ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
103 The Center shaft design is less straightforward as one would expect due to the thermal environment  requirements and the variation in thermal coefficient of the damper stainless steel input shaft and the aluminum 7075 Center shaft. The damper has a steel input shaft unlike the Center shaft, the main bracket and the cable pulley, which are of aluminum. An additional stainless steel shaft, the damper engagement shaft is introduced to interface with the damper input shaft and the Center shaft. The damper engagement shaft is coupled with the Center shaft by the lock bolt shown in Figure 10.   

Figure 10:  Damper Spring Unit shaft design.     Qualification program   The qualification program consisted of various functional tests at ambient and at the extreme temperatures (-115 ˚C / +115 ˚C), static load tests, random- and sine-vibration tests and a life test. As the  damper has been qualified by RUAG, not all tests are performed with the damper attached to the main bracket. This allows a full characterization of the DSU, which is required to create the ADAMS model. Figure 11 gives the test sequence of the DSU, which indicates when the damper was not used.    

Figure 11: DSU qualification test sequence.
104  

Figure 12: Functional test of the DSU qualification model (ambient). 

Random vibration tests have been done and the test requirements of Table 1 and 2 have been met.  

Table 1: Random Vibration Test requirements (all 3 axes)  

Frequency Range [Hz] PSD Qualification  10 – 80 + 6 dB/oct until 0.6 g²/Hz  80 – 400 0.6 g²/Hz  400 – 2000 - 6 dB/oct  Overall random vibration level [G RMS ] 20  Duration [seconds] 120 

Table 2: Sine Vibration Test requirements (all 3 axes)  

Frequency Range [Hz] g-level  5– 22 Maximum shaker amplitude  22 – 100 30  

During the functional tests, the behavior proved to be very sensitive for the alignment of the DSU. This was therefore monitored carefully during the qualification program, especially prior to the TV cold testing. During TV “Cold” testing the retarding torque exceeded the max. required value: 284 N-mm versus the required 247 N-mm. This requirement had been set conservative and fortunately the torque balance analysis shows that the 284 N-mm is acceptable. Because this Non Conformance has been resolved and all other pass/fail criteria are met, like the ultimate load of cable A of 3360 N, the DSU qualification test program has been successful and the DSU design is qualified for use within the specified environment (temperatures, forces and number of on ground cycles).
105 Torque Balance Analysis 

The torque balance analysis presented here shows the torque balance for the critical situation of the  deployed position during cold deployment. In the cold condition the friction coefficients and the damper retarding torque are at their highest, and the deployed position has the lowest available DSU drive spring torque. Figure 13 shows the torque balance calculation flow and Figure 14 shows which cable forces and deflection angles are taken into account. The torque balance budget for the DSU and its control cables, given in Table 3, shows that the DSU motoring torque is a factor 3.3 higher than the worst case retarding torque. A motoring factor (torque margin) >3 is required, so this requirement is met. The DSU spring unit can provide up to 7400 N-m, consequently a higher motoring factor can be achieved if so desired.    

Figure 13: Torque balance calculation flow.    

Figure 14: DSU cable forces and deflection angle definitions.
106 Table 3: The DSU system torque balance budget.  

Sentinel 1 Deployment Simulation in Adams 

The deployment of the Sentinel-1 solar array is studied with the commercial multi-body package ADAMS which is part of the mechanical engineering software of MSC. In ADAMS, a model of a solar array wing is characterized by Parts, Joints and Loads. The ADAMS Sentinel 1 model is shown in Figure 15.  The parts in the model description are:   Panels: the yoke panel and the solar panels.   Synchronization system: pulleys, guide-blocks and begin and end parts of the synchronization  cable.   The damper unit    The constraints in a solar array wing model define:    revolute joints (or hinges) between the panels   Constraints relating two markers that together define a marker for the (virtual) synchro-cable (The  contributing load is implemented and a graphical element. No actual body is representing the cable)     The loads in the model are:   Deployment torques, consisting of  o Torques to deploy the stack of panels  o Kick-off spring which pushes the complete stack during the first 60 degrees of the root  hinge   Non-linear loads to prevent that the hinges between the panels rotate beyond the structural  boundaries (The panels are not able to cross each other)   Linear loads activated when the hinge locks (=at the end of the hinge deployment)   The damper load applied by the damper   The loads in the synchronization cables   COLD  CASE DEPLOYMENT Torque Safety Factored factor torque [Nmm] [-] [Nmm] RETARDING TORQUE DSU DEPLOYED damper start torque -1240 3 -3720 Resistiv e torque of the DSU shaft + pulleys -284 3 -852 TOTAL retarding -1524 3 -4572 TOTAL motoring torque DSU springs 5100 1 5100 Motorization factor 3.3
107  

Figure 15: The Sentinel 1 SA wing representation in ADAMS.  

The functional components are modeled in ADAMS as User Defined Entities (UDE’s). Since the functional components are fully parameterized the UDE’s can be re-used in a very efficient way. The actuation springs are straightforwardly represented by moments in the hinges. The characteristic moment – angle behavior is introduced as a prescribed curve. Unfortunately, in ADAMS cables and pulleys are not standardly available and therefore a toolbox developed by Sayfield International has been used. With this toolbox the synchronization system has also been modeled as a UDE with help of sub-models of the pulleys, slip-rings and cables from the toolbox. In these sub-models, the cables are not physically introduced but as forces. The spring damper unit is also modeled with the cable toolbox. Figure 16 shows the complete DSU system: the damper disk, where the damping load is applied, the outer disk, the cables and the pulley. Only the spring unit of the DSU is not shown. The spring is modeled by a moment as a function of the rotation angle of the outer disk. 

Figure 16: The DSU representation in ADAMS.
108 The two stage deployment of Sentinel 1 is shown in Figure 17. Such a trajectory is achieved by placing  the pulleys off centered and using different sizes for the pulley radii. This adapted design changes the roll-on and roll off characteristics of the cables such that the wing opens approximately as required. This pulley configuration is shown in Figure 18.   

Figure 17: The two stage deployment.    

Figure 18: The Root Hinge Roll-off and Roll-on pulley configuration 

Figure 19 shows the ideal roll-on and roll off characteristic and the realized characteristic as a function of the root angle. The root hinge pulleys do not display any significant roll-on and roll-off in the first 90 degrees. Because there is no roll-on and roll-off, the next hinge lines are not able to open. The stack opens as a complete package. After about 90 degrees the length of the cables starts to change with the root hinge angle forcing the next hinge lines to follow.   30 30 30
109  

Figure 19: The roll-on / roll-off characteristic.  

The parameters in the ADAMS model are now tuned to the baseline Sentinel 1 design including the flexibility of the synchronization system. The model is used to predict the deployment characteristics under various thermal conditions and provides the following output: forces in the synchro-cables and control cables of the DSU, the torque provided by the DSU, shocks in the hinges when they lock, deployment time and path of the wing and the loads at the satellite interface. The ADAMS model is validated with the deployment test results of all relevant units and consequently the ADAMS results could be used to qualify the use of the DSU within the Sentinal-1solar array (verification of the occurring loads). It showed that the loads and shocks were within specification, the deployment time is within 88 and 284 seconds and deployment path of the Solar Array stays within the allowed deployment volume. Figure 20 shows the panel positions of the SA wing during deployment case 2, which provides maximum forces in the synchronization system.      

Figure 20: The deployment path of Sentinel 1.     Roll-off cable  Roll-on cable
110 Conclusion 

The DSU for the Sentinel-1 solar array was successfully qualified and the flight models are produced (see  Figure 21). Using engineering models proved to be essential to identify the deployment behavior of the Sentinel 1 wing and the required pulley diameter of the DSU. The qualification program showed that the alignment of the DSU needs to be controlled carefully to limit the friction of the control cables. The spring unit of the DSU design is over dimensioned for the Sentinel 1 SA; it can provide 1.5 times more torque than required, allowing application of the DSU in less favorable conditions / deployment systems. Representation of the Sentinel-1 deployment system (including DSU) in ADAMS allowed the actual wing deployment tests to be limited in both complexity and number of tests.    

Figure 21: The DSU flight model.     Acknowledgements 

The presented work has been carried out at Dutch Space B.V. The authors would like to thank Mr. B. Busz, Mr. J. Cremers, Mr. P. Duyster and Mr. T. Konink of Dutch Space B.V. and Chris Verheul of Sayfield for providing data and support.  

References 

1. ESA Requirements and Standards Division.”Space engineering mechanisms.” ECSS-E-ST-33-01C,  March 2009.  2. Ruag Space AG. “Datasheet Eddy Current Damper.” RUAG Space AG, Schaffhauserstrasse 580,  CH-8052, Zurich, Switzerland.
112 transformation) takes place in the range between As (Austenite Start Temperature) and Af (Austenite  Finish Temperature). Both effects are related to the thermoelastic martensitic transformation, which is a diffusionless reversible phase change characterized by a change in the crystal structure [3]. Thus, these characteristics allow SMA to be applicable for force and strain generation, in the case of shape memory, and for mechanical energy storage, in the case of pseudo-elasticity. 

(a)                                                                              (b)  Figure 1. SMA characteristics: (a) microscopic diagram of shape memory effects; (b)  martensite variation with temperature.  The main advantage of the SMA technology for its use in actuators is its great strength in relation to its ultra-low weight, optimizing the mechanical work performed by the device with a minimum mass. Furthermore, this technology has the advantage of being immune to electromagnetic interference, its noiseless actuation, and that it does not require lubrication to work.  The alloys currently available in the market, mainly NiTi based alloy, are limited in their operating temperatures. Arquimea suggests using a new SMA with further capabilities to overcome these temperature limitations. SMARQ material, a novel proprietary Shape Memory Alloy, is able to work in an extended temperature range, with transformation temperatures, tuneable during the manufacturing process, up to +180 ºC. We present in this work the results obtained from the SMA characterization. In order to perform the complete characterisation of the SMA material for its use as actuators in space environment, this document explains the way the tests have been carried out by considering the SMA material as the key element of a SMA actuator.  This approach requires the space characterization of the  complete actuator element in order to identify the performance of the different elements in the space environment. The main objective of these tests is to obtain technical information about the actuator performances. Special attention will be paid to the main advantage of the Arquimea’s SMA (SMARQ):  its
113 higher working range of temperatures with respect to the SMA materials used until now. This temperature  restriction is one of the most critical limitations of the current SMA devices for their application in space missions and other markets. The present project is expected to demonstrate the capabilities of SMARQ to overcome this limitation.   As above-mentioned, in order to perform the complete characterisation of the SMA material for its use in space environment, this document proposes to perform this characterization by considering the SMA material as the key element of a SMA actuator. A block diagram with all the constituent parts of a typical SMA device is shown in Figure 2. It must be taken into account that some of the blocks are basic blocks that must be part of any SMA device and others are optional, being or not a part of the device depending on the application requirements. The tests carried out as part as this work have been focused on the basic configuration of the SMA device, this is, the combination of the SMA Material and the mechanical and electrical interfaces. The characterization of the rest of the elements in space environment can be performed as independent blocks. The performance of the SMA material and the mechanical and electrical interfaces has been tested during the current work, in order to detect possible incompatibilities of the materials involved or mechanical limitations in the SMA actuators in space environment as well as to obtain information about the SMA actuator main parameters and working behavior.  As an application of SMA technology, a Dust Sensor (DS) instrument will be presented in this work. The Dust Sensor instrument of the MEIGA MetNet Mission will perform airborne dust opacity measurements on the Mars surface. The Dust Sensor is designed as a lightweight device (41 g) that performs an active measurement, using back scattering to estimate the concentration of particles in the airborne dust. The unit integrates an infrared (IR) optical active detector (there is an IR emitter and an IR detector) for a spectrographic (discrimination in wavelength) measurement of dust in suspension.  

Figure 2. Block diagram of Arquimea SMA Device.
114   Figure 3. Qualification Model of the Dust Sensor 

The emitter points to the airborne dust and the detector collects the scattering signal produced by the  emitter light when interacting with particles similar in size to the light wavelength. Figure 3 shows the qualification model of the Dust Sensor device. The DS includes an in-flight calibration system based on a reflector stick that allows directing the emitter optical signal directly to the detector. The reflector stick is inserted in the optical path, when the actuator system is commanded to do so, by means of an actuator based on a Shape Memory Alloy. This in-flight calibration system is used to compensate the error in the entire DS acquisition chain. There is a weight budget of 40 grams for the Dust Sensor instrument, due to the small weight budget of the full MEIGA mission. This has greatly affected design decisions such as the lack of an enclosure, reduced support structures, limitation to the number and size of electronic components, and mainly in the actuator design. The mass budget for the actuator was less than  10 grams, which implied a challenge in design. The Dust Sensor device is controlled by the onboard computer of the MetNet Lander. Periodically, according to a preprogrammed schedule, the Dust Sensor device is powered on and a sequence of instructions is executed. Table 1 summarizes the main characteristics of the Dust Sensor device.   A novel Pin Puller mechanism based on SMA technology will be presented. A Pin Puller is a mechanical device in which a pressure cartridge causes a pin or piston to retract inside the structure frame, usually against a side load. In the extended position, the pin or output shaft can be seen to be loaded by a compression spring.   Table 1. Dust Sensor main characteristics. 

Characteristic Value Notes  Mass 41.2 grams   Main dimensions 85x65x20 mm   Power consumption 360 mW nominal 2250 mW peak consumption during 750 ms  max.  Voltage operation 5 V   Communications 422 serial comms Command oriented instrument.  Temperature accuracy ±1.5°C   Reflector position accuracy ±7.5°   Operational temperature range -90 °C to +25 °C Also in vacuum   The pin remains firmly locked in this position due to mechanical components that block the stroke. Once actuated, however, the actuator drives specific mechanical components releasing the stroke and allowing the pin to retract under the force of the drive spring. The Pin Puller is reset by manually moving the pin back into the extended position. This is done by either pulling it out from the top or pushing it from the bottom. The Pin Puller acts as a trigger for deployment mechanisms. The inertia force of the Solar Array
115 panels, antennas, booms, aperture door covers, etc. will be reacted through the release mechanism –  through the Pin Puller in the shear direction across the pin and the housing to the structure. Pin pullers are normally used in pairs for redundancy. Applications include “Hold Down and Release” of numerous satellite deployable including solar panels, communication antennas, instrument cover doors, radiators, heat shields, tether experiments, and isolation systems. SMA Pin Pullers provide a number of advantages compared to traditional pyro devices:  (1) There is no (or minimum) shock associated with SMA devices (note that there have been reported instances when pyro pullers have tripped relays when fired). (2) The number of safety personnel providing oversight during the installation of SMA devices would be greatly reduced. (3) Deployment tests can be performed repeatedly without having to remove/reinstall Pin Pullers.  The product tree for the SMA Pin Puller mechanism, including its classification by trigger mechanism is shown in Figure 4.  

Figure 4. Product tree for the SMA Pin Puller mechanism.     A novel Shape Memory Alloy   SMARQ is a fully European material technology and production processes, based on a low cost production procedure, which allows the manufacture of high quality products. Both material and production processes are currently being evaluated for use in space applications. In order to perform the complete characterization of the SMA material for its use in space environment, the SMA material (SMARQ) has been considered as the key element of a basic SMA actuator, consisting in the SMA material and the mechanical and electrical interfaces (Figure 2). The tests carried out as part as this work have been dedicated to the basic configuration of the SMA device. The tests have been focused on the actuator behavior when heated by means of an electrical current. The characterization of the rest of the elements – such as external heater elements, mechanisms, sensors or control electronics – in a space environment can be performed as independent bl ocks. A transitio n temperature (A s) of +145ºC has been  selected for the test samples in order to satisfy the initial environment requirements of operation temperatures between -70ºC and +125ºC.
117 sample immersion in the thermal bath. Special attention was paid to the sample thermal isolation, in order  to avoid the direct contact of the SMA with the thermostatic liquid, which would produce an important change in the convection conditions and therefore in the actuator behavior.  

    (a)                                                                            (b)  Figure 4. Devices used in the characterization campaign of SMAR: (a) Arquimea Actuators  test bench; (b) test bench for cryogenic test.   Lifetime tests  The scope of this test is to obtain an initial approach to the lifetime capabilities of the material. The tests have been limited to probe that the material is able to complete 100 cycles at the worst case temperatures (-70ºC and +125ºC) without degradation of its performance. A lifetime of 100 cycles is a typical requirement for one-use space actuators, such as release and deployment actuators. The test bench of Figure 4(a) was used for the lifetime test at +125ºC, while the test bench of Figure 4(b) was used for -70ºC.  Vacuum tests  Due to the importance of the changes in power consumption and response times under vacuum conditions, a vacuum test was included in the SMA characterization test campaign. The objective of this test is to obtain technical information about the behavior of the SMARQ actuator in vacuum conditions at ambient temperature. A specific test bench has been developed for the vacuum tests. The equipment allows the strain and force measurement as well as the actuator electrical activation in vacuum conditions under constant loads. The test has been done controlling the pressure inside the vacuum chamber (~5·10 -3 mbar) and at ambient temperature (22ºC). 

Test Results 

Strength characterization tests  The Maximum Strength Test was successfully completed. The material has shown its capabilities to withstand high loads, both in martensite and austenite phases. The material fractures at 500 MPa in its martensite state, showing an elongation over 12%. In the case of the test at higher temperatures, a stress of 400 MPa was applied without fracture of the sample.
118    (a)                                                                        (b)  Figure 5. Results of strength characterization and transition temperature test: (a) reconditioning  force – superelastic behavior; (b) SMARQ Transition Temperatures - Strain vs. Temperature.  Actuation above 125ºC is shown. 

These stress levels are far beyond the elongation and loads used in the actuators, so the material shows good properties for its application in terms of strength. The results show that a maximum superelastic strain around 10% would be feasible.  During the Reconditioning Force Test, important information was  collected about the force required to deform the material at different temperatures. The Superelastic Effect test has shown the typical behavior of superelasticity, Figure 5(a), with the existence of upper and lower load plateaus at different temperatures, where the deformation takes place with approximately constant stress. A maximum deformation of at least 8% could be achieved during superelastic cycles. Information about the Clausius-Clayperon curves was also obtained after this test. The Pull Force Performance Test has successfully been completed. The transformation has been observed and the stress levels obtained are high. The material output force capabilities are shown to be over 120MPa. 

Transition temperatures test  The test results of Figure 5(b) show that the material is able to work at environment temperatures over 125ºC, since the austenitic transformation starts, under the applied load conditions, around 145-155ºC. Besides, the material shows a martensitic transformation which finishes over 135ºC, so the material would be completely in martensite state along the whole environment temperature range (-70ºC, +125ºC). A thermal hysteresis of 15ºC has been obtained during the transformation and an 8% strain has been recovered.  Electrical activation tests  The test has shown the capabilities of the material to be electrically activated in the extreme temperatures in the working range, as well as in ambient temperature. This test shows that the material is able to work perfectly at these temperatures. The main differences between the tested temperatures are the energy necessary to complete the actuation, in terms of power and time, due to the higher difference between environment and transition temperatures. Information about the maximum non-firing current, activation currents, response times, and power consumption at each temperature have been obtained. The behavior of the wire during the electrical actuation was similar to the one produced with the oven in previous tests (transition temperatures test). A good stability and repeatability has been obtained along the cycles. Information about the wire resistance and its relationship with strain has been obtained. Moreover, an estimation of its relationship with the temperature has been carried out.