (the Netherlands). From that moment on data has been gathered onboard the satellite and downloaded to the ground station. Due to the orbit around the earth communication with the ground station located in Delft is only possible three times a day during a period of a about 10 minutes.
Figure 11, location of the PEASSS satellite during experiment number 14, also showing the location of the ground station in Delft (the Netherlands). It is visible that the satellite is at the end of the sunlit period just before entering eclipse.
Before payload experiments could be started, first the correct operation of the satellite had to be established during LEOP (Launch and Early Operations Phase), followed by Platform commissioning and Payload commissioning. A critical aspect of Payload Commissioning was the release of the launch-lock of the Optical Bench. This launch-lock prevented damaging of the delicate mechanical parts of the Optical Bench resulting from the violent vibrations during launch. The launch lock consists of a Shape Memory Alloy (SMA) beam that returns to its ‘programmed’ shape after heating. Comparing measurements from a tilting experiment of the Optical Bench, performed before and after opening the lock, showed that the lock had opened as planned and that it had protected the delicate mechanics during transport and launch.
The short time between satellite launch and the end of the PEASSS project meant that only a limited number of experiments could be performed and analyzed. The main limitation is the relative low effective download bandwidth resulting from only 3 overhead passes of about 10 minutes each day. Communication is only possible during these 3 passes. Of course the PEASSS satellite will stay operational also after the end of the PEASSS project and it is expected that regularly data will be downloaded and analyzed to observe performance over time.
Interrogator and Fiber Bragg Gratings (FBG’s)
The limited effective download bandwidth was mainly a problem for downloading the data from the interrogator which was used for reading out the Fiber Bragg Gratings (FBG) used for strain- and temperature measurement. Due to the experimental state of the system it was decided not to integrate the algorithms that perform this analysis into the satellite, but to download the raw data and perform this analysis on the ground. This provides the possibility to tune the algorithms but the drawback is that relative large amounts of data have to be transferred from the satellite to the ground. The limited effective bandwidth available in combination with only a limited time per cycle that the satellite is in view of the ground station makes that downloading the complete data set of an experiment with multiple tilt angles takes several days. The result is that in the short time between the launch and the end of the project the data of only a limited number experiments could be downloaded and analysed.
Piezoelectric Actuation Mechanism
An experiment consists of changing the attitude of a sun sensor mounted on a gimballed structure that can be tilted by means of two orthogonally mounted piezo bimorph actuators. A sun sensor measures the attitude of the sun with respect to the body of the sun sensor. By applying different voltage levels to the bimorphs, different tilting angles can be obtained. The bimorphs can be actuated independently which makes it possible to generate a 2D array of tilting angles. In the figure an example pattern with discrete voltage level combinations is presented that is used in an experiment to direct the sun sensor to different pointing angles. Even this limited array consist already of 41 different tilt angle combinations. During the limited time available data from 2 experiments with 5 tilting angle steps was downloaded and from 1 full experiment with 41 tilt angle combinations.
Figure 12, example of 2D array of possible actuation voltages that result in tilting angles of the sun sensor mounted on the gimballed structure
It was demonstrated that a sun sensor could be tilted in two orthogonal directions by means of two piezo bimorph actuators. Feedback of the tilting angle is possible by means of strain gauges glued on top of these bimorph actuators. Strain was measured by means of classic resistive strain gauges but also by means of Fibre Bragg Gratings (FBG’s). The sensitivity of FBG’s to temperature can be cancelled out by mounting an FBG both on the top- and on the bottom side of the bimorph beam. The use of strain measurement on both top- and bottom sides of the bimorph also proved to be necessary to obtain a good measure for the bending of the bimorph and with that for the tilting angle.
Figure 13, for 1 rotation axis from top to bottom the applied voltage to the piezo bimorph actuator during the tilting steps, the strain measured by the differential FBG strain gauge and the resistive strain gauge (SG) and the tilting angle measured by means of the sun sensors.
Pyroelectric Power Generator
For analyses of the Pyroelectric power generation payload experiment only data with a relatively low sample rate is needed to observe the effect of temperature changes. These temperature changes are the result of the satellite entering the shadow of the earth (eclipse) each orbit of about 1.5 hours. A faster temperature change of the piezo power generator may result from the tumbling of the satellite. The temperature change of the Pyroelectric Power Generator is however also related to the attitude of the satellite. An example of the measured temperature change of the Pyroelectric power generator and the resulting output voltage is presented in the next figure. The voltage is measured over a 200 kohm load resistor and therefore represents generated power as well. The power output level is relatively low which is the result of a rather slow temperature change. It is expected that a higher power output level can also be obtained by further optimization of the design. As far as known this is however the first time that pyroelectric power generation is demonstrated in space.
Figure 14, temperature of the pyroelectric power generator resulting from the eclipse cycle and the resulting generated output voltage measured over a load resistor.
The temperature measurements performed in space were used for thermal analysis. The measurements were compared with thermal models. However due to the limited effective download bandwidth the data available for thermal analysis comparison purposes allowed only a comparison for external solar panels and a limited number of internal subsystems of the satellite (battery, transceiver and iMTQ). The pictures show both the simulated solar panel temperatures for cold case 01 and the measured solar panel temperatures. The temperatures’ profile on the panels is comparable with the analysis outcome. In the cold parts of the orbit (eclipse) the temperatures of the panels tend to equilibrate since no different radiative loads from external sources are present on the different sides of the satellite. The behaviour of each side is different once the satellite is exposed to the Sun. This is due to the different spinning rate of the satellite around each axis.
Figure 15, simulated solar panel temperatures for the Cold case 01.
Figure 16, solar panel temperatures measured in space.