ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
93 the real devices  can be obtained. These tests were completed in different environmental conditions  for each  model: ambient  conditions, minimum temperature with air (worst case for the lubricant) and  minimum and maximum temperature in vacuum. The models showed margins above 6X  in all cases,  as required by ECSS [4], with some models showing more than 11 X margin. This value shows the  reliability of the device to operate under worst conditions.  

    Figure 8.  REACT Qualification – Left: Vibration test. Right: Missalignment test  

Launcher Separation System  

For the qualification of REACT as the release device for  an Airbus DS separation system, a fully flight  representative clamp- band type LPSS937* was chosen as  the test item. The LPSS* Separation System is  the Airbus DS solution for clamp- bands where low shock and high load carrying capability are essential.  Airbus DS is a provider of Separation Systems in the European market for Ariane- 5, Soyuz and VEGA  launchers. Also, in the American market , Airbus DS provides Separation S ystems for Space- X and Orbital - ATK. Finally, in the Japanese market , Airbus DS clamp- bands fly in HIIA and will soon fly  in HIII launchers.  

    Figure 9. Airbus DS LPSS* Separation Systems (left). Endurance results during LPSS937* qualification.  

To enable the use of  REACT on LPSS*, some  mounting/ adaptation hardware was  designed and  manufactured. Later, the LPSS937* clamp- band, employing a REACT 15kN Extended Temperature model,  were subjected to the standard qualification test campaign applied to all Airbus DS flight separation  systems. The test campaign began with a functional release test, to verify that the REACT enabled proper  performance of the clamp -band. Although the functionality of the system for minimum, nominal and  maximum band tensions was correct, an unexpected hooking of the  “key” was detected. The “key” is a part  that trasmits the tension load of the clamp band to the REACT through a shock reduction mechanism. The
94 main outcome of these  tests was the need for a device to eject the key faster than the movement of the  shock reduction mechanism. This device could be a simple pull -out spring.  

The next step was the life  test campaign consisting of  ten tensioning/release tests to demonstrate that the  clamp- band will survive to launch after ten releases. The main outcome of this campaign was the  confirmation of the hooking issue (Figure 10 left).  

The LPSS937* plus REACT  assembly  was subjected  to a random vibration level of 20- G RMS showing a  stable and adequate dynamic behavior without high amplification.  After the successful environmental test,  the clamp -band was released to demonstrate the survival of all its components  including the REACT  release device.  

Figure 10. LPSS93 7* qualification setups using REACT. Vibration test (left). Thermal Vacuum Test  (center). REACT detail connected to the clamp- band (right).  

The thermal qualification of the LPSS937* plus REACT  assembly  was performed in vacuum conditions.  The assembly was submitted to four cycles between LPSS* extreme cold and hot qualification temperatures   of -55°C to 120°C. Functional releases were performed both at cold and hot conditions demonstrating the  correct function of the REACT as the LPSS* release device. Finally, the clamp- band was tensioned to its  nominal band tension and this parameter was measured and recorded for  more than one month to  determine its relaxation. The observed behavior was normal , with a relaxation  of just 0.3 kN in the  mentioned time period.  

Conclusion   • Tensioning operation torque/preload relationship and the repeatability of this parameter in  consecutive operations is similar using REACT to the release devices  previously used on LPSS* .  • It is necessary to implement an extractor for the key  bolt (bolt retained by REACT) either installed in  REACT body or as part of the LPSSS937* mechanism. The lack of this extractor creates an  incompatibility between the velocity of the key bolt driven by the opening of the band mechanism and  the movement of REACT fixed to the clamp- band main- beam. The result of this incompatibility is an  unexpected hooking between two pieces that causes an off -nominal band release.   • Main outcome of this work is that, although there are still a set of activities to be done prior to the  conclusion of  formal REACT qualification for Airbus DS clamp- bands, it is clear that the device is a n  option, at least for a given type of missions wher e the activation time of several seconds is acceptable.  

Planetary Exploration Subsystem  

Added Value Solutions  (AVS)  participated in the development and qualification of a Hold Down and Release  Mechanism for a planetary exploration subsystem. AVS designed a HDRM based on the REACT 5kN to  hold and release the Sample Canister of an interplanetary sampling tool. The “Sampling Tool Mechanism  for Low Gravity Bodies” developed by AVS collects a minimum of 100  g of regolith from the surface of a
95 low gravity celes tial body for later transfer to the next step in the sampling chain. Therefore, once the regolith  has been sampled successfully, it is necessary to perform two disengagement sequences:   • The first disengagement releases some parts of the mechanism leaving the canister exposed. For  doing this three HDRM's located 120˚ apart are needed.   • The second disengagement releases the sample canister and here is where the new REACT Hold  Down and Release mechanism (HDRM) design is considered.  

 Figure 11. Sampling Tool Mechanism (left), 1st Disengagement (cent er) and 2nd D isengagement (right)  

Two models were developed in order to qualify the 5kN REACT HDRM: an Engineering Model ( EM) to test   the functionality of the design and a Qualification Model ( QM), which was subjected to a qualification  campaign.  

The functional test campaign was  successful as all the requirements were met during the execution of the  complete series of measurements. The HDRM EM released every time after the REACT EM 5kN actuator  fired. The onl y issue encountered was related to the resetting procedure of the actuator, but they were  solved for the QM phase. The results of the qualification tests are explained hereafter.  

A total of 22 releases were carried out, two of them in vacuum (10-6 kPa) an d at a temperature of +120°C  and -90°C. All of the deployments were successful.  

The HDRM survived to 25- G quasi -static acceleration along all axes, as well as to 31- G RMS during the  random- vibration testing. Low -level sine-sweeps were performed before and after each random- vibe test.  The pre-  and post - sine sweeps were compared to verify that there were no eigenfrequency shifts greater  than the predetermined tolerance of 5%.
96       Figure 12. Vibration Y Direction (lef t). Shock Test Setup (right)  

    Figure 13. Sine Survey Comparison ( Before and After 2 S hock Tests) at SA1 Channel (left). Test Setup  for the Thermal Vacuum Cycling (center) and Cycles C arried Out During the TVac (right)  

Conclusion   In conclusion, it can be stated that the 5kN REACT successfully passed the functional and qualification  campaign. The issues with the resetting procedure that occurred during EM testing were resolved prior to  starting the qualification campaign. The QM testing vallidated thes e changes and this time the reset s were  simple and repeatable.  

Large Structure Deployment System  

As part of the H2020 REACT project, SpaceTech GmbH (STI) verified the use of the new release  mechanism for the separation of large solar generators, as typi cally used in space travel. A wide- ranging  test campaign examined the possible uses for this actuator. The main objective was the qualification of the  REACT actuator, preferably for multiple deployment jig designs for STI projects. The tests were performed   using a 15-kN series actuator (QM). The basis for the physical design of the actuator mounting to large  deployment structures were the interface points, for example as in the Sentinel S5P -project. Adaptat ors  and test benches (GSE’s) for all mechanical tes ts were designed (CAD), analyzed (FEM) and manufactured  by SpaceTech GmbH accordingly. After the test series was carried out, the following was determined for  the respective tests:  

Electrical Test:  The REACT actuator (QM) has two redundant electrical connections in case of failure of  one circuit. The resistanc e of the nominal and redundant lines were measured. Both circuits indicated the  required values. Additionally, isolation tests were performed to ensure that the conductive lines and the  mechanical structural parts maintained electrical isolation. The insulation measurement revealed full  isolation.
98   Figure 14. REACT Actuator (QM) T est Setup Vibration Test, Sensor Positions  

Conclusion   The initial problems lay with the misunderstanding of the narrow tolerance for the screw -in depth during  system assembly, which led to functional failures during the tests. This was identified and solved. Based  on this, suggestions were made that should lead to an improvement in the monitoring of the actuator during  the various test phases.  

The release shock test results indicate an exceedance of the required limits at the interface points (input),  especially in the upper frequency range. Whereas on the dummy mass, the shock loads remained well  below the requirement. These exceedances can be explained by the fact that the interface measurement  points were located directly on the bracket for the tests performed in this project. At the output interface  point s, the shock response spectrum levels are likely to be significantly reduced by the damping of the  structure.  

The final actuator mass and dimensions are significantly higher than the actuator of other competitors,  which makes it especially difficult to use the REACT as a standard deployment solution for SpaceTech at  the moment. If the mass of the actuator could be reduced, it would be an extremely attractive solution  compared to the products of the competition. Above all, the extremely fast -recoverable pr eload is an  enormous advantage. As already mentioned above, the reliability has been tested at SpaceTech GmbH in  a lifetime  comprising either electrical or mechanical releases far above the qualification limits.  

Small Satellite Deployable Subsystem  

The project undertaken at SSTL tested the 35kN standard temperature variant of REACT from Arquimea.  SSTL were involved from the start of the project, defining requirements necessary for use in small satellite  deployable systems. Test campaign of an EM could not be completed due to the mechanism not holding  the maximum preload after a certain number of tests. This was mainly due to the selection of an excessively   fine thread pitch. Arquimea further refined the design after these issues and sent a Q M for SSTL to test.  The qualification tests were completed and confidence was gained that the mechanism could perform  consistently.
100                Figure 15. Proof Test (left). Function Deployment Test Setup (center). Use of Reset Tool (right).  

The resetting of the actuator (Figure 14 right) was easy and the feedback from the mechanism gave the  operator confidence that the device had been reset properly. At no point was there any concern that the  reset may not work or that the mechanism may not hold preload after a reset.  

Throughout the testing the REACT QM device was released 15 times (6 Pre- EVT, and 9 either during or  Post-EVT), against its specified life of 10 releases on ground and 1 in orbit, which gives good confidence  of the device’s life.  

Conclusions  

A whole family of HDRAs has been developed and qualified in this activity, including their quailification in  systems for different applications. One of the key advantages of REACT is its resettability, which allows the  users to perform integration and multiple tests in a short time, espec ially when compar ed to other existing  solutions. This characteristic is especially interesting as a cost -efficient solution to improve the  Manufacturing, Assembly, Integration and Test  operations at system and satellite levels. Moreover, the  resettability enables a test -what -you-fly and fly -what -you-test philosophy.  

In addition to that, the temperature range is competitive for most applications, giving different actuation  options for different applications and uses (shorter actuation time and power for lower temperature  environments, different load levels, mechanical interface options including compatibility with existing  solutions,…). Another important advantage is its proven reliability, the high mechanical and force margins  shown by the devices in the w orst environment conditions.  

The devices have shown successful operation under a wide range of preload levels, from 0N to levels over  their nominal preload, allowing loads over proof levels to be released during the campaign. In addition to  that, the devi ce showed  capabilities to withstand loads much larger than nominal, maintaining functionality,  which gives an additional robustness level , showing survivavility to any unexpected event. Finally, the  device shows good stiffness and creep behaviour.  

During the work, several lessons learn ed have been identified:   • A good analysis of the concept and its potential failure modes should be done as early as possible in  the development.   • Reducing the contact pressures in the critical friction points is very import ant.  • The selection of lubricants together with a proper definition of contacts, materials and preparation of  the surfaces is critical for a good and reliable space mechanism, as shown after the issues detected  on the design selected at the beginning of t he activity.
102 References  

1.  Collado, M., Rivera, C., Inés, J., Sánchez, J. "  Evolution of a resettable Hold- Down and Release Actuator   Based on SMA Technology ."  18th European Space Mechanisms and Tribology Symposium,  2019   2.  Collado, M., Cabás, R., López -Ferreño, I., San Juan, J. "  Functional Characterization of a Novel Shape  Memory Alloy. " Journal of Materials Engineering and Performance, Volume 23, Issue 7 (July 2014),  2321- 2326; DOI: 10.1007/s11665- 014-1104- 7.  3.  I. López -Ferreño, U. Urrutia, P. Lorenzo, M. Collado, C. Rivera, N. Escudero, T. Breczewski, M.L. Nó,  J. San Juan, “Ultra- High-Vacuum experimental equipment to characterize shap e memory alloys for  space applications”. Materials Today: Proceedings 2S  (2015) S953- S956.   4.  ECSS -E-ST-33-01C - Mechanisms.
104   Figure 1: PSP Side view: The primary components of the Antenna Retent ion System (ARS)  and FIELDS  instrument  on Parker Solar Probe  

Figure 2: The three variations of the clamshell mechanism mounts for 4 different  mounting locations  

Clamshell Mechanism  Design  

The PSP clamshel l mechanism is comprised of two, half -circular machined aluminum ha lves, one  deployable or active hinge, two follower hinges, redundant  deployment  telltale switches , and plastic antenna  guides  that are mounted to  the top end of the clamshell.  The clamshell halves are 762- mm (30- in) long,  have a 22- mm (.875- in) external diameter,  and are fabricated from 6061 aluminum, with aggressive mass - reduction features  on the inside. The clamshell mounts to the  PSP SC bus via features on  the top and  bottom hinges, on  brackets that position each clamshell on a natural  tangent of each stowed antenna.  These brackets were designed  with multiple bolted interfaces  to provide multiple degrees of  in-plane   angular and translational  adjustability to ensure proper in-situ alignment with each stowed antenna.     Clamshell Mechanism   Containment Forks   FIELDS Antenna   FIELDS Antenna   SC +Z Axis
105 

Figure 3:  Top view:  The clamshell mechanism that corresponds to the V1 FIELDS antenna position.   

When deployed, the clamshells hinge open along the length of the tube created by the halves . They do not  open much; only 15 degrees per half ( Figure 4) , but this is more than suffic ient to release the 3.2- mm (.125in) diameter antenna. They are not designed to physically clamp  or restrain  the antenna, but  only to loosely  contain it. The 6.4-mm (.25 -in) clamshell ID enables to the antenna to rattle  slightly  and freely translate   axially , as not to impart any additional loads i nto the FIELDS instrument from the SC  deflections that occur  during SC  vibration  events . The closed side of o ne clamshell half features an “anti -trapping” tab along its   length, to prevent the antenna from getting hung up or caught in the gap formed when the mechanism is  deployed.  Similarly, t he yellow Ultem®  blocks on the end of each clamshell half ease the transition of the  antenna  into the end of the tube; a location d etermined to have the highest dynamic contact. Extensive  material testing was performed to ensure Ultem® compatibility with niobium  antennas .            

Figure 4: Top view: The clamshell mechanism depicted in the  closed (left) and open (right) condition.   

The heart of the mechanism, the active hinge,  utilizes a TiNi Aerospace P10 P inpuller to provide the motive  force to deploy the greater mechanism.  A TiNi P10 pinpuller is designed to retract 9.5  mm (.375  in) with   44 N ( 10 lbf) of pulling force2. Due to the shear load limitations of the pinpuller  pin as specified by TiNi , it  was necessary to isolate the vibration loads imparted by the mass of the clamshell  from the pinpuller pin.   This isolation is performed by a titanium slider cup, whic h is fastened  to the end of the pinpuller via a loose   spherical interface and engages two titanium legs built  to each clamshell  hinge  half. When the pinpuller is  actuated, the sliding cup translates  along the axis of the pin, disengaging the clamshell legs,  and enabling  each spring- loaded clamshell half to pivot  open  (see Figure 5) . The spherical interface between the
106 Pinpuller  pin and the cup prevents binding of the sliding cup in the housing by eliminating all potential over - constraining degrees of freedom. The intentional loose fit of the spherical interface ensures that no external  moments or side loads can be imparted onto the pinpuller pin due to assembly tolerances or launch shifts.   Only the retraction force can be imparted into the cup.   

Figure 5: The details of the active clamshell hinge, in stowed and deployed positions .  Left: Cross -section, Right: Housing not shown  

The other two hinges are spring- loaded follower  hinges that simply provide additional opening torque via  torsion springs at each location. These hinges feature no release mechanism or lock, and the mass of the  center hinge is supported entirely by the clamshell halves. The bottom hinge is designed to accommodate  ±3.2 mm (.125 in) of  change in the axial length of t he clamshells ( Figure 6), and t his eliminates the potential  for the mechanism to bind from temperature- induced dimensional changes (CTE).  Furthermore, the bottom  shaft was cut with a profile of spherical undercuts, to further minimize the potenti al for binding, and provide   additional  angular adjustment during SC  integration.      

Figure 6: A side view of the bottom hinge illustrates the bottom mounting bracket and the gaps designed  to accommodate CTE effects  of the aluminum clamshells     “Slider Cup”  Engagement Legs   TiNi P10  Pinpuller  Spherical  Interface
107 Also a ttached to  the active hinge  bracket are four Honeywell 9HM1 micro switches to provid e redundant  deployment telltale signals  of each clamshell half. Resistors of different values were placed in- line in the  harness, as to indicate “left -open”, “right -open” and “fully -open” states. A small aluminum lever arm attached  directly to the clamshell engaged the switches directly when stowed (Figure 7)  and disengaged only when  the cl amshell reached the fully open, 15-degree state. These switches functioned perfectly upon the actual  deployment in space but proved to be very difficult to set and tune on the ground.  The #2 attachment screws  needed to be torqued enough to prevent shifting of the switch, but not so much to crush the thin exterior  case. The risk of a false telltale  reading was also partial ly mitigated by analyzing the SC  IMU data to show  the shock of each at deployment.  

Figure 7: A bottom isometric view of the redundant telltale switches  

Materials and Coatings  

The mounting brackets and clamshell halves are fabricated from aluminum,  and are iridite coated and  electroless nickel plated, respectively. All of the mechanism parts are 6Al -4V titanium, including the active  hinge housing, the sliding cup, the hinge halves and shafts. This fact becomes an important detail in the  anomaly investigation to be detailed later. The decision to uniformly fabricate all mechanism parts from  titanium was made to eliminate the effects of CTE mismatch withi n the tightly -fitting, high- precision  mechanism components , and in some cases, for  strength considerations . The two halves of the titanium  spherical interface screwed  onto the pinpuller are unfilled PEEK, the springs are E lgiloy ®, and all the  fasteners are  stainless steel.    All titanium mechanism components are titanium anodized with a type 2 T iodize ® process, with local  application of Dicronite ® (tungsten disulfide)  dry film lubricant  on all moving or sliding surfaces.  No  additional lubrication was used in the assembly, except on the fasteners.
108 Development Effort  

As a result of the late schedule addition of the Antenna Retention S ystem, the clamshell mechanism was  developed on a very aggressive timeline. Two engineering model (EM) versions of the clamshell  mechanism w ere built and tested before the flight units were fabricated. The first EM provide d a successful  proof of concept of  the mechanism, and the environmental testing scheme. T he second EM implemented  most of the flight -like mass reduction efforts  and added the telltale switch assembly . The differences  between the second EM and the flight implementation were intentionally  very minor, including  maintaining  EM2-to-flight continuity with the fabrication shops  selected for each part . 

Each iteration of the design was fully vibration  (Figure 8)  and thermal vacuum ( TVAC ) tested, and  dozens  of ambient temperature, post -vibration and hot/cold deployments  (Figure 9)  were executed. Additionally,  ARS system -level containment tests were conducted  at three points during the PSP development effort:  during the SC modal test, the full SC  vibra tion test , and a specially  built SC  “stiffness simulator”, as shown  in Figure 10. This simulator provided a test bed to perform ARS -level tests at higher vibration levels, while  eliminating the risk to the actual SC  bus. In each case, s urrogate FIELDS antennas  were captured in EM  clamshells  and forks. Efficacy of the retention system was determined through high- speed video and   measurements from the video of the antenna maximum excursions  as observed during the modal test, and  the SC stiffness sim ulator vibration tests . The flight clamshells were qualified separately and added to the  SC after the complete SC  vibration test, but before the complete SC  TVAC test.  

Figure 8: V1 and Spare flight clamshells on a vibration table
109   Figure 9: Flight clamshell mechanisms staged for TVAC testing.         With the exception of some erroneous telltale indi cations observed during the EM2  clamshell development,  no anomalies occurred during mechanism- level testing. The telltale design was modified and follow up  thermal and vibration testing was conducted to verify the correction. This was considered to be a reasonable  step in the development process, and while additional attention was given to the telltale configuration as a  result, it was not  considered to be a serious  anomaly.  No deployment failures occurred through EM2  development, and the clamshells enjoyed a perfect record of functionality.  

                                         Figure 10: Left: EM Clamshell with preliminary fork configuration.   Right: EM Clamshells on structure “ stiffness simulator ”