ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
Resul ts and Discus sion 

The re sults of the testing are given in T able 1 and ill ustrated in Fi gure 1. It is notewo rthy that only one of  the nitroge n gas te sts fail ed (2 8.4 x 106 revs). Th e remai ning three te sts u nder nitrog en gas were  stopp ed. We stopp ed on e test at 17.32 x 106 revs prematurely, co mpared to  the other te sts. If we do not  include this result, the av erage du ration for the other th ree te sts is 3 0.9±2.4 x 106 revs. Thus , we  conclude th at the nitrog en gas tests la sted at lea st a factor of 5 lon ger tha n the tests under va cuum   conditions. This is si milar to the result s of our testin g usin g lead naphth enate additive. Und er esse ntially  the same  conditions, th e lead naphth enate test s in nitrogen  lasted at le ast a factor of 1 0 longer than  the  tests i n vacuum [7]. In th e case of  lead n aphth enate, we  believe that the  differences in  wear life are  related  to th e co mpetitio n between oxygen and t he additives for metal atom su rface sites on the  conta cting metal pa rts. In  nitrog en g as, the oxyg en concentration is high  enoug h that o xygen ca n  comp ete with  the additive f or freshly exp osed metal and pa ssivate the su rfaces . On the oth er hand, u nder  vacuum, where the oxygen  concentration is very low,  the additive comp etes successfully for surfa ce site s  with whi ch to react, its co ncentration in  the oil is reduce d at a higher rate tha n it is under nitrogen g as,  and thicke r Pb-containi ng films build u p.   

Table 1.  We ar Test Results  Environment Test Dura tion (revs x 106) Ave.   (revs x106) Failure (Y/N)  Nitrog en ga s 28.4 27.5±7.1 * Y   17.3 (or 30.9±2.4 **) N   33.2  N   31.2  N  Vacuu m 2.45 5.9±3.6 Y   10.79  Y   2.45  Y   6.08  Y   3.92  Y   9.76  Y  *Includi ng  all  tests.  **Excludin g test that wa s stopped p rematurely at 17.3 x 106 rev. 

05101520253035 ac  Total Revs (x1 06)  Nitrogen                                                                                                                      VNitrog en uVacuum  Figure 1.  Wear Tes t Results 

The me chanism of a ction of aryl ph osphate e ster additives is not compl etely understoo d. Aryl phosphate  esters ca n react with oxidize d iron a nd steel  surface s [9,10]  at tempe ratures above 100°C, a nd  pretreatment  of bea ring parts wi th phosphate additives at such tem perature s, leaving  pho sphate-  359
10. D. R. Wheel er and O. D.  Faut, "The Adso rption and The rmal Decomp ositio n of Tricresyl phosphate  (TCP ) on Iro n and Gold," Appl. Surf. Sc i. 18, 106-1 22 (1984).  11. E. S. Forbe s, N. T. Upsdell, an d J. Batters by, "Current Th ought s on th e Mechani sm of Action of  Tricre syl Pho sphate as a Load -Carry ing Additive," Inst. Mech. E ngrs. Tribol ogy Conventio n Papers,  C64172 (197 3).  

 361
The X-38 V-201 Fin Fold Actuation Mechanism 

Christian Lupo *, Brandan Robertson * and George Gafka* 

Abstract 

The X-38 Vehicle 201 (V-201) is a space flight prototype lifting body vehicle t hat was designed to launch  to orbit in the Space Shuttle orbiter payload bay. Although the project was cancelled in May 2003, many  of the systems were nearly complete. This paper will describe the fin folding actuation mechanism flight  subsystems and development units as well as less ons learned in the design, assembly, development  testing, and qualification testing. The two vertical tail fins must be stowed (folded inboard) to allow the  orbiter payload bay doors to close. The fin folding ac tuation mechanism is a remotely or extravehicular  activity (EVA) actuated single fault tolerant system  consisting of seven subs ystems capable of repeatedly  deploying or stowing the fins.   Introduction 

The X-38 Project consisted of multiple unmanned drop test vehicles of various scales and one full-scale  unmanned space flight proto-flight V ehicle 201. The vehicle shape, derived from the X-24, is a lifting body  that lands via a parafoil and skids. The project’s purpose was to perform the development work for an  operational International Space Station Crew Return Vehicle. In order to launch V-201 in the shuttle  payload bay, the two vertical fins must be folded inboard to allow the payload bay doors to close.  

Figure 1.  V-201 (Fins Deploy ed) on Mobile Transport Fixture after Static Testing  Fin Fold Actuation Mechanism Overview   The fin folding actuation mechanism is a complex system consisting of seven subsystems housed in a  structure that narrows to 13.97 cm (5.5 in.) at the highest load point, the mid hinge. The system must  react greater than 11,300 N•m (100,000 in •lbf) at the fold line from aerodynamic loading and seal the fold  line for re-entry from orbit. The mechanism is requi red to be a remotely actuated single fault tolerant  system operated via laptop in the orbiter cabin. The mechanism is also required to include an interface for  unplanned EVA actuation of the fin. Motors are used bec ause hydraulic power is not available on V-201.  The upper fin structure also served as the housing for the electromechanical actuators for the rudder.                                                    * NASA Johnson Space C enter, Houston, TX 

Proceedings of the 37th Aerospace Mechanisms Symposium, Johnson Space Center, May 19-21, 2004  363
This added additional challenges to the design due to t he necessity to accommodate large power cables  across the fold line and around the fin drivetrain. The rudder drivetrain will not be discussed in this paper.  The fins are stowed (folded and locked) for launch and deployed (unfolded and locked) for free flight. The  system is capable of being re-stowed for return in the shuttle and then re-deployed if required. 

The fins are folded via two four-bar linkages t hat toggle over-center in  both the stowed and deployed  positions. In the deployed position, three additional  latches are engaged to transmit load from the upper  fin to the lower fin. All of these operations are ac tuated with a common drivetrain. Each fin drivetrain  consists of a Power Drive Unit (PDU), an EVA interf ace, two commercial off- the-shelf (COTS) torque  limiters, two secondary gearboxes , and various connecting shafts.    

Figure 2.  Ov erview of Fin Actuation  Starboard  Upper Fin  Frame  Structure  Fin Hinge Line Rudder   Electromechanical  Actuato r Rudder  Hinge/Latch Location (1 of 3)  Fin  Actuation  MechanismSecondary  Gearbox (1 of 2)Torque Limiter (1of 2) Power Drive Unit (PDU) The PDU is the combination motor assembly/gearbox that provides to rque to the fin folding and latching  mechanisms. The PDU gearbox consists of planetar y and translation gears that provide the mechanical  logic to allow the fin to be actuated by a prim ary gearmotor, a secondary gearmotor, or an auxiliary  manual/EVA input. A layout of the PDU is presented in Figure 7. 

From the PDU, the torque is sent forward and aft with a torque tube to the torque limiters. The torque  limiters are ball-detent slip clutches that are set to disengage when the torque in the torque tubes reaches  a level that could potentially  damage downstream components.  

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upper rod end is keyed to the sleeve with a N AS1193 locking device engaged in the castellated end of  the sleeve and a jam nut locked with safety wire. The standard rod ends are cadmium plated, which is  generally avoided in space vacuum environments. Ho wever, the JSC Materials and Processes Branch  gave approval to fly these rod ends provided the ex posed cadmium plated surfaces were painted with  Super Koropon Epoxy Primer and the exposed cadmiu m plated threads were sealed with RTV-142.  

The upper fin stow angle is set by two drag links, one forward and one aft of the four-bar linkage. The  drag links are also turnbuckles. The crank and coupl er linkage are housed in a sub-assy module to  minimize assembly on the vehicle.   50º±5º Mechanical  Stow  Indicator  Flag  Hinge Drag Link  (Turnbuckle) Crank Coupler Link  (Turnbuckle) Outboard  Upper   Fin

Figure 4.  Fin Actuation Mechanism Linkage - Stow ed Stowed  Hardstop  Crank  Input  Shaft  Deployed  Hardstop  Setscrew Drag  Link  Guide  As the crank drives the fin to t he stowed position (Figure 4), the coupler link must compress to rotate  overcenter through the on-center posit ion until the crank hits a hardstop.  The link is over-center when the  axis of the pin connecting the coupler and crank is rotated past an imaginary line drawn between the  upper coupler pin axis and the Crank input shaft axis . A cam on the input shaft engages the limit switch  assembly pivot arm which trips the limit switch. T he hardstop is designed to stop the crank motion while  the coupler link is still compressed. This compress ed link preloads the coupler link in the overcenter  position. Inertial and vibro-acoustic forces act on the fin during the shu ttle launch. The direction of force  may be outboard or inboard. Outboard acting forces only result in attempting to drive the linkage further  overcenter. Inboard acting forces must overcome the compression preload in the coupler link and deform  the drag links enough to get the linkage to an unstable on-center position. The compression preload can  be increased by either lengthening the compression link or decreasing the fold angle by shortening the  drag links. Reference SKK51356551 for rigging procedure.  As the crank drives the fin to t he deployed position, the drag links retract into the lower fin along guides  and the upper fin contacts hardstops on the lower fin. Due to space constraints, the deployed position is  not shown; however, components described are visible in Figure 4. The crank continues to drive the  coupler link over-center. Since the upper fin can no longer move, the crank must stretch the coupler link  to rotate through the on-center position until the crank hits a hardstop in the crank assembly module. As  the crank approaches the hardstop, a cam on the i nput shaft engages and trips the limit switch. The  366
The TPS mockups led to redesign of the lower tile hinge fairings so that they were mounted on carrier  panels and could be removed to facilitate fin removal. 

The coupler link calibration worked best when the stra in gage bridge was placed at the center of the  turnbuckle sleeve over the hollow section where neither the upper or lower rod end thread was engaged.  This resulted in smooth linear plots with no jump  between tension and compression. However, in the  second prototype set, the strain gages were locat ed near the bottom of the sleeve. This was done to  improve the wire routing. However, this resulted in a jog between tension and compression. Further, it  was noted that to perform stiffne ss calculations based on the linkage calibration, it was necessary to add  displacement instrumentation bet ween the two pin connections of t he rod ends. Simply using the head  displacement of the tension/comp ression machine did not provide accu rate results. Reference “X-38 V201 Fin Linkage Calibration Report” for the comp lete calibration report on all the linkages.  Fin Deploy ed Position Latch Mechanism  Latch Mechanism Description   The fins are latched in the deployed position via th ree latch mechanisms installed in the three hinge  fittings: forward, mid, and aft. Many naval aircraft use two hinges with two hydraulically fired pins. The  pins are long pins with a tapered nose. This long stroke allows the taper to self-align the lugs with the pin  as it is fired. However, V-201 did not have hydraulic power as an option. In order to meet the single fault  tolerance requirement, it was decided to drive all systems with a single PDU. See the PDU section for  further description of single fault tolerance. This restriction limited the stroke and force that was available  to drive a pin. Initial concepts all focused on us ing three pins. Because the fin environmental seal  required significant force to compress, the pin concept was abandoned in favor of a latch that could  compress the seal. This eliminated the pin and lug alignment issue.   It was decided to use three hinge/latch pairs instead of two for several reasons. The fin fold interface is  approximately 2.13 m (7 ft.) long. The vehicle has thr ee main frames that provide a load path all the way  through the vehicle. Structurally, the main aerody namic load from the rudder transferred down the mid  spar of the upper fin frame. The cleanest load path was to have a hinge/latch pair at this mid location with  a frame lined up with the spar. A single hinge/latch pair forward or aft would not be sufficient to react the  remaining loads. Finally, three latches in conjunction with the two actuation linkages are necessary to  compress the environmental seal and minimize gappi ng at the fold line under aerodynamic loading.  

Upper  Fin  Lug  Coupler  Crank  Input  Shaft Hinge  Latch  Hook  Crank  Figure 5.  Fin Latch Mechanism   368
then drives the sun gear of the second planetary trai n and a manual EVA interface drives the ring gear of  the second planetary train. This allows either gearmotor or the manual EVA interface to independently  drive the overall gearbox. Each of the gearmotors has an electromechanical power-off brake (power to  the gearmotor releases the brake) and the EVA Interf ace is locked when not in use. For overall fin  actuation, the PDU provides full two-fault tolerance against motor failures. The CAD view in Figure 7  displays the geartrain.   Pinion  Gear  Manual  Input  Ring  Gears  Translational  Gears  Output  Shaft  Motors  Figure 7.  PDU  

Each gearmotor assembly is equipped with a power-off br ake that locks the motor shaft when not in use.  This power-off brake is required due to planetary gear kinematics. Not locking the motor shaft of the unpowered motor would result in the powered motor backdriving the un-powered motor as opposed to  supplying torque through the gearbox for fin actuation. The manual EVA interface is also locked while not  in use for this same reason (see EVA Interfac e section). Originally, COTS DC brus h motors  were  selected for use, but several problems were encountered. Problems included improper motor/brake  assembly, unknown and unidentifiable materials usage,  no secondary locking features for fasteners, nonvibration rated electrical interfaces, and inefficient  volume usage. This eventually drove the team to  specify custom flight-certified gearmotors for the final  flight assembly. These flight motors were never  purchased due to program cancellation. Each gear in the PDU is coated with a Dicronite TiS 2 dry film  lubricant, and due to the low rotational speeds their dual-supported shafts are housed in aluminum  bronze bushings lubricated with Braycote 602EF, with the exception of the planetary carriages which  were supported by thin-section ball bearings. Because hi gh rotational accuracy of the PDU output shaft is  not required, a small backlash was designed in and the center distance for all gearing was fixed without  provision for adjustment.  PDU Testing   The PDU underwent many informal f unctional tests after assembly, during which time it was discovered  that five of the six COTS motors had been sent from the vendor incorrectly asse mbled, causing them to  seize after about a minute of running. Seizure was due to motor operation with a partially engaged brake  that would heat and swell until torque/current cut-o ff. A work-around was implemented allowing the PDU,  torque limiter, and secondary gearbox to undergo therma l-vacuum testing. The PDU worked as designed  during and after all phases of testing. Later, a fin- level random-vibration test was conducted and the PDU  passed this test with no problems as well. 

PDU Observations   370
Figure 9.  Cam/Roller/Linkage Timing Mechanism  The design was completed using solid models which were used for preliminary kinematic analysis. A  detailed dynamic analysis was then performed. The strength analysis was performed with Mathcad based  on the dynamic analysis results. An initial unit was then manufactured so that it could undergo operational  proof testing. The proof test was planned to run to 100% of design limit load in the forward and reverse  direction three times each, but during the initial cy cle, the unit failed in the reverse direction. Upon  opening the test unit, it was discovered that the pi n upon which the roller rotated had sheared and the lug  in which the pin rested experienced an ultimate beari ng failure. In addition, other pins and holes in the  linkage had yielded to various degrees.   Upon closer examination of the failu re, it was found that the load applied to the linkage was not consistent  with the magnitude or direction predi cted in the dynamic analysis. An in vestigation of the dynamic model  uncovered a small error in the application of the motion driver in the reverse direction that resulted in an  incorrect load condition. Correction of the error produced new load vect ors consistent in both magnitude  and direction with that seen by the failed linkage.  The mechanism was modified in an attempt to correct the deficiencies in the or iginal configuration by  using a stronger material for the pins and links. Analysis indicated that this would result in very low, but  positive margins in the pins so the test was r epeated with the new pins. No failure occurred but  subsequent inspection showed slight yielding in the pi ns was still occurring. The design was more broadly  modified to increase the pin diameters slightly to  increase their load capability. This was a design  challenge given the tight confi nes but a solution was found.  The new design was then retested to the same loads . No failures were experienced and the mechanism  operated as designed. A post-test inspection of the internal mechanism revealed no problems.  

Timing Mechanism Lessons Learned   • The simplest concept doesn’ t always work. Mechanisms that work kinemat ically in the CAD system  may not work when under the loads and friction of the real world.  • Be careful with your analysis . Be s ure to review all as sumptions   • Always t est!  Analysis may not be sufficient, and testing may be the fastest way to iterate a design. 

Drivetrain Testing   The test unit shown in Figure 9 was thermal/vacuum tested. The sub-assemblies were later used in the  fin engineering unit and have continued to operate reli ably without incident. Reference “X-38 V-201 Fin  Drivetrain Thermal/Vacuum Test Report” fo r description of thermal/vacuum testing.   373
Torque Limiter 

 Gearbox  Power   Drive   Unit  Timing  Mechanism  Figure 10.  Driv etrain Test Unit  

Limit Sw itch Pair Assembly  

Limit  Switch Pair Assembly  Descript ion  In order to indicate successful stow or deploy, fi ve limit switch pair assemblies are used. First, the  assembly positions the two limit sw itches side-by-side so that a singl e cam actuates both switches. The  pair is required to meet the single fault tolerance requi rement for remote actuati on. Second, the assembly  is designed to prevent damage to the switches. T he pivot arm is spring loaded to engage the switches  which are off-loaded as the cam engages it. In this way, over-travel of the cam could not damage the  switches. See Figure 11.   Cam engages here  Setscrew pivots    away from switch   as cam engages   setscrew above.  Figure 11.  Limit Sw itch Pair Assembly  

Limit Switch Pair Assembly Lessons Learned   Attem pting to m eet requirem ents for two locki ng features can result in other problem s.  • The pivot arm has small # 6 setscrews (3.5 mm) m ade of 304 stainless which c ontact the limit switch  arm in order to trip both limit switches simultaneous ly. The setscrews are locked in place using a selflocking insert with a self-locking jam nut. Howe ver, the 1.5 mm hex in the setscrew would  occasionally strip while attempting to overcome running torque. Later versions of the assembly  374
The X-38 V-201 Flap Actuator Mechanism 

Jeff Hagen*, Land on Moo re**, Jay Estes**, and Ch ris Layer+ 

Abstract 

The X-38 Crew Rescue  Vehicle V -201 space flight test a rticle was designed to achieve a n  aerodynami cally controll ed re-entry fro m orbit in part through the  use of two b ody mounted  flaps on the   lower rear side. These flap s are actu ated by an elec tromechani cal system that i s partially exposed to the   re-entry environment. The se actuator s are of a nove l configu ration and are uni que in their requireme nt to  function while  expose d to re-ent ry con ditions. Th e aut hors are not aware of any  other vehi cle in whi ch a  major a ctuato r system was requi red to f unctio n throu ghout the co mplete re -entry profile whil e part s of the  actuato r were directly expo sed to the am bient enviro nment. 

Introduc tion 

The X-38 Project consisted o f multiple unmanned drop test vehi cles of vario us scale s and on e full-scal e  unman ned space flig ht prot o-flight V ehicle 201 (V-201). The vehi cle shape, de rived from th e X-24, was a  lifting body th at land s via a  parafoil and  skids (Figure 1). The  purpose of the p roject was to perfo rm the  developm ent work for a n operatio nal Internat ional Sp ace Station Crew Retu rn Vehicl e. 

The config uration, shap e, size, a nd fu nction  of th e X-38 co ntrol surfa ces (rud ders and bod y flaps) were   chosen to mi mic that of the X-23 / X-24 a s closely as p ossible. The X -23 re-entry b odies had fixe d  rudders an d wed ge shap ed body flaps t hat repli cated only the lower surfa ce of the X-24 flap s whil e filling  the othe rwise void area b etween th e top  side  of the fl aps and th e lower skin of the aft fuselag e. The X-38   flaps mo re closely re semble the flaps o f the X-24 low sp eed te st vehicle, with  a void area betwe en th e  lower skin of the aft fusel age and the top surfa ce of the relatively thin flap. Fl ight control required the  capability to drive the fla ps with up to 28,811 N-m (255, 000 inch-pounds) of torqu e - including ma rgins,  and to rotate throug h an a ngle from zero to forty-fi ve degree s from  the lower side of the aft fu selage. 

       V-201  Actuato rV-201 Figure 1.  X-38 F amily of Vehicles (V131R,  V201, V13 2)         Figure 2.  Actuator (foreground) and V201 

                                                  * Lockhe ed M artin Spa ce Operatio ns, Ho uston, TX  ** NASA J ohnson Space Center, Houston, TX  + Moog, In c, Motion Syste ms Divisi on, East Auro ra, NY 

Procee dings of the 37th Aeros pace Mec hanisms Sy mpos ium, Joh nson S pace Ce nter, May 19-2 1, 200 4  377