ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
356 Flight Concept of Operations   The ADEPT SR -1 vehicle was designed for the following concept of operations:   It would  be launched to  an altitude of approximately 100 km (62 mi) in the payload module of an UP Aerospace SpaceLoft XL  sounding rocket . Shortly before apogee, ADEPT SR -1 would be ejected by a spring- loaded sabot assembly  in a “tail first” orientation, so that as it fell to earth, it would re- enter nose first . Approximat ely 40 seconds  after being ejected (to reduce risk of re- contact), the on- board timing circuit  would command  the carbon  fabric decelerator to deploy . SR-1 would reach a peak Mach number of 3, at approximately 70 km (43 mi)   altitude  with no significant heat ing. Supersonic and subsonic stability would be determined using recorded  accelerations, body rates, and orientation references for stability analysis at a later time. A hard impact of  25 m/s would be attenuated with crushable material . Data stored on hardened SD cards would be recovered  from the impact site (no telemetry  was used) .  Design Constraints   The deployment mechanism design was constrained by mass, volume, and C enter of Gravity (CG)   limitations . An internal volume with dimensions of  10 cm x 10 cm x 30 cm (to simulate a possible 3U  CubeSat  payload) was  reserved in the center of the vehicle  (for this demonstration flight, this  payload area  housed batteries, electronics, sensors, and impact attenuation foam) . The external boundary  was d efined  by the 25.4- cm (10- in) internal diameter of the UP Aerospace launch vehicle payload module (Fig. 3). The  deployment mechanism, including the cloth and ribs, was limited to the space between these two volumes .  The overall length of the vehicle was constrained by the need to avoid high-speed flow impingement on the  aft payload as the wake expands off the aeroshell shoulder.  

Figure 3. ADEPT SR -1 Prototype within  Simulated Payload Volume   Cloth shape and mass properties were constrained. For aerodynamic purposes, cloth deflections were  required to be limited, and similar to those characterized in the wind tunnel test . The ribs and struts had to   lock into position when fully deployed so the cloth maintained  a constant tension during re- entry . In order  to maintain stability during flight, the vehicle’s CG location and M ass Moment of I nertia (MMoI) were closely  tracked and adjusted.   ADEPT generally employs a rigid nose cap, which is used to cover the central payload as well as the  deployable aeroshell when it is in the stowed position. In a typical high- heating re- entry  environment , such  as for Venus , it would  carry a traditional thermal protection system , however for SR -1 a metallic nose cap  was sufficient . As the carbon cloth skirt transitions  from a stowed config uration to a fully open shape, the  cloth must slide underneath the nose,  and the nose cap must contact the cloth securely at the end of travel  to prevent gaps that could allow hot gases to enter the payload region.
357   Figure 4. ADEPT SR -1 Flight Unit in Deployed Shape   Design O verview   ADEPT SR -1 uses  coil springs to deploy the ribs and fabric . Long extension springs provide most of the  deployment with a small force and long displacement . A retention cord constrains the rib tips . It is severed  after launch and a moving ring which pushes the ribs outward is pulled towards the nose of ADEPT . Near  the end of travel, a trip mechanism releases  plungers containing short , stiff compr ession springs, which  provide the final high force small displacement latching. In addition, the nose cap clamps  down on the fabric  (which needs to float as the ribs are deploying) just before latching . The time between retention cord  severing and fully -locked deployment is less than one second.  

Figure 5. ADEPT SR -1 Layout
358 Framework   The basic mechanism works like an umbrella, with rigid ribs and struts . The carbon cloth is attached to the   eight ribbed structure resembling an umbrella frame. The cloth is secured at the tips of the ribs  and the  cloth rests on  the ribs . It is not attached at the center, so the tension in the cloth distributes evenly over the  eight gores as the ribs are deployed.  

Figure 6. ADEPT Prototype Showing Deployment   A moving  ring containing  the anchorage for the eight struts  slides  along the axis of the vehicle  on four linear  slides . As the ring moves from the tail to ward the nose, the struts push the ribs out to a 70- degree angle.  Unlike a conventional  umbrella  where the struts meet at a small ring, the SR-1 design employs a much  larger square- shaped moving ring which encircles the 3U Cubesat payload volume. The ribs pivot from the  nose of the spacecraft and are pushed outwards by struts located 1/3 of the way along th e ribs  (Fig. 7) .  

Figure 7. ADEPT SR -1 Ribs and Struts
359   Figure 8. ADEPT SR -1 Centerbody with Moving Ring and Nose Cap   Cloth Considerations   The three- dimensional woven carbon cloth used on SR -1 is not strictly part of the deployment mechanism,  but its properties are important to understand. The carbon cloth used for this version of ADEPT is a four - layer version, which demonstrated suitable folding behavior and tensioning ability (the cloth is st iff and   becomes too thick to fold in 6-  and 8-  layer weaves) . Eight triangular shaped gores are stitched together  using carbon thread to form an eight sided “pyramid” . A carbon cord is threaded through pockets at the  trailing edge, to provide a defined trailing edge shape and tension. As mentioned earlier, t he clot h is free to  float at the nose (Fig. 9 ). It is clamped at the underside of each rib tip  (Fig. 10). 

            Figure 9: ADEPT Floating Cloth under Nose Cap       Figure 10: ADEPT Rib Tips with Kickoff Springs   Rib retention and release   When stowed, the eight ribs are parallel with the body axis, so there is no mechanical advantage for the  long travel springs at the beginning of travel . For this reason, spring  steel  “kickoff springs” we re added to  each rib tip to initiate rib movement . These push the rib tips away from the centerbody far enough to initiate  rib movement . The rib tips are sculpted to guide the cloth into a trailing edge geometry that matches the
360 projects’ CFD model . In order to increase the vehicle’s Ixx (rotational inertia  about the central axis) to  facilitate spin stabilization, t he rib tips are stainless steel . 

ADEPT SR-1 is secured in the stowed condit ion by looping Vectran cord through hooks on all eight  rib tips  (Fig. 11). The Vectran is threaded through a pair of redundant spring- loaded Nichrome hot wire cord cutters  [2]. 

Figure 11: Vectran C ord Loop and Nichrome Burn Wire Assembly   Motive force springs   The project decided against using a motor and gearbox due to packaging, complexity and power budget  reasons . Precluding an electric gearmotor to drive the umbrella open led to the use of springs as the  deployment motive force. However, meeting the  cloth  tension requirement of 17.5 N/cm (10 lbf/in) derived  from the wind tunnel testing required  high force as the cloth was  close to full deployment . This is contrary  to the properties of extension and compression springs, which have their highest force at their largest  displacement and a much smaller force when approaching their  free height.   The mechanism designed to solve this conundrum  uses two sets of springs to a ffect the required  displacement and large final force. Eight pairs of long extension springs are used to provide the gross  movement of the moving ring. A second set of stiff compression springs are triggered near the end of the  range of motion, to creat e the necessary high force to tension the cloth and lock the ribs into position . The  stiff compression springs also provide the motion to pull the nose cap firmly against the cloth and clamp it  at full deployment .  

Figure 12. ADEPT Prototype showing Long- Travel Extension Springs
361 Figure 12 shows the long travel extension springs . They are stretched between the forward end of the  spacecraft and the moving ring. There are two extension springs at each location, one nested within the  other (for a total of 16 tension springs) , to provide suffici ent force without excessive  free length.   Moving ring   The moving ring (Fig. 1 3) is an assembly of four identical machined quadrants , guided on four small linear  slides, and fits in the space between the payload and the cloth. The stiff compression springs are housed  in plungers that slide within the moving ring. The moving ring is aluminum and the plungers are stainless  steel . No bushi ng or lubricant is used.  

Figure 13. ADEPT Moving Ring Assemblies (both sides shown)   Nose Cap and Pins  The nose cap (Fig. 1 4) was machined from stainless steel, to shift the CG forward . The cap contains four  pins that are aligned with the plungers that are on the moving ring . The nose and pins slide in flange  mounted linear bearings that are mounted to the forward end of  the ADEPT  centerbody  (Fig. 1 5). The nose  cap has approximately 4 mm of linear motion. It is lightly sprung away from the centerbody to allow for the  cloth to move underneath the edges of the cap. The deployment mechanism pulls  the nose cap down tight  against the cloth after the system has reached full deployment.          Figure 14. ADEPT Nose Cap with Pins and          Figure 15. ADEPT Centerbody Forward Structure with                         Proximity Switches                                                Rib Pivots and Nose Pin Bushings
362 Figure 16  shows the changes in section of the aluminum pins. The pins are the “triggers” for the multiple  functions that occur during deployment.  At the tip of the pin is a point, then a relief, and at the shoulder of  the pin is a taper ed ramp with a recess on the back side . As the moving ring slides from the tail of ADEPT  towards the nose, the pins pass through the center of the plungers .  

Figure 16. ADEPT Nose Pins and Latches   Spring Plungers  and Latches   One end of each compression spring is seated against the base of the hollow plunger, and the other end is  constrained by a tabbed washer that fits through slots in the plunger  (Fig. 1 7). The ID of the compression  spring is large enough to clear the pin.  

Figure 17. ADEPT Spring -Loaded Plunger (Compressed State)
363 This tabbed washer also limits the plunger rotation and displacement. As the pins pass through the spring  plungers, the pointed pin tips spread t wo sliding, sprung plate catches  (Figs. 18, 19) at the ends of the  plungers . The plate catch es capture the tips of the pins by engaging in the relief . As the moving ring  continues  towards the nose of SR -1, eventually the taper ed ramp on the pins spreads a pair of l atches that  have held the spring -loaded plungers in a compressed position (Fig. 20) .              Figure 18. ADEPT Nose Pin and Plunger Catches             Figure 19. ADEPT Nose Pin Sliding Plate Catc h

Figure 20. ADEPT Plunger Assembly  with Notch for L atch 

The four plungers extend,  and since they have captured the tips of the pins, the nose cap is p ulled aft . This  causes the now -deployed cloth to be clamped at the nose, which was a requirement . Additionally, once the  nose cap reaches its  end of travel, any further movement of the pl unger will cause the moving ring to be  forced towards the nose of SR-1. This is the last bit of high- force movement required to pull the cloth to full  tension and latch the deployment mechanism.
364                 Figure 21. ADEPT Moving Ring Quadrant - Plunger in Latched (left) and Triggered (right) States   Once the plungers have reached their full end of travel, the two latches  that released the plungers are pulled  back against the pins and capture the backside of the large diameter tapered ram p section of the pins  (Fig.  22). This is the primary “locking” mechanism that prevents the moving ring from moving under aero load  (the secondary locking mechanism is when the tips of the nose cap pins are captured).  

Figure 22: ADEPT Plunger Assembly, Latched in Fully Deployed Position
365   Figure 23: Fully D eployed ADEPT Showing Extended P lungers with Captured Pin T ips  Instrumentation   SR-1 contains  a basic flight computer with an Inertia Measurement Unit and a Global Positioning System  (GPS) to measure and record the vehicle’s accelerations, spin rates and position vs. time. In addition, a C - band transponder and antenna allow for range tracking, and a GPS beacon help s aid recovery . A Go -Pro  camera is aimed at the trailing edge of the cloth skirt to observe any flutter  if it occurs and records  a set of  status LEDs (Fig.  24).  The deployment mechanism utilizes two switches and a burn wire. To sense ejection from the spacecraft,  a pair of proximity switches are mounted in the nose cap. The corresponding trigger magnets are mounted  in the sabot that cradles  the nose cap within the payload volume. Once SR -1 is pushed free of the payl oad  module, the switches change state and start  a timer to release the ribs . After 40 seconds (sufficient time to  guarantee no recontact with the payload module or nose cone), the burn wire severs  the Vectra n cord and  the mechanism deploys . At 69° rib angle (out of 70° total), a bypass style micro switch is tripped by the  moving ring, which signifies  “full deployment” . The separation, burn wire power, and “fully deployed and  latched”  events are indicated by LED’s that are visible in the field of view of the camera (Fig.  24). 

Figure 24: View of Trailing Edge of ADEPT Carbon Skirt after Exo- Atmospheric Deployment
366 Testing   This dual spring mechanism design was extensively tested thru multiple prototypes, from a single plunger  of printed plastic, through a 360- degree demonstration unit, culminating in the fabrication and testing of two  flight units .  

The nested pairs of long travel extension springs were vibration tested at the component level, in the ir  extended position, to check for possible resonance at launch vibr ation levels . The entire SR -1 was also  vibration tested after electronics installation. Sine sweeps and three axis random vibration tests were  performed on the stowed SR -1 to simulate the launch environment, followed by a full functional check.  

Functional  testing of the mechanism (separate from integrated testing after the electronics were installed)  included: 1) full deployment within 10 seconds of restraint release, 2) verification of full deployment via rib  angle measurement, 3) verification of deployme nt diameter of 0.7m, 4) measure ment of  gore deflections  under known pressure loads using a vacuum bag  (Fig. 25 ), 5) full deployment confirmation via limit switch,  and 6) plunger assemblies latched/locked after deployment .  

Figure 25: Cloth Deflection Measurement Test using Vacuum Bag   During a simulated payload ejection, the onboard Inertia Measurement Unit was used to estimate SR-1’s  tip-off rates  with respect to the payload module. Both stowed and deployed configurations were tested for  CG location and radial offset, plus Mass Moments of Inertia .  The soft goods presented a significant challenge in that the cloth is stiff and develops a “memory” after  prolonged storage . ADEPT was shipped to the launch provider months in advance of the SR -1 flight, so it  was important to be sure the cloth would unfold after long storage. The damping effect of the stiff cloth  forced the cloth tension to be at the lower end of the allowable range. Adjusting for  maximum cloth  tension  sometimes prevented the mechan ism from reaching full travel and latching, so a balance was required to  meet deflection limits while achieving full deployment . Long- storage tests of 85 and 90 days  were used to  ensure that SR -1 would still deploy.  

Other than some non-deployments due to long storage at high cloth tension targets, t he deployment   mechanism went through hundreds of cycles without other issues . To prevent possible contamination of  the carbon cloth, initially no lubricant was used on the rib and strut pi vots. No lubricant was used on the  plunger or catch assemblies.  The linear slides and linear bushings were pre- lubricated by the manufacturer.
367 Lessons Learned   Many design decisions were limited by the project budget . Pivots were simply aluminum holes rotating on  steel shoulder bolts . The selected springs were chosen from commercial, off the shelf hardware, rather  than custom made for optimal length/spring rate. There were no binding issues associated with the SR -1  shortcut of nesting the two long- trave l extension springs . One of the spring steel kickoff springs broke during  testing and some lost their shape. The heat treat shop had trouble reaching 48- 50 Rockwell hardness, and  suspected the spr ing steel was 1075 and not 1095 as indicated on the material  certifications.  An alternative  to a flexible spring steel design, or more rigorous material certification , could prevent this .  

The carbon fabric  was prone to unravelling at the edges  of the gores , since it has no edge treatment . Carbon  fibers would occasionally get caught in the linkage joints  (Fig. 26). This was mitigated as best as possible  by trimming any stray fibers before each stow.  

Figure 26: Carbon Cloth Fibers Trapped in Joint                        Figure 27: Wear at Pivot Joints SR-1 easily met its mass target of 25 kg (payload mass target of 15 kg) . The entire spacecraft weighed  10.9 kg ( 24 lb) . 0.61 kg (1.34 lb) of tungsten ballast was required at the front of the payload volume to meet  axial (forward) CG location  requirements . There is plenty of opportunity to wring weight from the design,  especially the ribs, rib tips, and moving ring.  

Additional antennae and connectors were required aft of the burn wire assembly, which made the repetitive  threading of the Vectra n cord (required during testing) difficult . Aft mounted antennae or interfaces may be  a regular necessity . Design for accessibility would be helpful because of the need for repeated system  testing.   

After repeated deployments during testing (50+), the mechanism began to fail to reach full deployment and  latch . A small amount  of Polytetrafluoroethylene (PTFE) grease was applied to the rib and strut joints, which  helped  (Fig. 27). For more repeatable deployment, the extra cost and complexity of bushings at the pivot  joints  and plungers  would be useful.  

The Nichrome burn wire assemblies performed flawlessly . This very simple system was well suited for the  SR-1 application .
368 Conclusions   ADEPT was launched from White Sands Missile Range on September 12, 2018. It was recovered after its  planned hard landing and inspection showed that the decelerator deployed and fully latched (Fig. 28). The  on-board video recovered from SR -1 showed a fully  deployed shape and an illuminated  deployment  indicator  LED. Analysis of SR -1’s flight data showed angle of attack oscillations, but that it was stable from  Mach 2.8 to Mach 0.8.  

Figure 28: Post -flight ADEPT  Recovery   Of the ADEPT SR -1 Success Criteria related to the deployme nt mechanism, Criteria D was defined as :   “ADEPT achieves fully deployed and locked configuration prior to reaching 80 km (49.7 mi) altitude on  descent. ” This was verified by evidence from the onboard GoPro video.  

Of the ADEPT SR -1 Key Performance Parameters related to the deployment mechanism, P arameter 1  called for : “Exo-atmospheric deployment to an entry configuration of the 1m-  class ADEPT. ” This was  met  with a confirmed, fully locked deployment to 70 ° rib cone angle prior to entry, as evidenced  by the onboard  GoPro video, and post -recovery inspection.  

This mechanism proved successful and was well suited for our specific ADEPT case:  an Earth re- entry  aero- stability demonstration of a small diameter vehicle, with minimal heating  and a low cloth tension  require ment . Larger applications of the ADEPT concept, or entry environments with high heating or dynamic  pressure, may require a motorized deployment design.   References   1. Smith, B., Cassell, A., Kruger, C., Venkatapathy, E., Kazemba, C., Simonis , K. “Nano- ADEPT: An Entry  System for Secondary Payloads .” IEEE Aerospace Conference, June 2015.   2. Thurn, Adam “Nichrome Burn Wire Release Mechanism for CubeSats .” Proceedings of the 41st  Aerospace Mechanisms Symposium, May 16- 18, 2012.
© 2020 Lockheed Martin Corporation   371 that supports the overall solar array. The third major subassembly is the Blanket Container Assembly. Like  the Deployer Assembly, the Blanket Container Assembly utilizes composite panels with aluminum  honeycomb core and carbon- fiber facesheets. The purpose of the Blanket Container Assembly is to protect  the Blanket Assembly during launch, position the Blanket Assembly for deployment once on orbit, and  provide tension to the blanket during its on- orbit life. The fourth major subassembly is the Blanket Assembl y.  The structure of the Blanket Assembly consists of a thin fil m of Kapton. The solar cells are bonded to the  Kapton in segments called panels. Panels are then connected to each other via hinge pins and Z-folded in  the Blanket Container Assembly to create the Blanket Assembly.  

Deployment  

The MM Solar Array deploys in two phases. The first phase of the deployment starts when the low -shock   non-pyro release mechanisms  that hold down the solar array to the vehicle are released. Once these 8  release devices are fired, 3  hinge lines deploy the Boom Assembly and Deployer Assembly. A f ourth hinge  line rotates the Blanket Container Assembly and positions it for deployment of the Blanket Assembly. The  end of phase one of the deployment is achieved when all four hinge lines lock out.  

                           Figure 3 – Stowed Configuration                                         Figure 4 –  First Stage Deployment  

Phase two of the MM Solar Array  deployment consists of deploying the Blanket Assembly. Like previous  generations of Lockheed Martin flex arrays, the Blanket Assembly is deployed usi ng a central backbone  called a mast. The mast is a deployable composite structure that is connected to the outermost panel of  the Blanket Assembly. When power is applied to the actuators, the mast is deployed and as it deploys, it  pulls the z -folded blanket open. At the end of the mast’s travel, tensioning devices  called the Tension  Mechanisms  are engaged. The Tension Mechanism is a constant force negator spring device that is  mounted to the static structure of the MM Solar Array. A cable is wrapped around a central spool of each   Tension Mechanism and then connected to the most inboard panel of the Blanket Assembly. As the mast  nears its end of travel, the Tension Mechanism pays out a short length of this cable, thus applying a known  tension to the blanket. The Tension Mechanism transforms the Blanket Assembly into a tensioned  membrane, which is imp ortant to the satellite’s guidance and control because if the blanket is not tensioned,
© 2020 Lockheed Martin Corporation   372 then control of the satellite is compromised as it is very difficult to operate a vehicle that has two 23 -meter  (75-foot) long “floppy” appendages. For the remainder of  the mission, the Tension Mechanism is left in this  “partially deployed” state, where it can pay out more cable or retract some cable while maintaining a  constant force. This is important because the Blanket Assembly will expand and contract as the temperature  changes, especially as the vehicle goes into and out of the Earth’s eclipse.  

Figure 5 – Fully Deployed MM Array in the Background, Stowed MM Array in the Foreground