ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
The SRS's of the shock pulses are shown in Figure 6 and the time histories of the pulses are shown in  Figure 7.  The response of the payload mass was measured with accelerometers at the locations shown in Figure5.  The peak response of the mass was measured at the channel 8 (in axis) accelerometer location. The  peak response from the shock pulse was peak input value of 805 g's and a peak response of the payload  mass was 21.2 g's. This was a reduction of 31 dB. The amount of isolation due to base shock input can  be seen by viewing the base input time history on the same plot as the mass response as plotted in  Figure 7.  The SRS of the 18.1-kg (40-lb) mass to both of the shock pulses is shown in Figure 6. The SRS of the  response reflects the fact that, in the higher frequencies, the first input shock pulse is higher and this is  reflected in the response of the mass. One noticeable artifact of the mass response in Figure 6 is that  there is a peak between 560 Hz and 630 Hz even though the break frequency of both pulses is around  800 Hz. The 630 Hz peak is very close to known surge frequency of the isolator main springs. The sine  vibration data only goes to 2 kHz; therefore, the correspondence to the peaks in the SRS data can only  be tracked to 2 kHz.  Conclusions about the Shock Beam Test Results  The following conclusions can be made about the test results.  0 The shock beam test was able to achieve the input levels required by the potential isolation  system.  The isolators in the bipod configuration were able to decrease mass responses relative to the  maximum input by 30 dB or more. These Isolators eat shock!  The isolator surge modes do contribute to the response of the payload mass due to the base  shock input; but, the contribution is not as great as previous testing indicated.  10  -Reference - - -LowerLimit - - -Upperliml -  -Beam Shock Input (g), Control Chn 1,4W. P1 -Beam Shock Input (g). Control Chn 1,4oW. Pa -Beam Shock Response. (9). Chn 8.40#. Pi - Beam Shock Response. (a). Chn 8.40% P2  Figure 6. Bipod 18.1-kg (40-lb) Mass response SRS, both Pulses  The isolated calculated SRS of the payload mass is significantly lower than the input SRS. The spring  surge mode at around 620 Hz can be observed in the above plot.  145
Dstrutm Bi-Pod Shock Beam Test, Both Pulses, 40 Ib Mass,  Base Input compared with Mass Response, Time History  800  700  600  500  -100  -200  -300  -400  -500  0 5 10 15 20 25 30 35 40  Pulse Time (rn-sec)  I-Shmk Beam, CHN 1, 4M, Pulse 1 -Channel 8, Pulse 1. 40 Lb Mass, Time History  Figure 7. Comparison of Shock Input and Mass Response, 18.1-kg (40-lb) Mass, Shock Beam Test  The time history for the recorded shock pulses is shown along with the mass response. This plot shows  the amount of reduction in the response of the bipod mass due to the isolation of the shock input. The  effective reduction of the mass response due to the base input is over 30 dB.  Sine Vibration Test  Sine Vibration Test Setup  The test setup for the sine and random vibration test is shown in Figure 15. The picture in Figure 16  shows the actual setup of the sine and random vibration testing. The bipod and mass were instrumented  with 15 accelerometers. The accelerometers were placed to measure the response of the isolators and  payload mass with-in the limitations of the available instrumentation. The locations of the accelerometers  were selected to provide information of the in-line spring surge modes, the isolator spring body lateral  modes, and what the mass response was at those modes. The locations of the test accelerometers are  shown in Figure 8. The testing was performed with only the 18.1-kg (40-lb) mass. This was done to  characterize the equivalent system bounce mode and to characterize the isolator spring body lateral  mode, spring body in-line surge mode and harmonics, which are not affected by the bipod suspended  mass.  Figure 8. Sine and Random Vibration Accelerometer Location, South Side  The north (top) and south (bottom) isolator with the location of the visible accelerometers are shown. The  slip plate and the adapter plate are to the right with the suspended 18.1 -kg (40-lb) mass to the left. 'I  146
Transfer Function, 5Hz to 2KHr Sine Vibration Test of Bi-Pod with 40Lb Mass,  Mass response at Channel 8 only Relative to the Base Input  I I IIIII  I I IIIII T--T i 1-rii  I I IIIII I 1 IIIII I I IIIII  100 1000 10000  Frequency (Hz) 10  1-8tol,lst.25g -8to1,2.5g’s -8to1,7.5g’s -8to1, 12g’s -8101 (16.25g’s) -8tol,Final.25g’s 1  Figure 9. Transfer Function of the Mass Response only at the Bipod Interface.  Since the in-line response of the mass is much the same in both accelerometers, the plot in this figure is  just the mass response at the bipod interface to the 18.1 -kg (40-lb) mass. These Isolators work so well  that for low inputs (0.25 g’s), the high frequency response is less than the accelerometer noise floor  Final Conclusions  A new D-Strut@ isolator has been developed for a Laser Communication Terminal experiment. This new  D-Strut@ isolator advances the state of the art for dual mode isolators in the space environment. This new  D-Strut@ isolator increases launch load capacity by at least a factor of two with the limiting factor now  being the structure and not the damper bellows as in previous designs. This new D-Strut@ isolator was  exposed to a sine vibration environment of 16.25 g’s and still functions properly afterwards without  changing it’s isolation capabilities. This new D-Strut@ isolator reduces the shock environment by a  minimum of 31 times (-30 dB).  A customer concern with the new D-Strut@ isolator is the main spring surge mode at 615 Hz. This mode  may affect the line of sight pointing capability of the Laser Communication Terminal instrument. To solve  this problem we have completed testing using constrained layer damping, or the using Tuned Mass  Dampers on the spring to reduce the main spring surge mode. At the time of this writing, we have not  completed the evaluation of the test results but firs indications is the constrained layer damping has no  effects on the surge mode and tuned mass damping has significant effects. We will have some  conclusions about the additional damping methods available in a future paper.  References  1 “Performance of a Launch and On-Orbit Isolator”, Jim Boyd, T. Tupper Hyde, Dave Osterberg, Torey  Davis; Presented at the Smart Structures and Materials Conference; SPIE, March 2001  2 “Advanced 1.5 Hz Passive Viscous Isolations System”; Porter Davis, David Cunningham, John Harrell;  presented at the 35th AlAA SDM Conference, April 1994.  148
Spring Mechanism Bolt Catcher Shell  Retraction Nut &  Retraction Spring  .-.-.-.-.-.-.-.-.-.-  Separation Plane  Figure 1. Discrete Point Separation System Components  The heritage component of the separation system is the Spring Mechanism. The Spring Mechanism has  extensive flight history and was also used in the previous iteration of this separation system design. The  Spring Mechanism’s primary purpose is to provide the separation tip-off force from the booster vehicle. It  delivers this separation force by utilizing a spring-loaded plunger configuration of which there are nine  configurations capable of 225 - 525 N of force. The Spring Mechanism also has adjustment capability that  allows each unit to exert an exact force and allow for mounting adjustments. These design features allow  the Spring Mechanism to be used in a variety of applications and meet a wide range of requirements with  only a slight impact to the weight of the unit.  Several years ago, NASA and ESA expressed a need for an alternative to traditional explosive actuators.  NEA Electronics Inc. satisfied this objective by developing a reliable, fast-acting, sure release, low shock  output, redundant, non-explosive separation mechanism called a Split Spool Release Device. The SSRD  is an electromechanical separation nut-release mechanism that eliminates the residue and shock  produced by explosive devices and are factory refurbishable for extended use. The device has significant  flight heritage primarily as a hold down and release mechanism of deployable structures (solar arrays,  antennas, etc). This is the first spacecraftlbooster discrete point separation system design utilizing the  SSRD as the restraint device. The SSRD is a mechanically and electrically redundant and in this  application is capable of 55.6 kN of preload. The typical generated shock output of a single device at  maximum preload is less than 500 G’s. Table 1 shows a brief list of capabilities of the device used in the  discrete point spacecraft separation system.  150
Table 1. Specification for the 55.6-kN Split Spool Release Device  Requirement  Ultimate Load  Max. Rated Release Load  Source Shock  Actuation Circuit  Actuation Time  Qual. TemDerature Ranae Capability  69,500 N  55,600 N  400 G’s Q 55,600 N preload  4 A Q 6 VDC per circuit  25 msec max.  -60°C to +105”C  The device consists of a load-  bearing spool that is split in half.  These spool halves capture a  spherical rod end and represent  the primary load path of the unit.  The spool halves are wrapped with  a restraining wire that keep the  spool halves from moving laterally  when load is applied to the  spherical rod end. The restraining  wire is fixed to one of the spool  halves on one end and restrained  with two fuse wires on a toggle on  the other. The restraining wire puts Figure 2. External View of 55.6-KN 33~~ a small preload on the-fuse wires  when the unit is “set”. Upon receipt of the actuation signal, the fuse wire heats up which decreases its  tensile strength. The toggle breaks through one or both of the fuse wires depending on the type of  actuation signal and the restraining wire starts to uncoil from around the spool halves. The unrestrained  spool halves are then gradually driven apart by the preload until the rod end is no longer captured and  “drops” through the device. The actuation of this device is very unique in that it gradually relieves the  preload, which is the biggest contributor to shock output, and releases the constrained end of the joint.  Refurbishment of the unit is achieved by replacing the restraining wire and fuse wire assemblies after  actuation and can be done numerous times to allow for ground test. In the Discrete Point Separation  System, the SSRD assemblies are jettisoned with the Booster Adapter Assembly after actuation. Figures  2 and 3 show external and internal views of the 55.6-kN SSRD.  Fuse Spherical Rod End  u Restraining Wire  Figure 3. Internal Components of a Typical SSRD  151
Four configurations of the aluminum honeycomb were tested in order to find the configuration to minimize  the shock output. Each honeycomb configuration went through two test runs and the data was evaluated.  Table 2 lists the aluminum honeycomb configuration, and Table 3 lists the maximum shock levels at the  three accelerometer locations, and crush depth in the honeycomb due to the rod of each configuration.  -  Density (kg/m3) 16.0 32.0 110.5 104.0  Cell Size (mm) 9.53 4.76 4.76 9.53  Crush Strength (kPa) 172.4 482.6 3930 351 6.3  Foil thickness (mm) 0.018 0.018 0.064 0.064 Table 2. Aluminum Honeycomb Configuration  Data  Information  I 1X Honeycomb Configuration  -007 I -013 1 -015 I -017  Honeycomb Configuration  -007 -01 3 -01 5 -01 7 None  167.631 27.025 1240.894 1233.103 622.312  1 Y  1 Z  2X Table 3. Aluminum Honeycomb Shock Data  138.727 17.238 1465.1 70 11 17.588 382.61 6  233.268 29.820 2143.589 1270.930 51 6.098  170.565 18.846 1367.840 1238.71 9 51 6.545  1- -1 22  3X  3Y 257.538 23.989 1863.343 1664.746 431.61 1  41 0.757 289.466 1985.1 65 1739.393 1635.51 6  1090.883 530.638 1321.61 0 1538.292 191 5.1 43 Run I 2~ i 101.467 i 11.751 j 787.916 j 564.251 j 239.809 1 IG'S\  The data collected during the development testing showed that the -01 3 configuration honeycomb had  the best shock absorbing characteristics. This was based on data collected on Accelerometers 1 & 2  which were mounted at the base of the Bolt Catcher Can. As the data in Table 2 shows, the shock levels  using the -013 configuration honeycomb are much lower than any of the other configurations for all test  runs. Data collected at Accelerometer 3 was used for reference purposes but is also lower for the -013  configuration in all but one test run. It's interesting to point out that when compared to no honeycomb the  levels significantly lower but the levels seen for the -015 and -017 configuration were higher. This was  154
possibly due to the honeycomb increasing the shock impact surface area. These two configurations had a  significantly higher stiffness which lends some validity to the observation.  Another testing challenge occurred during the component qualification testing of the SSRD. Since this  was the first time an SSRD was used in a discrete point booster separation system, the shock output of  the unit needed to be characterized fully and extensive generated shock testing was performed. The test  setup consisted of a 75-cm square aluminum plate that had a flight-like interface. The interface included a  flight like cup-cone interface as well as the A-frame bracket. The SSRD was mounted to the A-frame  bracket and the preload was applied through the center of the aluminum plate. The test was performed  with the flight designed Retraction System and the crushable honeycomb shock absorber. Tri-axial  accelerometers were used to collect data at locations 7.5 cm, 15 cm, and 30 cm from the center of the  plate. Figure 6 shows the test setup for generated shock testing.  45"\  \ J  L  Figure 6. Generated Shock Test Setup  The design of the test setup was meant to mimic the exact flight configuration in the preload path;  however, for conservatism, the large aluminum plate was incorporated to collect shock levels at varying  locations and did not represent the flight honeycomb panel configuration. This created a major challenge  during the component-level shock testing because configuration as well as accelerometer location had  some bearing on the data.  In addition to the component-level generated shock tests, there was a system-level separation and shock  test. The purpose of this test was to demonstrate the separation system and to measure the shock output  of the system. The BAA was attached to the spacecraft with all the component parts of the separation  system. The spacecraft bus was instrumented with 255 tri-axial accelerometers to measure the shock at  various locations on the spacecraft bus. The entire assembly was suspended from an overhead crane for  the separation test. The SSRD's were actuated in pairs that were diametrically opposed from each other.  Upon actuation of the last pair, the BAA separated successfully from the spacecraft and fell into padding  below the suspended spacecraft. Data collected during the three actuation events showed shock  readings well within specified limits for the spacecraft.  155
Faying Surface Lubrication Effects on Nut Factors  Deneen M. Taylor* and Raymond F. Morrison**  Abstract  Bolted joint analysis typically is performed using nut factors derived from textbooks and procedures from  program requirement documents. Joint specific testing was performed for a critical International Space  Station (ISS) joint. Test results indicate that for some configurations the nut factor may be significantly  different than accepted textbook values. This paper presents results of joint specific testing to aid in  determining if joint specific testing should be performed to insure required preloads are obtained.  introduction  During review of International Space Station analysis documentation and build paper it was discovered  that one of the International Partners had overtorqued the installation bolts on the Common Berthing  Mechanism (CBM) Ring on each of their elements. The build paper also raised questions concerning  lubrication and whether or not it had been applied to the faying surfaces of the bolt head/washer in  addition to the threads. The CBM is a mechanism built and qualified by Boeing and furnished to the other  partners for use with the pressurized elements. Too high of a preload could overload the bolts when  combined with on-orbit thermal and mechanical loads and potentially cause failure. Too low of a preload  could cause the joint to gap and leak atmosphere overboard on orbit. Due to this out of configuration  condition, it was decided that joint specific testing would be run to quantify the effects of the torque and  lubrication.  Background  The testing was implemented to determine the nut factor, K, of three different bolt materials, A286,  MP35N and lnconel in combination with A286 helicoil inserts and nuts, with and without lubrication on  faying surfaces. To remove any uncertainties of previous contamination to the bolts, all bolts were cleaned  before testing. All bolts are about 6.3-mm (exactly I4 inch) diameter. Countersunk CRES washers were  placed under bolt heads and A286 passivated flat washers under the nuts. Helicoils and nuts came from  the manufacturer with standard locking features and lubrication. Lubrication applied during testing to bolts  was Braycote 81 52 oil. The bolted material is anodized aluminum. The different permutations of the tests  are shown in Table 1.  Although the objective of this testing was to verify that the higher torque did not yield the bolts and to  obtain actual nut factors for these specific configurations, due to the numerous tests and various  configurations some interesting trends became apparent. It is important to note that the lubrication was  added between cycles not just for the first cycle, reflective of the flight hardware. In addition, the max/min  nut factors cited are inclusive from all previous cycles, not just the max and min from the noted cycle.  * NASA Johnson Space Center, Houston, TX ** Boeing, Huntington Beach, CA  Proceedings of the 38th Aerospace Mechanisms Symposium, Langley Research Center, May 17- 19,2006  157
A286 CRES Bolt, CRES helicoil, CRES countersunk washer  0.6  0.5  0.4  Y i B  0.3  LL  0.2  0.1  0  0 / /  /--- - -Lube Thread only, K rnax  Lube Thread & Head, K mir - -Lube Thread & Head, K ma  / / -.-  2 4 6 8 10  Cycles  Figure 1. Comparison of test cases 1-1 and 1-2. 12  ,  i ..  Figure 2. Photo on left from Test Case 1-1, max preload 15 kN (3400 Ib)  Photo on right from Test Case 1-2, max preload 8.0 kN (1800 Ib)  159
0.6  0.5  0.4  Y  0.2  0.1 MP35N Bolt, A266 CRES nut, CRES countersunk washer  Lubrication on Threads Only  /- / /  ~ .c  0 - -Tghtened via Nut, K max  -Tightened via Bolt, K rnin - - Tahtened via Bolt. K max  -I-- -++-  -+-e- -------  C  0 2 4 6 8 10 12  0.6  0.5  0.4  Y  5 0.3 L s  0.2  0.1  0 Cycles  Figure 3. Comparison of test cases 2-1 and 2-2.  MP35N Bolt, A266 CRES nut, CRES countersunk washer  Bolt lightened via Head  ~~ -.-- ____ ~- ._ ~---- __.- - ~~~ ------- ~~~ . /-  - -Lube Thread only, K rnax  Lube Thread & Head, K rnir / /  ~~  0  0 2 4 6 8  Cycles 10 12  Figure 4. Comparison of test cases 2-2 and 2-3.  160
0.6  0.5  0.4  Y  0 0.3 U  z ii  L  0.2  0.1  0  0  0.6  0.5  0.4  Y  L-  2 0.3 B  0.2  0.1 Inconel718 Bolt, A286 CRES nut, CRES countersunk washer  Lubrication on Threads and Faying Surfaces of Nut  Tightened via Nut, K min  Tightened vla Nut, K max  -Tightened via Bolt, K min  __ - _._  0 0 -. 0 0 0 0 0 0 0 0 0 // // 0 # / ------------- 0  -4--  _. .  1 2 3 4 5  Cycles  Figure 5. Comparison of test cases 3-1 and 3-2.  Inconel718 Bolt, CRES helicoil, CRES countersunk washer  L - c-  - -Lube Thread only, K max  Lube Thread 8 Head, K min  Lube Thread 8 Head, K ma -e-- /- 0 / / /  0  0  0 1 2 3  Cycles 4 5 6 6  Figure 6. Comparison of test cases 4-1 and 4-2.  161
Torque Loss and Stress Relaxation in Constant Torque Springs  Robert W. Postma^  Abstract  Constant torque springs are manufactured from spring steel strip that in some applications is stressed  beyond the yield strength to achieve maximum torque-to-weight ratio. An adverse consequence of a high  state of stress is torque loss resulting from stress relaxation, which typically occurs over a prolonged  period of time, accelerated by thermal cycles or continuous elevated temperatures. This poster paper  discusses a case of torque loss resulting from thermal cycling of a spacecraft hinge spring manufactured  from Type 301 corrosion resistant steel strip, cold worked to the extra hard condition.  The equations governing the design of the constant torque (Neg’ato?) spring are reviewed. Included is a  discussion of ongoing work to better understand the design, manufacturing, and stress relieving of  constant torque springs, particularly in regard to stress relaxation and delayed cracking from sustained  high stress levels in aerospace environments.  Introduction  Constant torque springs are sometimes stressed beyond the yield strength, by design, in order to obtain  maximum torque-to-weight ratio. The springs are initially fabricated as a tightly wound coil of steel strip,  and may have multiple laminates. The material is stress relieved, possibly before and after forming into  the coil, at temperatures up to 425°C. When installed on the hinge, the spring is reverse flexed from the  take-up spool onto the spool driving the output shaft (Figures 1 and 2).  Stow Direction  (Shown in Deployed Position)  carp. I  Figure 1.  Two Constant Torque Springs  Installed in 90-degree Hinge  The Aerospace Corporation, El Segundo, CA Figure 2.  Constant Torque Spring  Critical Dimensions  Proceedings of the 3dh Aerospace Mechanisms Symposium, Langley Research Center, May 17- 19,2006  163
This work was initiated following loss of torque in spacecraft hinge mechanisms after a number of  temperature cycles during thermal vacuum testing. The torque loss at the end of the deployment motion  for one of the 180-degree hinges was 10 percent, as shown in Figure 3. (For a 90-degree hinge, as  pictured in Figure 1, the torque loss was 7 percent). Note that the torque at the stowed position remains  unaffected. Evidently, it is the slope of the deployment curve that is affected by the thermal cycling during  this particular test. On tests of other hinges, however, the slope remained approximately the same, but  the mean torque level decreased.  Empirical Analysis of Torque Loss  After considering a number of other possible explanations, it was concluded that the 10 percent thermal  cycling torque loss was due to stress relaxation. The most likely other explanation, increase in  interlaminate friction, was eliminated by a special test using only one laminate. (The torque loss was  about the same amount). Since stress relaxation is a function of time, temperature, and stress level, it can  be further surmised that the 37°C elevated temperature portion of the thermal cycles, combined with high  bending stress levels, are the primary factors contributing to the observed torque loss.  The Figure 3 data also shows 5 7 percent hysteresis and an initial negative slope of 15 percent of the  maximum torque. It is also surmised (tentatively) that these adverse effects can be attributed in some  measure to high bending stresses.  I I  I 501 Courtesy  Lockheed Martin  Gorp. (1 in4b = 0.1 13 N.m)  -190 -180 170 -160 150 -140 -130 -120 -110 100 -90 -80 -70 60 50 -40 -30 -20 -10 0  Angle (dog)  Figure 3. Torque vs. Angle for 180-degree Hinge  As the torque loss analysis proceeded, calculations were made that indicated that the springs were  stressed beyond the yield strength (discussed in the following section). This conclusion is supported by  examination of springs used for development testing, shown in Figure 4. Yielding of the spring stock from  the high operating stress and strain levels after being reverse flexed is evident from observed distortion of  the springs. The material is Type 301 corrosion-resistant steel, 0.0229-cm (9-mil) thick, cold worked to the  extra hard condition (60 percent cold reduced). It is seen from Figure 4 that after being removed from the  hinges, the springs are no longer in a tightly wound coil, but are expanded out of shape into a loose  spiral. It can be concluded that being stressed beyond yield, combined with thermal cycling, contributes to  stress relaxation and the associated reduction of torque.  164