ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
344 The need f or a new separation system development is linked to  the drawbacks of the current  system s:  • A relatively high shock load during the separation event is transferred to the rest of the launcher  structure, including the payload. This necessitates that all the components attached to the PLF,  including the rest of the launcher and p ayload, are d imensioned  to sustain this load case.   • The fairing separation typically takes place at over 80 km of altitude, at the last layers of the  atmosphere, with conditions similar to vacuum. For ground testing, however, t he reaction forces  from the pyrotechnic separation s ystem are affected by the air resistance, driving the long-term  clearances. For this reason, the pyrotechnic separation systems  typically  need to be tested in  vacuum conditions to be fully representative of the separation event. Today there are few locations   worldwide where this test can be carried out due to the size of the test items, structures over 20  meters in height. One of them is NASA’s Glenn Research Plum Brook  Station in Sandusky, OH.   • Complexity of triggering: a drawback is the demand in accuracy  to synchronize the horizontal and  vertical separation systems in order to control the clearance.  

Figure 1: RUAG fa iring test at NASA Glenn Res earch Center Plum Brook in San dusky, OH [1] 

The main focus of th e new development s described in this paper  is to overcome the above drawbacks of the  pyrotechnic based separation systems used today . The system under development is designed to cover the  separation of different PLF classes with low shock, enabling on- ground tes ting while ensuring the reliability  and robustness of current pyrotechnic based solutions. The envisioned system consists of a scalable  horizontal and vertical separation system (HSS & VSS , not part of this paper ), compatible with a passive  jettison system using hinges and actuators that once triggered, rotates and pushes away both PLF halves in  a controlled maneuver . The focus of this paper is on the  developed  mechanisms of the passive jettison  system.   System Description  

The within this paper  presented mechanisms  are developed for the payload fairing’s  passive rotation and  jettison after the HSS and VSS release. The system consist s of two independent components :  • Pneumatic Actuators:  to provide the kinetic energy for  the PLF rotation.   • Preloaded Hinges: define the PLF rotation axis and provide the jettison kick -off energy.
345 The functioning principle of the PLF separation system under development includes three phases as visible  in Figure 2:  1. Low shock actuated horizontal and vertical separation system (HSS and VSS) , which are not part  of this paper.   2. Passive PLF rotation: actuators push the PLFs to a controlled rotation, guided by the hinges.   3. Passive PLF jettison : the hinges release the connection of the PLF to the launcher at a given angle.  Additionally,  a jettison energy is implemented to the PLFs by pre- loaded springs, in order to ensure  clearance.  

Figure 2: Separ ation System Principle  

This approach has several  advantages:   • Signals do not need to be synchronized: Due to the use of passive systems, there is only a single  triggering signal to release the HSS and VSS and initiate the fully passive separation.   • Modular:  The separation and jettison systems are independent , which allows  individual customiz ation   or even replac ement of  each sub-system for a different product without affecting the other subsystem.   Additionally , the actuators and hinges can be outlined and tested independent ly.  • Scalable: The solution under development and described here is  meant for small and medium size  fairing and may also be applicable to larger firings as well,  keeping the same working princi ple and thus  reducing the development effort.   • On-ground testable: The non- pyrotechnic separation is slower than pyro- based separations. As a  result, the friction with air becomes less dominant, which makes ambient verification possible.   • Simpler modelling: The functioning principle allows for a simpler kinematic modelling.   • Low shock: The k inetic energy for the jettison is not provided by an instant impulse but by a continuous  force from the actuator and the springs inside the hinges.   • Clearance control: the entirely  guided rotation allows a higher control of the clearance to the launch  vehicle.  

In the following paragraphs the selection and development of the actuators and hinges for the separation  system are described separately.  

Actuators  

Given the system concept described above, the actuators  need  to provide the thrust to overcome the a pex  of the fairing half shel ls against the launcher’s  acceleration loads and provide the kinetic energy to reach  separation.  Commercial of f the shelf gas springs have been chosen as passive actuators .    Spring  Loaded  HingesPLF Half  Shell HSSVSSPneumatic  Actuators Jettison  Spring  ForceCoGCoG  1) HSS & VSS Separation  2) Rotation  3) Disengagement and Jettison
348   Figure 4: Actuator I/F schematic  

Characterization and Qualification:   Using a gas spring as the actuator brings the difficulty of a changing pressure and resulting force with  changing temperature. Pressurization needs to be adapted in a way to have the required force at  separation,  but the actuator needs to sustain a possible high temperature increas e. Thus , the  characterization and qualification campaign enveloped  a broad range of  temperatures , for different launcher  classes and different missions . This temperature range includ es possible low temperatures  when the  actuators are installed near a cryo stage of the launcher as well as the temperature increase caused by air  friction during launch.  

The qualification campaign (including lifetime  and thermal characterization) has been performed on  different COTS actuator  sizes, applying full mechanical and thermal loads as per medium launch vehicles  PLF qualification requirements . As a result, a fully qualified pneumatic actuator at low mass  and broad  application possibilities has been found.  

Lesson Learned:   One of the lessons  from testing is that the material for the sealing is critical in terms of heat capability and  sustainability to the vibrations loads during launch. Dedicated testing has been performed to select  a  suitable material enveloping the range of temperature, the range of pressure and the high vibration load  requirements.  

Hinges  

The hinges guide the PLF during the rotation and provide the kick -off force to jettison the PLF halves. Two  hinges are used per PLF half shell to ensure a balanced guidance of the PLF during rotation and enabling  separation in the launcher’s rotational accelerated cases by avoiding high moments. The hinge’s main  functions include the following:   • Provide the mechanical interface during the fairing rotation phase.   • Transform the actuators forces to rotational movement of the fairing.   • Guide the fairing during rotation with low friction.   • Accommodate external f orces during rotation.   • Disconnect the mechanical connection at a given angle.   • Provide a kick -off energy to jettison the PLF halves  ensuring clearance.   • Provide the correct force direction for the kick -off.
349 The developed solution is a very simple design to ensure passively the release at a given angle providing  a kick- off energy to ensure clearance to the launcher.  The hinge, in disassembled condition,  consists of two  parts  as visible in Figure 8:  • The launcher bracket  is attached to the launchers last stage. It includes the bolt  for the rotation  axis and the protuberances  that lock the hinge in closed configuration and counteract the k ick-offsprin g. This part is fixed on the launchers last stage and cannot be jettisoned with the fairing.   • The hinge housing  is mounted to the fairing halve- shell and is jettisoned with the PLF . It  accommodated the kick -off-spring and the guiding pins,  located  in the slid of the launcher brackets  protuberances  in closed configuration.  

During launch, the hinge is in ‘closed’ condition as visible in Figure 8. The geometry provides the separation  of the hinge parts and the release of the loaded kick -off-spring . The preloaded spring is  accommodated  within the hinge housing, pressing against the rotation axle bolt  and counteracted by the launcher brackets  protuberances.  

During the rotation phase, driven by the actuators,  the hinge gets tilted into ‘open’ position. During this   movement, the g uiding pins slides over the surface of the l auncher bracket’ s guiding protuberances. When  the guiding pins reach the end of the launcher  brackets protuberance the hinge is ‘open’ . The geometr ic  constraint s are not given any more, the hinge housing  is not locked to the launcher bracket anymore,  and  it gets pitchforked, with the fairing,  by the released kick -off-spring system.  

Two of the described Hinge systems are installed at each PLF half shell.  This avoids moments on the  hinges, coming from external forces on the PLF.  The alignment of the hinges needs t o be adjusted on each  fairing type. This can be done in a simple way by building interface brackets  to meet the launchers shape  within the following geometrical constraints, mainly respective to the alignment  of the fairing’s  rotation axis   during the separ ation event :  • The hinges common rotation axis needs to be outside the launcher’s envelope as visible in Figure  5. This is necessary as to avoid jamming or the need of tr ansversal movement before the rotation  phase.   • The rotation axis needs to be below the fairing’s interface line to the launcher last stage, in order  to avoid jamming at horizontal or vertical connection line. The correct placement is also dependent  on the geometrical shape of the horizontal and vertical interface line.   • The hinge kick -off-spring force direction is to be aligned with the fairing center  of mass. This avoids  additional moments introduced by the k ick-off and enables optimal use of the kick -off-energy for  clearance.   •   •   • Figure 5: Hinge alignment (1)  

In the following  section, the development from concept to qualified system  is shortly  described  including  research studies and lessons  learned.
350 Development :  In a preliminary development study , the principle of the hinge was  developed and a functional verification  was run with a modifiable demonstrator model , shown in Figure 6, to prove the concept functional ity. The  hinge for the demonstration model  is designed to have a broad range of adjustable functional parameters  as to accommodate different springs and featur ing an adaptable spring force and stroke  length to vary the  kick-off energy.  The demo nstrator consists of  a fairing dummy, down scaled from a large launch vehicle.  The dummy was constructed with adaptable mass distribution, in order to simulate possible design changes  on the fairing for the new separation system as well as asymmetric mass distribution to assess the  robustness of the system  by test . Only one hinge has been attached to the demonstrator, allowing to  investigate and test occurrence of jamming due to asymmetric actuator force introduction.  

Figure 6: Early stage hinge during concept test ing on the demonstrator  model  

The demonstrator  test verified the principle of the system as well as the concept of the hinges. Several test  variations in terms of initiated force direction, energy variation of actuator  and kick -off-spring were  conducted and showed the high robustness for the new jettison system.  By varying a broad range of  parameters, the concept also proved to be scalable to the use for different launcher classes  with low  additional work.  

Lesson Learned  and Research:   Along with the  successful verification of the hinge’s  principle and initial correlation  to analysis, t hose tests  showed that an improvement of the tribological behavior of the hinge’s gliding surfaces was necessary , i.e.,  the kick -off plunger to the axle and the suspension to the guiding pin. Consequently,  a dedicated trade off  and testing campaign was conducted to find a suitable material combination which ensures a low friction  and avoids abrasion, jamming or cold weldin g. 

Three different material combinations , including a bearing,  were selected by a tradeoff  and tested   extensive ly for gliding  movement  with high forces  towards each other . The involved materials for the gliding  surfaces are aluminum with Ematal coating and stainless steel. For the counter -acting surface t hree  different materials and corresponding coatings were tested:   • Material 1:  Titanium  with Dicronite coating   • Material 2:  Stainless steel  with fiber reinforced plastic composite bearing   • Material 3:  Stainless s teel with Teflon- filled coating
351 A hinge dummy has been buil t simulating the internal loads by the preloaded kick -off-spring and external  loads , simulating forces and moments on the hinge caused by  the separation actuators and the launchers  vertical and rotational acceleration at separation.  In Figure 7 the results  in terms of friction are presented.  

Figure 7: Material Test for the Gliding Surfaces; Test Set -up and Results  

Material  1 was discarded due to the high  friction increase occurred at high loads . In contrast,  both materials   2 and 3 achieved the test success criteria of low friction and wear. Comparing material  2 and 3 , combination  2 required additional parts (bearings), making the design more complex. In addition, material  3 also  performed to high loads without change of the friction coefficient and no degradation was seen after  exhaustive lifetime  testing . As a result , the gliding surfaces material selection process was completed by  the selection of material 3.  

In addition to the material selection, the interfacing geometry of the hinge dummy  for the material tests  were  built similar  to the qualification model  hinges,  including a preceding analysis for the hinge tolerances to  ensure that no jamming occurs in the full temperature range due to different material expansion. This  allowed preliminary functional tests to verify the concept against jamming in case of lateral or torsional  external loads on the hinge dummy  set-up, including life- time tests for the material and surfaces.  

Subsequently, dedicated flight model hinges were designed, built and tested to qualification loads according  to the customer requirements for a middle launch vehicle (MLV)  PLF. The campaign included  thermal tests,  vibration and lifetime . In Figure 8 the fully qualified hinge is shown, mounted on an MLV fairing with  dedicated adapter brackets.  

The generic layout of the hinge allows the use on several  different types of launchers . The only part needing  additi onal design work are the INTERFACE  adapter brackets to align with the fairings and upper stage  curvature with the above described geometrical constraints . Additionally, the hinge is outlined to be  scalable, the size is adaptable to accommodate forces from different launcher classes, if required.    1
352   Figure 8: Hinges for MLV PLFs  

Full Scale Test 

One of the advantages of the system is the possibility  to test under ambient conditions  without the need to  use a vacuum chamber . To verify the functionality  and correlation to the analysis , a full-scale  separation  test on an  MLV fairing has been performed. Figure 9 shows the test set -up at RUAG Space premises in  Zürich/Switzerland . For this test, the fully qualified mechanisms (hinges and actuators), from the  qualification campaigns as described above have been used.  

Figure 9: Image of the PLF during rotation of the  full-scale  test 

For several  separations runs, tracing the fairings movement  with high -speed camera targets and  accelerometers at  different  failure modes and degradation cases, the separation and jettison hinge and
353 actuator system have demonstrated to be a simple and robust system with repeatable and predictable  results.  

Figure 10 shows the trajectory of the fairing during and after rotation and jettison for several test runs with  varying energy. The picture shows the movement of the hinge in radial (y) and gravitational ( -x) direction,  captured and tracked  with a high -speed camera. The movement is robust against the variables that  reduce  performance of actuators and spring kick -off spring energy and can be adjusted very accurately.  

Figure 10: Full scale test trajectory  

Each run show ed similar rotation behavior , accelerations, trajectory and rotational speed. The test outcome   comparison to FEM prediction results showed good correlation,  considering the prediction of the effect of  air resistance.  

Besides the repeatability and good FEM correlation of the rotation and jettison system, the methodology  and procedure to test in ambient air by track ing the trajectory and accelerations has proven to be a reliable  methodology to verify the system for future separation tests  on different launchers . 

Conclusion  

This paper summarizes the activities carried out by RUAG Space within  the Future Launchers Preparatory  Programme – separation and jettison project. In the frame of this project, a low -shock jettison system has  been developed based on the functional principle of a rotation, disengagement and jettison by means of  passive gas springs actuators and spring loaded hinges. Q ualification models of the mechanisms have  been built and qualificatio n testing has been successfully completed on component level and on system  level in a full -scale test (TRL 6) . An extensive  test campaign was carried out to verify the suitability and  prove the advantages of the newly developed technologies with regards to simplicity and verification efforts.   Additionally, a full -scale test has been successfully conducted for the validation of the separation analysis  model proven functionality of the technologies developed. Overall, the test results are proving the  functionality of the new jettison system.  

Acknowledgements  

The authors would like the thank the European Space Agency  (ESA)  and in particular the Future Launchers  Preparatory Programme for supporting and granting a part of the investigation shown in this proceeding, in  the frame of the Contract No. 4000121430- 17-F-JLV.
354 References  

1. European Space Agency, 2012. ARIANE 5 fairing separation system undergoes testing.  http://www.esa.int/spaceinimages/Images/2012/06/After_separation2  (accessed 28.01.2020)   2. ESA Standards: ECSS- Q-ST-70C rev2: Materials, mechanical parts and processes   3. Sanchez A.,  Halter B., Gerngross T., RUAG’s development of a modular low -shock jettison system,  International Astronautic Conference IAC -19, October 2019, Washington, USA   4. Caldirola L., Schmid B., Cattaneo M., Schiaffini A., 2017. Test HDRM” development for separation  performance verification.. In European Space Mechanisms and Tribology Symposium.   5. Arianespace, 2014. VEGA User's manual issue 4. Online, available in  http://www.arianespace.com/wp- content/uploads/2015/09/Vega- Users -Manual_Issue- 04_April - 2014.pdf  (accessed on the 04 .06.2019) .
355 Deployment Mechanism for an Earth Re- Entry Foldable Cloth Decelerator  

Carl E. Kruger* 

Abstract  

Adaptable Deployable Entry and Placement Technology (ADEPT)  is a NASA technology development  project  that uses a folded, 3D woven carbon fabric decelerator to serve both as thermal protection and  primary structure . A sounding rocket test flight, with a 3U Cubesat sized payload , was used to evaluat e  ADEPT’s  in-space deployment performance and supersonic st ability . Size, mass and complexity limits  dictated a solely mechanical mechanism to deploy the folded  cloth decelerator surface. This paper  describes the design  constraints , system design components, hardware development, testing, flight test  execution, and flight test results  of the deployment mechanism used by this “Nano- ADEPT”  [1]. 

Introduction  

ADEPT is a technology development project  that uses a foldable carbon fabric  “skirt” to serve as the primary  drag surface for an entry vehicle decelerator . The carbon fabric acts as an aerodynamic surface to provide  drag while  also being able to withstand the high temperatures of re- entry . Using a folding design allows a  larger decelerator to be packaged with in any given launch vehicle ’s payload envelope. For small spacecraft  missions, a small version of ADEPT (sub- 1m) could be used to provide a lower heating entry environment  while still fitting within common secondary payload accommodati ons.   

This ADEPT Sounding Rocket One (SR -1) project  focused on  subsystem level testing of a 0.8-meter (NanoADEPT class)  decelerator . Eight gores of a multi -layer, woven carbon fabric are sewn together with carbon  thread to form a faceted 70° “cone shaped”  aeroshell . A wind tunnel test (Fig. 1) was used to characterize   the deflected fabric shape as a function of aero load, angle of attac k, and cloth tension . Arc jet testing   (Fig. 2) exposed carbon cloth samples and seams  to heating rates and heat loads expected for a Mars  mission . Lastly , a sounding rocket  test was executed to investigate spacecraft re- entry stability as the  vehicle decelerated from Mach 2.8 to Mach 0.8.                      Figure 1. ADEPT Fabric Deflection Wind Tunnel Test               Figure 2. ADEPT ArcJet Heating Test       

* NASA Ames Research Center, Moffett Field, CA  

Proceedings of the 45th Aerospace Mechanisms Symposium, NASA Johnson Space Center, 2020