ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
1 w , deg/s pedestal threshold sw412:34 run time, sec .5 radius..5radiusp1mc/mo1/12 in to ft -K-gcenergy*x/t   .5 a1 time OrionSp Wrksp211 W pedestal  height, inIn1 In2 In3 In4 In5 In6 In7Out1 Out2 Out3 Out4 Out5 Out6 Vertical VelocityIn1 In2 In3 In4 In5 In6 In7Out1 Out2 Out3 Velocity  Calculation 0 Thrus t, RCS, lbf 1 TVC angle, degIn1 In2 In3 In4 In5Out1 Out2 Out3 Subsystem1In1 In2 In3 In4 In5 In6 In7 In8Out1 Out2 Subsystem1 Spring limit, in Spr Sep Clr 111Residual Engine Thrust, lbfRel ative Clearance, inR2D Rads to Deg In1 In2 In3 In4In5Out1 Out2RET System  PertubationIn1 In2 In3Out1RCS Effects In1 In2 In3 In4 In5 In6 In7 In8 In9 In10Out1 Out2 Out3 Out4 Output Condition 111Orion length, in Cg to OMS engine bell0 Orion Rotation deg/s  - 11 Orion OMS Diameter, in1e8Orion Moment of Inertia about "z",  lbm in211111 Orion Mass, lb 1111 Orion Cg, in from CLV engine baseIn1 In2 In3 In4 In5 In6Out1 Out2 Moment  Subsystem100 K s pring, lbf/in1 In1 In2 In3 In4 In5 In6 In7 In8 In9 In10 In11 In12Out1 Out2 Out3 Out4 Out5 Geometry Separation Velocity SubsystemDivD2R Degr to Rads CLV w/ped length total, in CLV to Orion Cg  inCLV system  inertia 111 CLV length, in00 CLV Rotation deg/s  - 111CLV Pedestal ID, in1e9 CLV Moment of Inertia about "z",  lbm in211111CLV Mass, l b CLV Length total, in1111 CLV Cg, in from engine base In1 In2 In3 In4 In5 In6 In7 In8 In9 In10Out1 Out2 Out3Angular VelocityIn1 In2 In3 In4 In5 In6 In7Out1 Out2 Out3 Out4Actuator Force  Subsystem u2 u21time Figure 4. Simulink Separation Model  The basis of the Simulink analysis utilizes the conservation of momentum and kinetic energy equations  shown below:  /g11/g12 o2oo o c2cc c stack r mI rmI  I /g90/g184/g185/g183/g168/g169/g167/g14/g14/g90/g184/g185/g183/g168/g169/g167/g14/g32/g90 and /g184/g184 /g185/g183 /g168/g168 /g169/g167/g90/g14/g14/g184/g184 /g185/g183 /g168/g168 /g169/g167/g90/g14/g32/g184/g184 /g185/g183 /g168/g168 /g169/g167/g90 2 I v m 2I v m 2 I2oo2oo2cc2cc2stack Where the variables are defined as:  Istack = Stack (Orion + Ares 1 US) moment of inertia, kg-m2 Ic = Ares 1 US moment of inertia, kg-m2 Io = Orion moment of inertia, kg-m2 /g90 = body stack rate of rotation, deg/s  /g90c= Ares 1 US rate of rotation, deg/s  /g90O= Orion rate of rotation, deg/s  mc= Ares 1 US mass, kg  349
mO= Orion mass, kg  rc = Ares 1 US c g moment arm to system cg, m  r0 = Orion c g moment arm to system cg, m vc = Ares 1 US relative velocity to system, m/s vO = Orion relative velocity to system, m/s The results of solving these equations are that t he rotation rate of the separated components is  maintained at the same rate as the sta ck rotation prior to separation. The addition of a residual engine  thrust post separation does induce an additional mom ent onto the Upper Stage and results in an angular  acceleration, reducing the clearan ce during separation.   ADAMS Approach The SA/Upper Stage and Orion vehicles were also modeled using the ADAMS dynamic software code  [13]. This is a motion simulation code that allows the user to create a mechanism model and then solves  the simultaneous equations for kinematic, static, quasi-static, and dynamic simulations. Figure 5 depicts the ADAMS separation model during a simulation as Orion clears the SA. For the purposes of this study  the Upper Stage and Orio n were modeled as rigid bodies, each with six degrees of freedom. The SA was  modeled as rigidly linked to the Up per Stage since it never separates fr om it. However the geometry of  the SA, particularly the top flange was important for this analysis since the Orion engine nozzle needs to  be extracted from this cavity and translate beyond the top flange of the SA without impact. The Crew  Module and SM were modeled as a single rigid body (i.e. Orion) since, again, they never separate in this analysis. The modeling of the separation systems for the Orion to Ares 1 US included the single axis  springs or actuators located around the SA top flange, between the SM  and SA. These compression  springs produce a translational motio n when released. Once their free len gth is achieved they no longer  impart any force onto the vehicle. The Ares 1 US’s J-2X engine thrust is modeled at the bottom of the  Upper Stage. Auxiliary engine thrust is also accounted for on several design studies and these are also  modeled as point forces located at the current auxiliary engine locations near the separation plane. Vehicle “dump” or pitch rate is applied as an initial velocity condition to the Ares 1 US/Orion at the  combined vehicle stack center of gravity (CG). Joints were added to the model as follows: a fixed joint  was created between the Upper Stage and the CEV to allow them to pitch together at the start of the simulation (time = 0 sec.) and release when the separation event began arbitrarily at time = 1.0 sec.; a  hinge joint was created at each solar array anchor point under the avionics ring to allow for flaring of the  arrays in order to investigate different launch configurations to optimize array clearance from the outer  fairings at launch as well as avoid impact with the SA upon separation.   Figure 5. ADAMS Separation Model 350
Results The clearance requirement for this system was to provide a 5.1-cm (2-in) minimum clearance between  the Orion vehicle (engine nozzle, solar arrays, and other protrusions) and the SA. Several analyses were completed to size the actuator forces, determine  practical separation times and optimize the vehicle  geometry (engine size, solar array placement, SA diameter, etc.). Figure 6 plots the time that the  minimum clearance is reached versus the vehicle dump rate for a gas actuator system and for the case of no actuator forces, using the Simulink model. From this plot the two systems are seen to coincide at the  higher body rates where the actuation force needed for the 5.1-cm (2-in) clearance is diminishing as the  system approaches the no-force required condition.  0123456 0 1 02 03 04 05 06 Vehicle Dump Rate, deg/sTime to Minimum Clearance,  sec  00.051m Clear Constant Force No actuator force Figure 6. Clearance Time as a Function of Vehicle Body Rate  The initial actuation force required to meet the clearance requirement at the baseline 5 deg/s dump rate is  shown in Figure 7. The curves in the figure are the Orion radial clearance (red), the Orion axial separation distance (green), and the actuator force line (blue) which shows when the actuator force is terminated  (0.33 sec) and is reflected in the slope of the Orion separation distance curve which tends to be more  linear after this force is removed. The jog in the clearance curve occurs when the Orion vehicle clears the SA at approximately 4.8 seconds, which is when the change in clearance becomes a positively sloped  line as the vehicles move further apart from each other with no chance of contact.  351
00.20.40.60.811.21.41.61.82 00 . 511 . 522 . 533 . 544 . 55 time, secDistance/Clearance, mActuator  Force ClearanceAxial  Separation Figure 7. Time Required for Separation at 5 deg/s, (Simulink Model)  Figure 8 is a plot of the Simulink model separation force required to provide clearance at an initial dump  rate of 5 deg/s for both types of actuators, springs and gas (or pneumatic) thrusters. In the spring type  actuator the force is a function of axial displacement where the initial force is very high and then decreases along a power curve as a function of time. In the gas thruster system, which can be modeled  as constant pressure through its full stroke, the force is maintained throughout the action time at a  constant level. Both systems were sized to  separate the vehicle with exactly 5.1-cm (2-in) clearance  maintained at the engine nozzle. Duration of the const ant pressure actuator is determined by the 10.2-cm  (4-in) actuator stroke. The calculated time necessary to provide this clearance is shown to be  independent of the actuation method, since the vehicles are rotating and translating at a rate as a function of dump rate and the residual engine thrust on the Upper Stage. The 5.1-cm (2-in) minimum clearance  point in space or “gate” is reached at the same point in time, which varies from 24.0 deg (4.8 s) to   23.4 deg (0.39 s) of vehicle rotation for the 5 deg/s to 60 deg/s body rate respectively.   There is a very slight difference in the final velocity induced by the actuators since to make the clearance  gate time, the spring actuator provides a higher initial acceleration and then coasts at the resulting  velocity of 0.477 m/s (1.56 ft/s) with an acceleration rate of 0.144 g’s, while the constant pressure  actuator provides an acceleration rate of approximately 0.118 g’s over a longer time period resulting in a higher 0.485 m/s (1.59 ft/s) final velocity imparted on the Orion system. Therefore, less energy is required  for the spring actuator compared to the gas thruster mechanism. 352
05001000150020002500 0 0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4 0.45 0.5time, secForce, kgSpring  Force Gas Thruster1095 J  (808 ft-lbf) 1144 J (844 ft-lbf)Total impulse for Spring is 1095 J vs Gas thruster @ 1144 J Average accel for Spring is 0.026 g's greater Actuator induced velocity for Spring is 0.477m/s vs 0.485 m/s Both achieve clearance at 4.83 sec Figure 8. Separation Force Input for Spring vs. Gas Thruster  For the ADAMS model several cases and design studies have been run to determine the optimal  separation system. The two critical considerations in these dynamics analyses are to  ensure that the SM  engine bell can get extracted from within the SA without bumping (avoiding a Falcon 1 type of hazard),  and that the solar arrays, which are mounted down the sides can clear the SA without interfering from the  outside. Using the given mass properties the resulting spring stiffness case results are shown in Table 1  as well as Figure 9. All cases are assuming there is no separation assistance from the Orion SM main engine.  Table 1. Summary of Parameters Analyzed  case #dump rate  (deg/s)J-2X  Residual  ThrustCEV RCS  Thrust     kg  (lbf)Spacecraft  Adapter flange ID  m (in)Spring stiffness  kg/m (lb/in)Spring stroke  length            m (in)min clearance,  ADAMS model   m (in)min clearance,  Simulink model   m (in)model         delta          m (in) 1 35 yes 0 3.4 (135) 0 0.102 (4) 0.14 (5.6) 0.11 (4.5) 0.03 (1.1) 2 0 yes 0 3.4 (135) 10,724 (600) 0.102 (4) 0.80 (31.5) 0.72 (28.3) 0.08 (3.3) 3 10 yes 0 3.4 (135) 178 (10) 0.102 (4) 0.07 (2.6) 0.08 (3.0) -0.01 (-0.4) 4 10 no 0 3.4 (135) 178 (10) 0.102 (4) 0.15 (6.0) 0.13 (5.2) 0.02 (0.8) 5 20 no 0 3.4 (135) 178 (10) 0.102 (4) 0.15 (5.9) 0.13 (5.1) 0.02 (0.8) 6 20 yes 0 3.4 (135) 178 (10) 0.102 (4) 0.13 (5.2) 0.10 (4.1) 0.03 (1.1) 7 5 yes 0 3.15 (124) 4,147 (232) 0.102 (4) -0.08 (-3.3) -0.05 (-1.9) -0.04 (-1.4) 8 35 no 0 3.15 (124) 0 0.102 (4) 0.01 (0.2) 0.01 (0.2) 0  9 5 yes 366 (808) 3.4 (135) 0 0.102 (4) 0.18 (7.2) 0.11 (4.5) 0.07 (2.7) 10 5 yes 0 3.4 (135) 4,147 (232) 0.102 (4) 0.05 (2.0) 0.08 (3.2) -0.03 (-1.2) 353
00.050.10.150.20.250.30.350.40.45 2500 4500 6500 8500 10500 12500 spring stiffness, kg/mclearance,   m3.08m (121.6") flange dia.- array 3.82m (150.4") flange dia.- nozzle 3.08m (121.6") flange dia.- nozzle3.82m (150.4") flange dia.- array3.34m (131.2") flange dia.- array 3.34m (131.2") flange dia.- nozzle3.58m (140.8") flange dia.- nozzle 3.58m (140.8") flange dia.- arrayallowable Figure 9. Engine bell and 6-m Array clearance for 5o/sec dump rate  Figure 9 depicts the results of several ADAMS cases using different SA flange diameters to assess  clearing the solar arrays on the outside versus clearing the engine nozzle on the inside. Previous Orion  designs incorporated a longer but narrower vehicle with a likewise narrower SA. The redesign of the  vehicle allowed for a wider SA flange. As can be seen in the plot, the arrays have adequate clearance  (>5.1 cm) for any SA inner diameter of 3.58 m (140.8 in) or less, while the engine nozzle will have adequate clearance for a SA diameter of 3.34 m (131.2 in) or greater. Thus an inside diameter of 3.34 –  3.58 m (131.2 – 140.8 in) satisfies both. In these cases it is assumed that the J-2X engine residual thrust  is active and the flange width is 0.343 m (13.5 in) radially. A nominal separation system would include springs located at SA nodes as shown in Figure 2 outboard of the separation pyrotechnic device with a  10.2-cm (4 in) stroke and 3,842-kg/m (215-lb/in) stiffness. Six standard 1.27-cm (1/2 in) separation bolts  located directly inboard of the push o ff springs at each node will transmit launch loads through the  structure.  Figure 10 is a plot of the separation clearance as a function of time for the different models used for the  baselined configuration at a body rate  of 5 deg/s. The ADAMS model includes the Solar Arrays for  additional clearance studies while the Simulink model only considers the clearance for the Engine Nozzle to the SA, which becomes the limiting parameter for both models after approximately 2.4 seconds.  354
00.20.40.60.81 0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5 time, secSeparation Clearance, mADAMS Global - Nozzle Clr ADAMS Global - Solar Array Clr Simulink Global - Nozzle Clr Figure 10. Separation Clearance at 5 deg/s Body Rate  Figure 11 depicts the clearance achieved as a function of actuator force for the Simulink and ADAMS  models at a 5 deg/s dump rate. The interference at the low actuator force is due to the influence of the  applied Residual Engine Thrust moment and to the longer separation time required for the low dump  rates (<10 deg/s). This slower separation time allows the induced moment on the Upper Stage to rotate the SA reducing the clearance below the required limit. From the plot it is evident that an actuator force of  greater than 3500 kg/m (196 lb f/in) is required to assure the clearance is achieved and that use of a  grossly oversized actuator has diminishing returns since the actual clearance is not a linear function of spring stiffness.  00.10.20.30.40.5 3500 5000 6500 8000 9500 11000 12500 14000 actuator spring stiffness, kg/mclearance, mAdams Simulink 0.051m Clearance Re quired Figure 11. Comparison of Clearance vs. Actuator Force  355
Natural Separation It has been learned th rough this analysis that for two bodies rigidly fixed together and undergoing rigid  body rotational and translational motion which then separate, if body 1 has a protruding feature (like an  engine bell) tucked inside a recessed area of body 2 (such as the SA cavity) that, due to the centrifugal forces naturally propelling them apa rt, there exists a relationship between the diameter and length of the  protruding feature and the mating clearing radius Y B1 of the recess whereby for a recess radius greater  than Y B1, the vehicles will separate without collision at any dump rate with no additional kickoff force  required nominally. Figure 12 depicts the geometry definitions used.  Body 2  w/recessBody 1  w/protrusion Figure 12. Ares I US/Orion Vehicle Geometry and Definition  RB*COS(/g84Bt) = X A1+XAtrans+XArot ; where X A1is the initial axial distance from Body 2 CG to  Protrusion point A, X Atransand X Arotdepict the axial & rotation motion components of point A,  which represents the outermost point of the protrusion. RB = R A *(SIN(/g84At)/SIN(/g84Bt))  SIN(/g84A1) = YA 1/RA RB = SQRT[(X B1)^2 + (Y B1)^2] XCG1A+XA1= CG off;   CG offis the distance between Body 1 and 2 CG’s  prior to separation . XAtrans = /g90*Tcoll* CG off  ;Tcoll is the time needed for pt. A to separate and pass thru pt. B at X.  XArot = R a *[COS(/g84A1) - COS(/g84At)] XBt = R B *COS(/g84Bt) XBt+XAtA = CG off +X Atrans /g84Bt = /g84B1-/g90*Tcoll /g84At = /g84A1+/g90*Tcoll Thus, the minimal recess radius for natural separation, Y B1, can be solved for easily, if X A1, YA1 (protrusion  radius), and lengths X B1 (recess radius) and CG off(distance between CG’s) are known. This analysis  356
assumes the protrusion and recess are modeled as straight cylinder sections. This finding gives a  designer a useful preliminary size for the vehicle recess diameter (such as in a spacecraft adaptor cone  flange) or protruding diameter (such as an engine bell), and is independent of the vehicle’s dump rate at  separation.  While this does not account for secondary effe cts like residual engine th rusts and separation  event side loads, tank slosh or friction which can either help or hurt this clearance, these effects are  typically secondary to the overall conic area that the bodies follow dynamically upon separation. For our  case, assuming no external forces, the separation event has been determined to occur naturally for the approximately 28.5 degrees of rotation needed and will provide a minimum clearance of 0.173 m (6.8 in)  for all significant body rates. This is primarily due to the location of the Vehicle Stack system Cg, which is  located very close to the separation plane and is therefore very sensitive to any changes in that location.  Taking this concept further, separation cases were run (see Figure 13) in which the same conditions were  applied to a vehicle with an adequately large recess diameter (SA flange diameter) and to an undersized  flange diameter (124 in). As can be seen, cases were run with the US engine on or off for comparison. As  the lowest curve shows for a smaller SA flange of 3.15 m (124 in), spring force is dependent on dump rate as the higher dump rates require much larger spring stiffness to clear upon separation, while for a SA  flange ID of 3.4 m (top curve) no springs are required at any dump rate even with the US engine on (2nd  curve). The vehicle separates naturally without help of any kickoff device. -0.1-0.0500.050.10.150.2 10 15 20 25 30 35 vehicle dump rate, deg/sclearance, m Contact Positive Clearacne3.4m (135") flange dia., No actuation or J-2X .051 m Re quired Clearance 4,893 kg/m 8,018 kg/m 6,250 kg/m 10,358 kg/m 3.4m (135") flange dia. , No actuation w/J-2X 3.15m (124") flange dia., No actuation w/J-2Xadded spring  stiffness to meet req'd clearance Figure 13. Engine Nozzle Clearance with No Separation Forces (no springs or actuators)  Lessons Learned  One of the lessons learned was that intelligent preliminary sizing of spacecraft geometry can greatly  improve reliability and save on vehicle weight. Also for two connected bodies rotating at a fixed dump  rate, that same angular rate will be maintained by each body after separation, as angular momentum is  conserved. Another important lesson learned was that separation mechanism component mass can be  357
Focus Mechanism for Kepler Mission 

Kraig Koski*

Abstract 

The Focus Mechanism built for the primary mirror on the Kepler mission provides a method for  adjustment of the mirror position for the duration of the mission. The Focus Mechanism also provides  structural support for the 87 kg primary mirror. The Kepler mission requirements provided some interesting and difficult design tasks for the Focus Mechanism. This paper will describe the development,  design, function and testing of the Focus Mechanism.  

Introduction 

The goal of the Kepler mission is to survey our region of the Milky Way Galaxy to detect and characterize  hundreds of earth-size and smaller planets near the habitable zone. The habitable zone encompasses  the distances from a star where liquid water can exist on a planet’s surface. The transit method will be used for detecting extrasolar planets. A transit is when a planet crosses in front of its star as viewed by an  observer, resulting in a small change in the star’s brightne ss for a repeatable amo unt of time. Once  detected, the planet’s orbital size and mass can be calculated using Kepler’s Third Law of planetary motion (T 2 = R3). The size of the planet is found from the depth of the transit (how much the brightness of  the star drops) along with the size of the planet’s star. From this information, the planet’s characteristic  temperature can be calculated.   

Figure 1. Kepler Flight Segment [1]. 

                                                       * Ball Aerospace & Techno logies Corp., Boulder, CO 

Proceedings of the 39th Aerospace Mechanisms Symposium, NASA Marshall Space Flight Center, May 7-9, 2008  359
Kepler Mission Design   For a planet to create a transit visible from our solar system, the orbit must be li ned up edgewise to us.  The probability for a planet in an Earth-like orbit around a solar-like star to be properly aligned is 0.5%.  For this reason, one must look at thousands of stars to determine if Earth-like planets are common or rare. The time between transits for planets in the habi table zone is around 1 year and to reliably detect a  sequence, four transits are required. The Kepler instrument, called a photometer, has a large field of view  (12° diameter) in order to observe more than 100,000 stars in the Cygnus Region of the Milky Way continuously for the entire 3.5 year mission as shown in Figure 2. Kepler is scheduled to be launched in  February, 2009 on a Delta-II rocket into an Earth-trailing heliocentric orbit with a period of 372.5 days  which provides the optimum Sun-Earth -Moon avoidance criteria [1]. 

Figure 2. Kepler field of view in Milky Way Galaxy [2]. 

Kepler Flight Segment   The Kepler flight segment, which was designed and fabricated at Ball Aerospace & Technologies Corp.,  consists of the Photometer mounted onto a Spacecra ft as shown in Figure 1. The Spacecraft provides  power, pointing and telemetry for the Photometer. Pointing at a single group of stars for the entire mission greatly increases the photometric stability and simplifies the Spacecraft design. The Photometer, shown  in Figure 3, is a specially designed Schmidt telescope with a 0.95-meter diameter aperture and an array  of 42 CCD detectors. Each 50x25 mm CCD has 2200x1024 pixels, which are read every three seconds to prevent saturation. The CCD’s are not used to take pictures and the images are intentionally defocused to  10 arc seconds to improve the photometric precision. The instrument has a spectral bandpass from   400 nm to 850 nm. Data is stored on the spacecraft and transmitted to the grou nd once per week [1]. 

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Figure 3. Kepler Photometer, Section View. 

Primary Mirror   Three Focus Mechanism units are used to support and focus the primary mirror as shown in Figures 4 &  5. Each Focus Mechanism attaches to the primary mirror with two flexured struts resulting in a hexapod  configuration. The low CTE FRIT bonded primary mirror is made from ULE Titanium Silicate and is 1.45 meters in diameter with a mass of 87 kg. The three Focus Mechanisms are situated at 120º to each other  and mounted to invar inserts embedded in a composite bulkhead. 

Figure 4. Solid Model of Kepler Primary Mirror/Focus Mechanisms/Aft Bulkhead Assembly.  361