ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
Figure 5: Orbital Express vehicles, conceptual drawings  Significant Changes from Initial Prototype  Mechanism Capture Capability During the initial prototype design, a clear definition regarding the capture capability envelope did not  exist. In light of the emerging requirements, the prototype capability was not sufficient to meet the requirements of the OE program. Therefore, the capture capability would need to increase in order to accommodate the precision of the guidance and navigation controls. The required capture capability would be achieved from scaling up the design, with minimum impact to the overall mechanism configuration. In short, the mechanism size and stroke were increased to create a much larger “capture zone”. Mechanism Structural Capacity The initial prototype had also been developed for use with a microsat, and its structural capability was not  suited for large vehicle payloads. Thought was given to refining the load paths in order to handle the increased strength and stiffness requirements. The drive train also needed to be enhanced in order to achieve the necessary mated interface stiffness. As a starting point, it was decided that the capacity  would be designed to accommodate the largest operational vehicle payloads encompassed in the scope of the OE program.  Mechanical Docking System (MDS) From a functional standpoint, the design that emerged looked very much like a scaled-up version of the  original hardware. The design had nearly doubled with respect to the prototype. However, the design had not only grown in size, but had matured in complexity to accommodate its enhanced capabilities, both internally and externally. Figure 6 displays the second generation prototype hardware, officially designated MDS. The requirements that encompassed the design effort are displayed in Table 1.  MDS Testing  In-house Development Testing SRC completed in-house testing of the MDS design that contributed additional information for the parallel  effort of the OE flight unit design. In house testing included stiffness characterization, off-load testing, and  mated interface load. All tests performed proved the MDS unit design accomplished the main goals of the  program. Computer Simulation Latch arm stiffness measurement was one of the several tasks undertaken on the MDS hardware. The  lateral and radial stiffness of the linkage assemblies had been measured in both the deployed and captured position. By considering these measured stiffness values together with analysis of dynamic  211
simulation results, higher fidelity results and conclusions were produced. Dynamic Analysis and Design  System (DADS) software was used to model the capture and retraction features of the MDS. The model included complete mass properties of two satellites, a zero gravity environment, and contacting features of the MDS.  In excess of 400 cases were simulated to explore performance issues using a uniform distribution of  initial condition parameters. The analysis concluded that the mechanism was always capable of positive  captures given reasonable limits of relative initial test conditions.  Passive Side Active Side  Figure 6:  MDS, second generation prototype  Table 1:  MDS Phase II Design Requirements  Parameter Value Axial Capture Distance: 15 cm  Angular Capture Misalignment Tolerance  Pitch/Yaw  Roll±5 degrees  ±5 degrees  Lateral Misalignment Tolerance: ±5 cm  Linear Contact Velocity Tolerance: 3 cm/s  Preload: 11250 N Capture Time: < 10 s  Capture and Latch Time: < 240 s  Interface Outer Diameter: < 46 cm  Active Side  Mass: < 23 kg  Passive Side Mass < 11.5 kg  6 Degree of Freedom (6 DOF) Tests The main goals of these tests were to correlate/verify the dynamics analysis model and to demonstrate functionality and performance of the hardware in a flight-like environment. 6-DOF testing was performed with the test equipment as shown in Figure 7. The test simulated full, relative six degrees of freedom with the actual hardware in the loop. During testing the active half of the hardware was mounted to a  moveable platform supported by six hydraulic actuators. The passive half was mounted above the active half at a fixed location with a force/moment sensor in the load path. Software commanded the six legs to  212
start the test at a user-specified set of relative initial conditions between the active and passive halves.  Once the test was started, software simulated the complete relative dynamics of the two vehicles due to the real contact between the active and passive halves during the capture sequence. The test results were then compared to the dynamic model simulation predictions. Comparison of the test data confirmed the model accurately depicted the capture dynamics. The information gained from the MDS tests was critical to the flight design effort. Figure 7:  6-DOF testing facilities at MSFC  Docking System Mechanism Description  ollowing are more detailed descriptions of some of the significant features of the docking mechanism F final design as it was flown on the Orbital Express program.  Figure 8:  Docking system major components and schematic layout, as configured for Orbital  Express 213
Orbital Express Capture System (OECS)  The capture system consists of an active side and a passive side. The active side contains the grappling  arms and the drive system; this side would normally be part of the supply spacecraft. The passive side  provides capture features and a sensor to indicate proper engagement of the grappling arms; this side would normally be a part of the client spacecraft. Figure 8 shows a general schematic of the system configuration. Capture Sequence To understand the OECS unit, a description of its functionality is necessary. Figure 9 displays the typical  steps involved in a capture operation.  1. The linkages of the active mechanism begin in a deployed (open) state. The passive structure is  held in a station keeping envelope within the capture capability of the active mechanism.  2. Upon receipt of command, the motor begins to actuate a ball screw which translates an internal  piston along the length of the active mechanism canister. The piston moves three separate four-bar linkages that make up the grappling linkages. The linkages then move downward over a roller feature. The camming action of the roller causes the linkages to constrict, engaging the passive structure. The wedge shaped architecture of the passive structure guides the tips of the linkages into center grooves. Capture is achieved as the passive structure is constrained within the bounds of the linkages.  3. As the linkage tips move down along the grooves, they engage a shelf feature, allowing the active  and passive structures to be drawn together. As retraction proceeds, alignment occurs in stages. Continued motion of the linkages causes the interface plate of the passive structure to contact push-off rod struts that act as a three point alignment mount on the active mechanism. The alignment features each consist of a spring loaded pin with conical ends that seat into tapered cups on the passive structure. These features provide a gross alignment. With further retraction, additional features align the electrical couplers within their allowable tolerance.  4. At the final stages of retraction, the passive structure becomes fully constrained by a final set of  cup/cone features. Rigidization then occurs as the motor applies the necessary preload to provide the  required interface stiffness.  1 2 3 4 Figure 9:  Orbital Express Capture Mechanism Capture Sequence  214
Top Plate AssemblyGrappling LinkCone Mount, 3Plcs age , 3Plcs  Pu(shown retracted) Canister Assembly Motor, Gear Train, Brake Assemblysh-off Rod Assembly, 3PlcsFloating Electrical Connectors, 3Plcs Top Plate AssemblyGrappling LinkCone Mount, 3Plcs age , 3Plcs  Pu(shown retracted) Canister Assembly sh-off Rod Assembly, 3Plcs Motor, Gear Train, Brake AssemblyFloating Electrical Connectors, 3Plcs Figure 10:  Active mechanism cross-section  Active Mechanism The Active mechanism is composed of a top plate assembly to which the canister assembly is attached  igure 10). Within the canister assembly is a drive train consisting of a motor, ball screw, and  ke up a four bar connection. Mounted to the canister is a reaction roller that  guides the linkages through deployment. Alignment occurs via two separate features: push-off rods, and alignment pins. The three push-off rods are spring loaded features that seat into alignment cups on the passive end to provide gross alignment. The cone mounts are hard mounted features that also seat into cup features on the passive side. This geometry provides a statically determinate final position after mate.  Passive Damping(F radial/thrust bearings which transfer load into the housing. All three linkages are fixed to a piston, which is connected to a ball screw via a ball screw nut. The linkages consist of one upper grappling linkage and two lower linkages which ma The need for passive damping was identified during simulation and prototype testing. Concerns had been raised regarding difficulties that could arise during capture; specifically, the passive half impacting the push-off mount and rebounding. While it was agreed that the passive side would need to be held within a station keeping envelope during capture, a non-zero relative velocity would likely exist between the two spacecraft. This could result in slight amounts of contact between the two faces of the active and passive halves, making it difficult to maintain the passive structure in the required envelope. Also of great concern was unintentional impact involving even larger than nominal velocity deltas, and the possibility of damage to the mechanism.  Another motivation to introducing damping involved oscillations occurring during the capture event.  Testing of the prototype in a micro-g environment exhibited a tendency for the passive structure to oscillate within the constraints of the linkage tips and the push-off mounts. Computer simulations performed by Boeing confirmed this behavior. In order to address this issue, a spring damper system was added to the push-off struts on the active mechanism. Figure 11 depicts a cross section view of the push-off mounts with the spring damper system.  215
Adjustable Coulomb DamperCompression Piston SpringHigh Stiffness Piston HousingMounting FlangeDamper Rod SpringBushingDamper Rod Figure 11:  Push-off rod cross-section  Torque Sensing Mechanism  Running the motor/actuator until proper preload is achieved was accomplished using a torque sensing  mechanism (Figure 12). To apply a preload in the mated position, a specific torque from the motor is required. The torque sensing mechanism is adjusted to activate limit switches at the set torque needed for proper preload. Once the limit switches are activated, the motor is commanded to stop. In the stop mode, the motor/actuator applies a brake holding the preload developed.  Passive Assembly The passive assembly (Figure 13) is a three wedge shaped trefoil. The faces of each wedge terminate at  a center groove that runs along its length. The passive interface consists of an interface plate that is mounted to the three trefoils. Part of this geometry includes retention lips, or shelves, which provide the  linkage with a positive feature to grapple. The interface also includes kinematic cup features which help to align the structure and institute preload. The OE program required a sensor to confirm that capture had been successfully accomplished (Figure 9, step 2). Light Emitting diodes (LED’s) located on the passive assembly detect when the active assembly arms are in the captured position (Figure 14)  216
Figure 14:  Proximity sensor detail  OECS Flight Unit Testing  Acceptance Testing Based on the experience gained with the two prototypes, SRC developed and performed comprehensive  acceptance test program. In-house testing included many of the same elements that were present in the prototype and MDS testing: mate and de-mate verification, stiffness characterization, and off-load testing. In addition, thermal, thermal vacuum, and vibration testing were performed. Computer Simulation Dynamic models were updated to include all modifications incorporated after development. Significant  load cases were analyzed to verify that mechanism performance met the Orbital Express requirements. The same software and methodology were used as for the development unit design effort.  6 DOF Testing The final phase of qualification and flight testing included 6-DOF testing similar to that which was  performed on the MDS. Although the MDS and OECS mechanism features important for capture and alignment are almost identical, small changes were made to the design based on the results of 6 DOF testing of the MDS. Furthermore, the manufacturing fidelity of the OECS qualification components was flight-identical, unlike the MDS which was of engineering development unit quality. For these two main reasons, a second 6-DOF test of t med. The goal was to once again confirm that the dynamic modeling accurately depicted the dynamics encountered during capture and alignment.Sensor ElectronicsHousing Light beam Photo Transistor 3 plIR LED 3pl he qualification unit was perfor 218
Orbital Express Mission  AstroNEXTSat OECS Figure 15:  A uring flight  or the Orbital Express mission, the two spacecraft were launched attached together with an adapter  oupling ring. The OECS docking mechanism was engaged in a unique, preloaded condition for launch,  but was not the primary load path for NEXTSat. After system checkout and  tabilization, the robotic arm was used to attach the two spacecraft together. Prior to free flight  ic arm was used as a safety measure to keep the two spacecraft connected while  rmed. After proving out a  aptured” docking operation several distinct free flying docking operations were performed. In each case  spacecraft, the proximity flying, and the subsequent re-docking were performed  direct communication and electrical power transfer stro and NEXTSat mated by OECS d F c support of launch loads for  s operations, the robot the adapter coupling ring was released and jettisoned. The first docking operation was performed by  moving the spacecraft with the robotic arm (the two spacecraft were already de-mated and were held together by the arm) and operating the OECS docking mechanism. This operation was successfully performed and confirmed that further docking operations could be perfo“c the separation of the  autonomously by the two spacecraft without ground control intervention. While docked, key refurbishment operations were performed:  /g120 The electrical connections were utilized for  between the spacecraft. /g120 The fluid connector was engaged and transfer of propellant was successfully performed. /g120 The robotic arm was utilized to install an electronics box on a mating port on the NEXTSat  spacecraft.  Overall, all mission goals were successfully completed. The two spacecraft have since been  decommissioned. 219
Conclusions The OECS flight mechanism performed as expected and has supported demonstration of the  effectiveness of the soft docking approach enabled by the design. The blending of technology and vision  yielded a simple and effective concept. Through a series of design trades and system testing, the application of minimal mechanisms, and the use of common proven technology (i.e. motors, lead screws, and linkages), a highly functional design resulted. What began, during the primary phase, as a well developed concept quickly evolved into a commercially viable product. References 1.  Shane Stamm and Pejmun Motaghedi, “Orbital Express Capture System:  concept to reality,” SPIE  Conference, May 2004.  2.  DARPA websit ww.darpa.mil/orbitalexpress/ ehttp://w 220
- Occupied volume in the folded configuration should not exceed 140 x 200 x 80 mm3 (0.1 ft3) The main performance parameters are:  - Deployed stiffness should exceed 2 Hz  - Position accuracy at end of life should be better than 0.01° half cone  - Deployment repeatability should exceed 0.005° half cone - Latching shock on the payload should not exceed 100 N·m (75 ft·lbf)  The performance environment included the sine loading (15 to 20 g) and random vibration of a typical  launcher for these antennas (12 g RMS in-plane / 20 g RMS out of plane) as well as a qualification  temperature range between -90°C and +90°C (survival) with a deployment range of -40°C to +70°C.  Consideration of a storage period of five years in the folded configuration without loss of deployment  torque required accelerated creep testing.  Material and Geometry Trade-Offs  A key point of the development was defining a baseline conceptual design. The survey of state of the art  devices revealed advantages and disadvantages of existing technical solutions. A typical representative  of a compact mechanism is the elastic collapsible hinge (ECH) such as the Maeva [2] hinge depicted in  Figure 1.  Figure 1. Maeva hinge  At first sight the advantage of the hinge concept is its simplicity, but appearances are deceiving and the  overall behavior is complex and requires an extensive investigation of curved tape spring elements, which are cornerstones of the design. A short outline of the main characteristic of these elements follows.  Tape Spring Characteristics  We know tape springs from steel measuring tapes that utilize the snap-in effect to achieve a straight and  stable cylindrical shape. Curved tape springs provide a self-latching function. Fundamental studies of design and performance of these elements were carried out by Pellegrino and Steffen [2] [3] at the University of Cambridge. The schematic bending behavior is illustrated in Figure 2.  The tape springs are stiff until they buckle under a bending moment. In the buckling zone, the affected  cross section is flattened. The peak value is higher for bending against the tape spring curvature  (opposite sense) than for equal sense bending. The steady state moments are almost equal in both  directions The challenge in using an ECH design is building a high deformable hinge with sufficient deployment  torque and end stiffness. This was solved by using curved spring elements of different length. The longer  222
spring bent in an S-shape, in both equal and opposite senses, can be accommodated below the shorter  one resulting in a compact folded hinge with high torque capability.  MmaxMmax+M -/g60+/g60 Figure 2. Typical moment vs. angle graph for curved tape springs  Equal sense  bendin gShort spring  Long spring  Opposite sense  bendin g Figure 3. Baseline concept in deployed and folded state  Design The design of the ULMAAS incorporates four pairs of tape springs in a row, each with the curvature  facing inwards as shown in Figure 4. The rectangular arrangement maximizes torsion stiffness.  The major parameters for motorization and stiffness of the hinge are the free length of the spring, the  curvature and the width of the tape springs, and their thickness. The shorter spring (Spring B) can be  seen as a pivot for the mechanism. The distance between the springs controls the bending stiffness of the deployed hinge. The different length of the spring elements is compensated by four support bodies.  223
The support components provide large cut-outs with sufficient clearance for the flattened long tape  springs (Spring A) in their bent configuration.  Interface S/C side  Spring B  Interface washer  Spring A  End stop  Support  Interface reflector side  Figure 4. ULMAAS Design  Tape springs are connected to metallic fittings by adhesive bonding and an additional screw connection.  The bond minimizes stresses around the bolting hole while the bolt connection with curved washers reduces peeling loads on the adhesive. On the backside of the small spring, an end stop limits travel of the hinge during deployment after reaching the straight configuration (back buckling). It does not have effect on the alignment precision of the hinge.  A material trade-off for the springs included metallic and fiber reinforced materials, stiff materials with a  high elastic deformation ability. Achieving a high pointing stability over a wide range of temperatures  meant that the ratio of thermal conductivity to coefficient of thermal expansion was a main driver in the selection process. The decision to use carbon fiber reinforced plastic (CFRP) is based on the high thermal stability of this family of materials. After selecting CFRP as the material of choice, a direct comparison of several carbon fibers was made in  a search for a good balance between high stiffness and high elastic strain. The selected fiber is a Toray T800 carbon fiber with a 295 GPa (43000 ksi) tensile modulus and an elongation of 1.9%. The matrix system has been selected to meet stringent thermal stability and outgassing requirements. Bryte EX-1515, a cyanate ester system with considerable space heritage, was selected.  Spring Design  The performance of the hinge in terms of stiffness, torque and the general ability to bend is determined by  the design of the spring elements. The total thickness has to be in the range of 0.3 to 0.45 mm, corresponding to laminates with three to four plies. The availability of thinner prepreg material is deemed advantageous in the lay-up design of the spring elements as it allows greater tailoring of the composite.  224
A principle relation between the bending radius of the tape spring rand the curvature radius R of the  spring cross section was derived by Pellegrino [3]. The factor connecting these radii is derived from stiffness parameters of the specific lay-up. Equation (1) is used to assess the bending radii of the lay-up. R DDr /g152/g32** 2211 with [D*] = [D] - [B] [A]-1[B] (1)  [A], [B], [D] are the non-zero quadrants of the laminate compliance matrix. D11* and D22* are the diagonal  terms of the matrix [D*]. R is the curvature of the tape spring and r the resulting bending radius. The bending radius is important because it allows a design that fits in the desired envelope of the hinge.  With the approximate bending radius and the knowledge that the tape spring will nearly flatten in the cross direction, the peak strain can also be assessed.  The calculation considers only one single tape spring but the pair of springs will influence each other  during folding. The actual behavior is also driven by edge effects, material and manufacturing  imperfections that cannot be predicted properly.  Based on the performed analyses, the following lay-up combinations were selected for further  investigation: -  [0,+45,-45,90] and [0,-45,+45,90]  - [0,+60,-60] and [0,-60,+60] - [0,+45,-45] and [0,-45,+45] - [90,+45,-45,0] and [90,-45,+45,0]  Analysis  The activities in the field of simulation and calculation were intended as a support of the design and to  establish a verification approach to correlate and predict test results thus allowing scale-up of ULMAAS to meet specific end user requirements. Analysis was broadly based and included stress/strain, stiffness, buckling, deployment torque and thermal deflection predictions  Stiffness Natural frequencies of the coupled system hinge and antenna were determined in finite element (FEM)  calculations using I-DEAS software (see Figure 5a). The calculated frequencies showed good agreement with those derived from test data later on. The lowest natural frequency was 3.2 Hz and represents the torsion stiffness. The frequency in the folding direction (bending mode) is nearly the same (3.4 Hz).   225Figure 5a. I-DEAS Simulation modelFigure 5b. ANSYS Simulation modelFigure 5c. Simulation model PamCrash