ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
108 Bearing Active Preload System (BAPS)   Many of the lessons learned on the BAPS, especia lly the ones concerning manufacturing and assembly  are presented in Ref. 2 and will not be repeated here. However, there are some major lessons from the  usage of such a BAPS that will be discussed below.   For the CPA application with the very high requireme nts on torque noise, life, jitter etc. the BAPS is an  ideal solution. The measurements done so far c an only confirm and justify its implementation.   One major challenge turned out to be the used SMA ac tuator. This actuator type was chosen due to very  tight volumetric/geometric constraints. Several optimization loops were performed and a specific  screening process has been introduced in order to obtain a qualified solution.   Conclusions   The CPA currently under testing at RUAG Space has been presented in this paper. It has been shown,  that the very tight requirements ca n be met by a highly optimized design providing multi-purpose usability.  The main performance characteristics have been reported among which the low torque noise and the sustainability of high launch loads shall be pointed out.   For future “coarse” pointing mechanisms  it is recommended to use a BAPS (Ref. 2). 

References 

1. Mussett D., et al. (2003) “Contraves Optical Term inal – Coarse Pointing Assembly (CPA)”. In Proc.  10th European Space Mechanisms and Tribology Symposium (San Sebastian). (Ed. R.A. Harris) ESA SP524, ESA Publications Division, European Space Agency, Noordw ijk, The Netherlands  2. S.D. Lewis et al. (2009) “Development of an Adju stable Bearing Preload Enabled – Optical Terminal”.   In Proc. 13th European Space Mechanisms and Tr ibology Symposium (Vienna). (Ed. H. Lacoste)  ESA SP670, ESA Publications Division, European  Space Agency, Noordw ijk, The Netherlands 

Acknowledgements 

Special thanks are due to ESR Technologies (EST L) for their development of the BAPS, the numerous  hours spent for their work and their constant support.  Also thanks are due to the team of TESAT Spacecom for their continued interest and involvement during the development of this product. Furthermore, the German Aerospace Center (DLR) and the Swiss Space  Office (SSO) vigorously drive the development of optical communication forward with their support.    Last but not least, it shall be acknowledged with a hearty “thank you”, that the author team listed in this  paper would not have been able to accomplish this exceptional feat without the constant support of the rest of CPA team as well as the management at RUAG Space with their high motivation and  perseverance.

NASA/CP-2010-216272
110 of the clearance volume has been thoroughly analyzed to ensure sufficient clearance between the Gimbal  Ring and the coiled propellant lines over the full co mbined range of angular travel of both gimbal axes. 

Figure 1. Modular Construction     Rotary Actuator    Two rotary actuators are provided for the articulati on of the Thruster Gimbal (Figure 2). The rotary  actuators consist of four-phase unipolar stepper moto rs coupled with Harmonic Drive gear transmissions.  These actuators are also configured with large output flanges mounted on preloaded duplex bearing set. This bearing mounting provides a high degree of mo ment stiffness while causing minimum drag torque  reflected to the motor, at the input of the Harmonic Drive gear reducer. A potentiometer is provided for  position telemetry. It is located in the rear section of the actuator, with the potentiometer wipers directly  connected to a shaft that extends through the length of the actuator from the output flange. Since the output of the actuator is attached to the inboard/outboard platform of the thruster gimbal, the position indication of the potentiometer located inside the actuat or correlates directly with the gimbal axis angle. 

Inboard and Outboard Platform    The Inboard and Outboard Platforms are almost identical components, with the exception of the mounting  hole patterns designed to accommodate spacecraft and thruster mounting interfaces (Figure 3). Other  differences between the Inboard and Outboard Platforms are the use of different materials for the mounting spacers, which will be discu ssed in the “Thermal Control Features” section. Materials for these  brackets were selected carefully, due to the requirement for low weight and the critical dimensional requirements for orthogonality and parallelism of the gimbal axes  For simplicity and light weight, the brackets were designed as one-piece machined parts. The basic Ushape and the extensive lightweighting employed in t he design mean that a large amount of the original  billet material is removed in machining to the net shape. The need for stability of certain critical dimensions in the finished parts- required for gimbal performance- resulted in special design  NASA/CP-2010-216272
111 considerations. In addition to light  weight, a material exhibiting good dimensional stability over time was  needed. After researching the avail able materials, a 7000-series aluminum was selected as having the  best combination of properties. Th is material demonstrated high di mensional stability in the finished  brackets.  

Figure 2. Elevation and Azimuth Rotary Actuators with Heaters 

A number of other design features were incorporated to produce a high stiffness-to-weight ratio. These features were carefully implemented in order to avoid weak points in the structure and to avoid  interferences over the large arti culation range of the gimbal.  

Gimbal Ring   The Gimbal Ring is the structural member connecting  the Inboard and Outboard Platforms of the Thruster  Gimbal (Figure 3). It is in the load path and is a critical structural component.  

Figure 3. Inboard and Outboard Platform with the Gimbal Ring   The gimbal ring is designed as a large round annular structure, creati ng a circular clearance volume in  the center of the gimbal. Although c hanging in shape, that volume remains clear over the full articulation  NASA/CP-2010-216272
113 in all operational configurations of the gimbal, not  interfering with coil movement. This solution was  implemented simply with tie cords spaced at optimum locations on the tubing assembly (Figure 7). These  two design measures produced a tubing assembly with  the flexibility to allow full gimbal deflection and the  support needed to withstand launch vibration.  Another requirement on the propellant line assembly was maintenance of tubing  temperatures conducive  to proper performance of the xenon propellant, and therefore heaters were required. Adhesive-backed film heating elements are normally used for this func tion; however, the 1/8-inch (3.2-mm) diameter  propellant tubing did not offer sufficient flat areas for application of the heaters. Further, the tubing coils  flex in operation, compromising adhesion of such heaters. A different solution was needed.  The heating elements to be used requir e flexibility similar to the flexibility of the tubing coils themselves.  Therefore, resistive heater wires fo rmed into coils matching the tubing coils were fabricated. The heater  wires were sandwiched between two layers of insulating Kapton film (Figure 5).  

Figure 4. Outer Propellant Tubing with Heater  Figure 5. Inner Propellant Tubing with Heater  

      Figure 6. Propellant Line Assembly   with Spacers  Figure 7. Fully Articulated Assembly  with Restraining Cords   Figure 6a. Propellant Line Assembly   NASA/CP-2010-216272
114 Thermal Control Features 

The operating environment of the Thruster Gimbal  presents a number of challenging thermal design  problems. As discussed above, the propellant supply lines are furnished with heaters. Heaters are also  provided for the gimbal actuators. For temperature monitoring and control, ther mistors are located on the  propellant lines near the Inboard and Outboard Platforms as well as on the actuators.  The location of the thruster gimbals on the spacec raft is such that they are exposed to extreme  temperatures and temperature gradients. In the worst ca se, the gradient from Inboard to Outboard can be  as high as 100 deg C. This extreme thermal condition was mitigated by tailoring thermal conductivity at the gimbal interfaces. At the Inboard Platform inte rface with the spacecraft, a set of four aluminum  spacers were used to promote conductivity. At the Ou ter Platform interface with  the thruster assembly,  four spacers were used to limit heat transfer from or to the thruster assembly an d to promote isothermality  of the gimbal structure.   The spacecraft operating scenario results in the Thru ster Gimbal being exposed to a deep space view,  and the resulting extreme cold, for extended periods of  time. A flexible multilayer insulation blanket  encapsulating the entire gimbal was used to protect against the resulting radiant heat loss, and, again to  promote isothermality of the gimbal . Innovative design of the Multilayer Insulation Blanket (MLI) was  required to accommodate the large travel angles of the gimbal. A camera bellows type of construction was used, and was found to offer minimum additional torque load on the gimbal actuators (Figure 8). 

Figure 8. Multilayer Insulation (MLI) Blanket  

   NASA/CP-2010-216272
115 General Performance of the Thruster Gimbal 

Thruster Gimbal Specifications 

Physical Characteristics  Dimensions 22.23 x 12.7 x 12.7 cm (8.75 x 5 x 5 in)  Weight < 5 kg (11 lb)  Payload Weight 22.7 kg (50 lb) (externally supported)     Performance  Total rational range of travel ±36.5 degrees in two axis  Angular resolution of gimbal 0.01125 degree  Angular velocity 1 deg/sec (nominal)  Angular accuracy of gimbal Under 0.02 degree  Operating temperature range -20° C to +80 ° C     Power Requirements  Power Consumption 22 watts max. per actuator 

Lessons Learned 

The motor in the elevation and azimuth actuators is  a four-phase 1.8-degree hybrid stepper, wound for  unipolar excitation. The motor design has high magnet ic efficiency, allowing it to produce high output  torque and unpowered holding torque at relatively low power levels. The high magnetic efficiency results  in part from a small motor air gap. The air gap is the radial clearance between teeth on the moving rotor  and opposing teeth on the stationary st ator core. Although it is important to maintain a small air gap in  order to achieve optimal motor perf ormance, it is equally critical to  ensure that the clearance between  parts in the air gap is maintained. It was recogni zed early in the development phase that diligent  measures would be required to insu re consistent air gap dimensions in production units. All elements  affecting air gap variations were thoroughly analyzed and a Critical Clearance Analysis report, based on worst case values, was generated.  Approximately 26 actuators were produced (some shown on the completed gimbals in Figure 9). In thermal cycle testing, one axis of a  thruster gimbal stopped operating during a high temperature portion of  the cycle. As a troubleshooting step, the chamber tem perature was lowered. The affected actuator then  resumed normal operation. The preliminary conclusion  was that the anomaly was temperature-related. 

Figure 9.  Completed Gimbals  NASA/CP-2010-216272
116 The gimbal was disassembled and the problem traced to one of the act uators. Actuator disassembly then  further isolated the cause of the problem to the motor. Motor disassembly showed rub marks on the rotor  with matching marks on the stator. These marks were  consistent with a condition causing the rotor-tostator clearance to be lost, with resultant physical contact.   The Critical Clearance Analysis was revisited, and no errors in the methodology or calculations were found. However, it was noted that the analysis assum ed a uniform coating of insulating material thickness  on the critical rotor and stator surfaces. This assumption was investigated further.   The magnetic materials susceptible to corrosion are protected by a thin film of epoxy material- a so-called ‘wash coat’- which forms a barrier coati ng over the otherwise vulnerable surfaces.   Wash coat application is a standard and well esta blished process used on virtually all Moog motors.  However, the majority of motors produced are of the three-phase permanent magnet stepper design, with  relatively larger air gaps not requiring ex tremely tight control on coating thickness.   The conclusion of the investigation wa s that the actual thickness of t he wash coat in places exceeded  dimensions assumed in the Critical Clearance Analysis and, together with differing thermal expansion coefficients and/or thermal gradients in the motor assembly at high temperatures, caused the closure of  the air gap and resulting stoppage of the motor. Figure 10 and Figure 11 demonstrate the radial location where the air gap closure occurred.       

Figure 9. Rotor      Figure 10. Stator  

Conclusion 

Design and development of the Two-axis Thruster Gimbal was completed largely free of interruptions and course corrections. However, a significant number of up-front analyses were precip itated by the stringent  operating, performance, and weight requirements. In  addition, a Monte Carlo analysis was performed to  predict performance in simulations using randomly selected thermal and operating parameters. To further  ensure successful production of flight hardware, several Engineering units of varying degrees of  complexity were built and tested. By the conclusion  of the Final Design Review, an extraordinary amount  of analytical and test data had been accumulated,  indicating a well conceived and low risk design  approach.      NASA/CP-2010-216272
117 A large part of the development effort was focused on the mitigation and suppression of, and  compensation for, the extreme thermal conditions to  which the gimbal would be exposed. Heaters which  could be mounted on moving and flexing members were developed, as well as complex, flexing thermal blankets designed to contain thermal energy and pr omote isothermality. Aggressive weight reduction  measures were also implemented in order to achieve an optimized design with adequate margin.  Unique testing methods were developed, enabling testing of the Thruster Gimbal with simulated payload and launch lock mechanisms in order to emulate f light-like test conditions with high fidelity.  Despite the up-front development effort and the exte nsive analyses performed to lower and/or eliminate  risk, an oversight in process selection allowed a test anomaly to occur. A motor fabrication process commonly used for Moog motors, but inappropriate for the optimized Thruster Gimbal motors, was used. The process was modified to specifically address t he unique characteristics of the Thruster Gimbal.  Numerous other process improvement s were implemented as well as the specific testing measures used  to isolate and correct the observed anomaly.  Moog has successfully produced a number of the Thruster Gimbals, and has recently concluded a second life test program, demonstra ting 2X life. This follows the original life test program which  successfully demonstrated 1.5X life.   Acknowledgements    The author wishes to acknowledge James Hammond , Steve Mallonee and Ruben Nalbandian of Moog  for their contributions in the design and deve lopment of the Two-axis Thruster Gimbal.   The author expresses his sincere appreciation to Da ve Putnam of Lockheed Martin Space Systems for  his significant contribution in the development of the Thruster Gimbal and his hands-on involvement during the complex Qualification Testing phase of the program.      NASA/CP-2010-216272
120 Other gravity offloading approaches use support fr om above cable arrangements that place a simple  linear spring in series with each cable. Unfortunatel y the very low stiffness required of the spring poses  the problem of static sag. This ma kes suspension cable systems only practical for vertical displacements  on the scale of centimeters, whic h is inadequate for deployment testing. The WAGM utilizes established  AGM technology to deliver vertical gravity offloading with  zero static sag and air spring stiffness limited  only by the volume of the external volume connected to the AGM. 

Figure 1: Mechanical gravity offloader   In its first application, WAGM replaced the mechanical pulley and counterweight gravity offloader shown in Figure 1. The pulley and counterweight were attached to a moving trolley which rode on a guide rail. The guide rail was attached by a hinge to the si de of the spacecraft. T he customer expressed  dissatisfaction with the fact that the counterweight  overhung the payload, the system took a long time to  set up and required bolting test e quipment to the side of the spacec raft, and most deleteriously, the need  for the deploying payload to pull the “jib crane” around and to move the counterweight inertia compromised torque margin measurements. Torque marg in of the deployment system is an important  criterion to the customer as it is fundamental to on- orbit reliability. In tests on the WAGM, friction in the  vertical direction at the payload end of the load cable was measured to be on the order of 0.25 – 0.35% of  payload weight. 1 In the horizontal degrees of freedom, drag forces were measured to be just 0.5% of  payload weight at payload speeds of 1.59 mm/s (0.0625 in/s). Moreover, interfacing to expensive flight  hardware is kept to a minimum as the deploying payload under test is single point suspended from a light load cable with no other interface between the WA GM and any part of the spacecraft. The latter  advantage greatly reduces setup and test time, an im portant factor since tests are performed with the  spacecraft very close to completion and the tests on the critical schedule path.                                                        1 The AGM itself is completely frictionless. The me asured friction is added by the 4:1 displacement  amplifying mechanism described later.   NASA/CP-2010-216272
121 Requirements and System Architecture 

The WAGM was commissioned to offload the weight of a deployable robotic arm during ground testing.  The requirements were:  • Minimal friction  • Minimal setup time  • Minimal interfacing to spacecraft  •   Offloading force of 355.86 N (80 lb) with an offload accuracy of + 89 mN (+ 0.02 lb) in x, y, z  • Insensitivity to set-up or set-up repeatability 

Figure 2: WAGM during acceptance testing at customer site 

Figure 2 shows the WAGM during acceptance testing at the customer site. The WAGM is composed of  five main subsystems.    • The gravity offloading device (AGM) that carries the weight and accommodates vertical motion of the payload.  • An X-Y stage system comprised of sealed linear bel t drives that move the AGM horizontally to  keep the load cable vertical during active tracking. The X-Y stages are driven by a pair of commercial off-the-shelf stepper motors.  • A COTS industrial lift whose fork has been replaced by a custom load platform onto which the X-Y stages are mounted. The industrial lift has a telescoping tower that allows for coarse adjustment of the height of the load platform. The base of the lift is equipped with casters that  allow the entire assembly to be rolled into place.  AGM  X-Stage  Y-Sta ge  Telesco ping Towe r  Frame Supporting  Deployable Payload  Load Platform  Load Cable to Payload Pick Point  Control Rack Stepper Moto r NASA/CP-2010-216272
123                            

Figure 4: AGM detail  

During active tracking, horizontal forces on the payload  are minimized by an acti ve system that keeps the  AGM centered over the payload such that the load cabl e is kept very close to vertical as the payload  moves. The key to doing this is a system for sensing the load cable angle relative to true vertical. This is done by routing the load cable up and over two additional pulleys and down through a guide tube as shown in Figure 6. The guide tube is part of a larger angle-sensing mechanism: the upper end of the  guide tube is fixed to the lower yoke of a precision universal joint (the upper yoke of the universal joint is fixed to the baseplate of the AGM) into wh ich are mounted two high-sensitivity, DC-coupled  accelerometers (i.e. tip-tilt sensors) with their se nsing axes oriented normal to the guide tube.  As the  payload moves horizontally the load cable departs fr om vertical, pulling the lower end of the guide tube  with it. The DC-coupled accelerometers produce output signals proportional to the tube’s deviation angle from vertical (for small angles). Using band-limiti ng lowpass filters in the signal conditioning, the  accelerometers have a resolution noise floor of about 1 micro-g, corresponding to 1 micro-radian of angle for small angles of the guide tube around vertical. The signals from the accelerometers are gained up to a sensitivity of 200 volts/radian. The ±10 volt ran ge at the amplifier output then corresponds to a  measurement range of ±50 milliradia ns. The basic advantage  of the sensing method is that it is  unaffected by small deviations of the load platform fr om horizontal, as will inevitably be caused by flexing  of the platform and the tower under load.   Moving  Carriage  Air  Cylinder Buffer Air  Tank Carriage  Position  Sensor  4:1  Displacement  Amplifier  Load  CableTip-Tilt  Sensors  Emergency  Cable BrakeBaseplate Guide  Tube  NASA/CP-2010-216272
124  

Figure 5: Cable-and-pulley displacement multipliers.  The WAGM uses the 4:1 system. 

The deviation angles are filtered through an analog anti-aliasing filter module before being digitized by a packaged motor controller card. Inside the motor controller card a feedback loop outputs step and direction commands to the stepper motors to position the X-Y stages such that the tip-tilt sensor is  maintained directly over the payload, keeping the lo ad cable as vertical as possible and minimizing the  deviation angles. A 1-Hz low-pass digital filtering on board the controller card is included in the feedback  loop to keep the system stable and to prevent the control loop from amplifying ambient ground vibration  and noise in the accelerometer signals. It is important  to note that the controller only has control authority  over the movement of the X-Y stages. It has no cont rol authority over the AGM piston pressure or the  AGM piston displacement. This is because the AGM is  a constant force device; once the AGM cylinder  pressure is set the uplift force on the payload remain s essentially constant over the vertical range of  motion.      1:1 displacement ratio2:1 displacement ratio to load to load to load~4:1 displacement ratioAGM NASA/CP-2010-216272
125  

Figure 6: Q-flex DC coupled accelerometers, cable guide, and payload pick point 

Figure 7 shows the control system block diagram for one translational axis. A difference in the horizontal  position of the pick point relative to the AGM creates an off-vertical angle in the load cable that is sensed  by the tip-tilt sensor. The analog tip-tilt sensor out put is digitized and digitally low-pass filtered with a  cutoff frequency of 1 Hz. Proportional gain is then applied to generate a velocity command (stepper counts/sec) that is digitally filtered and then inte grated. Stepper counts are turned into engineering units  and a gain is applied that accounts for the cable length variation as the pick point translates vertically. The result is the setpoint for t he native position control loop running on the card. Both control loops are  executed at a 2-kHz update rate. 

Figure 7: Control system block diagram  Load CableCable Guide Universal Joint  Guide TubeQ-Flex DC-Coupled  Accelerometers  NASA/CP-2010-216272