ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
37 forces of the deploying geometry of the antenna that was calculat ed by kinematic modeling. The  offloading force was provided by compression springs that were interchangeable, allowing offloading  forces to be varied quickly and different equivalent offloading to be tested easily. This offloader design  was straightforward and adjustable; however it interfaced to the base ring of the antenna and was incapable of being used once the antenna was integrated to  the rest of the satellite. This meant that the  antenna would need to be completely qualified before de livery to the system. “Pop and catch” deployment  tests instead of full deployments were done for a ll system level antenna deployment testing. 

Figure 7: Gravity Offloading Fixture      EDU Antenna   An engineering development unit (EDU) was fabricated and fully tested in order to gain confidence in the antenna design. The EDU antenna was the full scale flight design and was completed in February of 2007. Testing included in air deployment testing, bas eline RF patterns and PIM measurement, surface RMS  measurements, three axis quasi stat ic and random vibration, TVAC cy cling and deployment testing, and  finally post environmental RF pattern and PIM measurements. In total, the EDU antenna had a total of 22 open/close cycles including 4 in air and 3 TVAC deployment tests.  The EDU antenna passed all tests, however there were several noteworthy lessons learned:  Spring Cartridge Lock Pins

Initial deployment testing and surface mapping reveal ed that the locking mechanisms in the spring  cartridge design did not function as intended. The de sign did not adequately take out the end of travel  slop of the rib. The problem was found to be a combinat ion of both tolerance stack-up of all of the rotating  surfaces and the large moment arm of the deployed rib. This issue needed to be fixed in order to meet  the deployed frequency (>5-Hz first mode)  requirement. To fix the problem, it was decided to remove the  NASA/CP-2010-216272
39 did not contact, they would spring back into the deploying antenna. A damping method was required to  absorb some of the energy to prevent the ends fr om striking the gores or payload and becoming  entangled. The ultimate solution was to attach two welding rods to the Elgiloy band with heat shrink tubing as a damper/stiffener. The solution was tested with a 1G gravity assist to prove the end fittings  would never come into contact.  Reflector CTE

During TVAC testing of the EDU ant enna, a cold spike test  was conducted to dem onstrate survivability  and performance at the coldest predicted temperatures. During the test, the antenna was taken cold (~-50 deg C) and deployed with ~-0.3 g offload giving a wo rst case energy deployment. After the deployment  was completed, the shrouds of the chamber were set to -150 deg C and the antenna was allowed to be  exposed to a cold spike. The reflector has very lo w thermal mass and quickly moves with the shrouds of  the chamber while the rest of the antenna remains within  survival temperature limits. The purpose of the  test is to show that the EDU gores were properly de signed for on orbit coefficient of thermal expansion  (CTE) effects. The reflector material is made primarily of Kapton ® which has a fairly high CTE. This is a  potential issue because if the reflecto r material shrinks too much it would go taut and would exert force on  the deployable ribs. This force would act to move t he ribs back towards the stowed condition, altering the  deployed shape of the antenna. Without a latch, the ribs could move enough to change the shape and/or pointing of the antenna as a function of temperature. The advertised CTE of the reflector material is 25 ppm/degC and the gore was sized such  that the extreme cold temperat ure would not allow the gores to  go tight while meeting RMS requirements at the predicted hot condition. During the cold spike test, it was  visually observed that the deployed ribs did move in t he stowing direction indicating that the reflective  material did get tight. More slack gore was added to solve this issue. This in turn adversely affected the RMS of the reflector surface at hotter temperatures. T he solution to this problem was to loosen the RMS  requirement of the reflector to 12.7 mm (0.50 in) RMS,  resulting in a link margin reduction of ~0.2 dB,  which was deemed acceptable by the program. The flight gores added 6.35 mm (0.25 in) additional  material from rib to rib which proved to solve the CTE issue. 

Figure 9: Mockup FM Gore with Added Material     Gravity Offloading Issues    Deployment testing showed that the gravity offl oading was not acting as intended. This problem  presented itself as the antenna stopping short of the fully open position. The deployment force required to  Added Gore Material  NASA/CP-2010-216272
40 gap the gravity offloader in the deployed position was m easured and compared to the theoretical model.  These measurements showed a discrepancy between predicted and as-built  gravity offloading forces at  the open position. Further analysis showed that the as-built gr avity offloader force profile was offset which  resulted in both under-offloading in some position s and over-offloading in other positions over the  antenna deployment. The issue was traced to a small lin k chain that was used to transfer the offloading  spring force over the cam profile and onto the antenna rib. The chain restricted movement along the  profile of the offloading cam. The “as-built” offloader des ign was capable of testing either energy margin  or force at fully deployed position but not both simultaneously. The offl oader was used in the “as-built”  configuration for all remaining testing and was adju sted accordingly to get the desired energy or force  offload.  Surface Mapping

The reflector was mapped using laser scanning phot ogrammetry; a technique which is often used in  reverse engineering applications. Because of cost limitations the antenna was mapped at room temperature in a “cup sideways” orientation only; that is the boresight of the antenna is parallel to the ground. The loose RMS accuracy required of this an tenna allowed this simplified measurement setup to  be used. The surface accuracy of the antenna was loose enough that the antenna met the requirement  even with the 1g sag of the reflector. Several setups  were used to record the shape of the reflector and  the data was processed to find the RMS deviation of  the antenna from the ideal parabola. The antenna  RMS deviation was measured to be 6.35 mm (0.25 in) from the theoretical surface. The measurement  technique worked well, however it was necessary to find a laser scanner that was capable of imaging a surface with optical properties of both the Kapton used for the EDU antenna and the germanium  sputtered Kapton ® used for flight model (FM) antenna. T he laser scanner that was found to be adequate  was a Cyrax 2500.  

Figure 10: Surface Mapping Results     FM Antenna   The FM antenna was built following the completion of a ll EDU testing. The flight  antenna incorporated the  lessons learned on the EDU antenna discussed in this paper including changes to the spring cartridges,  gores, and deployment strap. The FM antenna went through identical mechanical and RF and PIM testing  as the EDU antenna. This included in-air deploy ments, surface mapping, RF and PIM performance  NASA/CP-2010-216272
43 A Completely New Type of Actuator  -or-  This Ain’t Your Grandfather’s Internal Combustion Engine 

Brian W. Gore*, Gary F. Hawkins*, Peter A. Hess*, Te resa A. Moore* and Eric W. Fournier* 

Abstract 

A completely new type of actuator – one that is proposed for use in a variety of environments from sea to land to air to space – has been designed, patented, buil t, and tested. The actuator is loosely based on the  principle of the internal combustion engine, except that it is a completely closed system, only requiring electrical input, and the working fuel is water. This paper outlines the theory behind the electrolysis- and ignition-based cycle upon which the actuator operates and describ es the performanc e capability test  apparatus and results for the actuator. A mechanism app lication that harnessed the unit’s power to twist a  scaled rotor blade is also highlighted.    Introduction 

Traditional electro-chemical actuators use electr olysis to produce hydrogen and oxygen to generate  pressure that can be used for mechanical work. These actuators can create substantial forces over large strokes but are inefficient and relatively slow. The Aerospace Corporation developed an actuator in which  the hydrogen and oxygen are produced el ectrochemically and then ignited when actuation is desired. The  process takes place in a closed volume so the water produced during combustion is contained and used again. The actuator essentially acts as a power amplifie r that converts a low electrical power into stored  chemical energy. The stored chemical energy can t hen be rapidly released (~10 ms) to produce high  mechanical power in the actuator.  The advantage of this device is the extremely high energy density that can be safely stored, then delivered instantaneously. It is essentially a power amp lifier – a low level of power can be input over an  extended period of time, then quickly reclaimed, on t he order of single-digit milliseconds. The time  duration of the actuation pressure pulse can be controlled from tens of microseconds to tens of milliseconds by adding variable quantities of inert “bu ffer” gas, which slows down the thermal activity  during combustion. In the design space of energy densit y vs. speed of operation, this actuator lies outside  the current state-of-the-art envelope, as shown in Figure 1.                                                    * The Aerospace Corporation, El Segundo, CA 

Proceedings of the 40th Aerospace Mechanisms Symposium, NASA Kennedy Space Center, May 12-14, 2010  NASA/CP-2010-216272
44 Figure 1.  Actuator Specific Power vs. Actuation Rate for State Of The Art  (Courtesy Of Bell Helicopter) 

The heart of the closed-cell electrolysis/combustion cycle (shown in Figure 2) uses Nafion® material  submerged in a small, fixed amount of water, which does not require replenishment.  

Figure 2.  Schematic Drawing of the Actuation Cycle  When an electric current is passed through the Nafion®, the surrounding water undergoes electrolysis to  produce hydrogen and oxygen molecules that bubble to the surface as gases. Both gases are contained  within the common volume until a spark- or glow-plug is energized which then ignites the hydrogen in the  system, using oxygen as the oxidizer. A large pressu re pulse results from the combustion, from which  mechanical work can be extracted, similar to that in  a piston/cylinder system of an automobile’s internal  combustion engine. The difference between the two is that there is no exhaust; after combustion of the  hydrogen and oxygen, the reactants are transformed in to high-pressure steam that eventually condenses  back into water, collecting in the reservoir. The proc ess can continue to cycle, as long as electricity is  provided to the Nafion® to drive the electrolysis. 

Development of Prototype Actuator 

A typical experiment during the actuator development  cycle used a prototype actuator unit, shown in  Figure 3, which was not optimized for weight or  volume. The actuator was a piston and cylinder  arrangement that pushed against a spring typically us ed in automotive applications. Electrolysis was  performed using 7.5 W (3.75 volts and 2.0 amps) for 180 seconds, followed by ignition.  s VH2H2 O2 V++H2H2 O2Steamcooling Cell   h a l f f ill e d  w it h  w a t e rCu rrent gene rates H2 and O2Spark causes detonationSteam e x p a n d  ce l l 0.0010.010.1110100 0.01 0.1 1 10 100 1000Power/Weight (HP/lb)    Actuation Rate (Hz)Actuator Power Available per Mass Electro-Chemical (Single Cycle) Rare Earth Motor Hydraulic MSMA AC motor cooledSwitch Reluctance MotorSMA PZT (MIT)DC motor SMA-cooled Terfenol 1991 Ball-Screw MagnetostrictiveState of the ArtElectro-Chemical (incl. charge time)Electro-Chemical (Single Cycle) Rare Earth Motor Hydraulic MSMA AC motor cooledSwitch Reluctance MotorSMA PZT (MIT)DC motor SMA-cooled Terfenol 1991 Ball-Screw MagnetostrictiveState of the Art NASA/CP-2010-216272
45  

Figure 3.  Piston/Cylinder Actuator (Left), And Spring It Compressed (Right) During Actuation  The resulting force and displacement profiles are shown in Figure 4. The force peaked at approximately  2.60 kN (580 lb), while pushing 1.8 cm (0.7 in) agai nst the 145 kN/m (825 lb/in) steel compression spring.  The force curve (Delta Load, in black) was measured by a load cell at the fixed end of the spring and  contains additional temporal structure due to compression waves generated within the spring. The full-width-at-half-maximum duration of t he force pulse was approximately 5.5 ms. Peak internal pressures of  31 MPa (4500 psi) have been attained inside the cy linder and are limited only by the strength of the  cylinder itself, not the elec trolysis/ignition process. 

Figure 4.  Load and Displacement of the Actuator Pictured in Figure 3  This example demonstrated the application of 7.5 W fo r 180 seconds and the achievement of >2500 W of  instantaneous usable power. Again, the amount of safe potential energy stored , and power delivered, by  the combusting hydrogen and oxygen is only limited by  the structural integrity of the cylinder. Gas  generation rate is proportional to elec trolysis current; the more time-int egrated current that is used, the  more gas will be generated in a given time, and the more power will be available.    Development of Trigger-based Power Meter 

One of the other metrics in this development effort wa s to determine and maximize the specific power of  the actuator, with a goal of 16.7 kW/kg (50 hp/lb). Initially, the piston/cylinder pictured in Figure 3 was  used with an original test “bomb” mounted to its underside as the combustion chamber (also not optimized for weight, but used readily available, off-t he-shelf parts). A power density for that combination  of parts – about 13.1 kW/kg (39 hp/lb) – was calculat ed through knowledge of pressure, surface area, and  piston velocity/acceleration.   A custom, lighter-weight chamber was then designed and built for the purpose of lowering the entire  Electrolysis Combustion Actuator (ECA) device’s mass, thus increasing the overall specific power density of the device. This lightweight ECA (195 g/0.43 lb) c onsists of a thin-walled gas generator and ignition 

02468 1 0 1 20100200300400500600700 0.00.20.40.60.8   Delta Load Time  from Ignition (ms) Di sp lac ement     Displacement (in.) Time From Ignition (ms) 

Displacement (in)    Force (lb)  NASA/CP-2010-216272
46 chamber, has ports for the igniter and buffer gas inlet, and has feed-throughs for gas generation. It  replaces the original, heavy flange-based CVR (Constant  Volume Reactor), (1.42 kg/3.14 lb). Figure 5  shows a close-up of the newly built, lightweight ECA chamber.  

Figure 5.  Lightweight Electrolysis Combustion Actuator (ECA)  A special test fixture was also developed. The fixture design needed to restrain the stroke of the piston  from the build-up of pressure during the electrolysis phase, but possess a trigger function to allow the piston to begin to stroke instantly after ignition. The g oal was to make this a passive device, rather than to  try to incorporate a complicated feedback release syst em. The trigger device in Figure 6 was designed. It  uses the basic principle of a quick-disconnect, ball rel ease mechanism which will not allow relative motion  until a specifically-applied external force is present; in this case, that force comes from the pressure pulse  above and beyond the electrolysis static pressure. Rollers were used instead of spherical balls to allow for greater Hertzian contact load capability without  brinelling. Accommodations were made for a variablesized spacer to allow for different threshold pressures at which the device would release.  

Figure 6.  Quick-Disconnect Ball Release Test Fixture     VariableSized Spacer  Electrolysis  Combustion  Actuator (ECA) Belleville  Washer stack Rollers (4),  captured in  light blue slots  O-Ring on  Piston HeadAccelerometer  Detents (4), in  orange block  NASA/CP-2010-216272
47 The following derivation was used to calculate the horsepower being generated by the actuator on the  moving mass inside the test fixture:  Power = Energy / time  Power = Force * distance / time Power = Moving mass * acceleration * distance / time Power = Moving mass * acceleration * velocity           [1]   For the test, the moving mass in the apparatus (everyth ing except the gray and green structural pieces) in  Figure 6 was 0.2 kg (0.45 lb). The total displacem ent allowed before an external crushable device was  impacted was 9.65 cm (3.8 in). Note that the device it self has an allowable stroke of 1.7 cm (4.6 in).    Figure 7 is a picture of the entire assembled test fixture hardware, including the ball release trigger  mechanism and the ECA. Note the inclusion of the ac celerometer mount at the top which serves a dual  purpose as both a structural mount for the accelerometer, and also as a hard stop for the moving mass as it impacts the external crushable honeycomb unit mounted to the rigid test stand above. 

Figure 7.  Complete Peak Power Density Test Fixture 

Figure 8 depicts the entire test station where the sp ecific power measurements were made. It includes  the location of the test fixture on the stand, the crushable shock absorber, and the black and white  measurement scale. The calibrated scale was placed behind the fixture and used as a reference to easily determine position and velocity from the high speed video camera output.      New ECA  Quick-Disconnect Ball Release FixtureAccelerometer   NASA/CP-2010-216272
48   Figure 8.  Peak Power Density Experimental Test Setup at The Aerospace Corporation  Figure 9 shows the first frame of the high-speed vi deo. The accelerometer mount can be seen at the  bottom of the frame, with a clear horizontal edge to us e as a reference on the vertically moving part. The  measurement scale is shown along the right side, and the crushable device is at the top of the frame.  

Figure 9.  First Frame of Peak Power Testing High-Speed Video   Hi-Speed CameraMeasurement Scale    Test FixtureCrushable  Honey-comb Unit Lights (2)   Crushable Unit  Measurement  Scale Accelerometer Mount NASA/CP-2010-216272
49 Each of the measurement marks (black or white) was measured to be 2.08 mm (0.082 in) (average of  many stripes divided by the number of them), and the observation points taken from the video were to  count the number of marks travelled by the ac celerometer stand in each frame (=0.1 ms).  

The actual “run-for-record” utilized 1.38 MPa (200 ps ig) of argon buffer gas in the chamber, and about  0.48 MPa (70 psi) of electrolysis-generated H 2 and O 2. When combusted, the gases propelled the moving  mass a distance of 9.58 cm (3.77 in) in 2.9 ms and compressed the crushable device (which had a   26–36 kN (6000-8000 lb) crush load) approximately 0. 5 cm (0.2 in). That translates to an average  velocity  of 33 m/s (1300 in/s), which for reference is almost  119 km/hr (74 mph). The raw displacement-vs.-time  data were surprisingly smooth, but when differentia ted to obtain velocity and then again to obtain  acceleration, the data became increasingly noisy. To solve this problem, a best-fit polynomial was found  to represent the displacement-time data, which could then be more easily differentiated. Since the displacement and velocity are initially zero, the constant and first-order coeffi cients of a fourth-order  polynomial were forced to be zero, and the result was Equation 2.    D = 51,746,636,636*t 4 – 396,904,167*t3 + 1192091*t2            [2]  where: D = displacement (in), and t = time (s)   This polynomial had an R 2 correlation of 0.9997 to the raw di splacement data. With such a strong  correlation factor, the best-fit method is strongly supported. The pl ot in Figure 10 was created after  Equation 2 was differentiated for velocity and then agai n for acceleration. The non-zero acceleration is  real, since there is a force acting on the body (f rom the combustion) before it starts to move. 

Figure 10.  Displacement, Velocity, & Acceleration of Moving Mass   As illustrated in Figure 10, the pe ak of instantaneous horsepower occurr ed approximately 0.7 ms after the  start of motion and was calculated, using Equation 1, to be 134 kW  (180 hp) . The maximum velocity  achieved before the end-of-travel impact was almost 160 km/hr (100 mph). However, since the milestone  to be met was specific power, not total power, that  value had to be divided by the mass of the entire  device that produced it. The total mass of test fixture’s mechanism/cylinder (all of the hardware above the “New ECA” in Figure 7), is 0.39 kg (0.85 lb) with the moving portion at 0.20 kg (0.45 lb) and the fixed portion at 0.18 kg (0.40 lb). The mass of that ECA is 0.43 lb (the old CVR “bomb” was 1.42 kg (3.14 lb),  for reference). Therefore, the calculation of the demonstrated peak specific power is  180 hp / (0.85 + 0.43) lb = 141 hp/lb (47.1 kW/kg) ,   [3] 0.0000.5001.0001.5002.0002.5003.0003.5004.0004.500 0 0.0005 0.001 0.0015 0.002 0.0025 0.003 Time (sec)Displacement, Velocity, & Acceleration Disp (in) Vel (1000 in/s) Accel (1000 g's)Instant of Peak hp  NASA/CP-2010-216272
51 Operation of the actuator requires power at constant current for each electrochemical unit to “charge” and  power at different constant current for each igniter to “fire.”     The electrochemical-combustive units contain pressurized argon gas that modifies the combustion and heat transfer characteristics which, in turn, optimiz es the transfer of energy of combustion to piston  motion. The electrochemical-combustive units have valv es that allow the argon pressure to be adjusted  prior to initial use and during maintenance.  The electrochemical-combustive units also contai n water to keep the electrochemical components  submerged. In a spinning rotor blade, known water positi on is achieved by centrifugal force. In the testing  environment, water position is achieved by orienting t he actuator vertically with the piston shaft at the  bottom.   Figure 12 shows a cross section of the actuator with the piston, shaft, and components of the two  electrochemical-combustive units labeled. Note that  the combustion chamber of the number 1 unit is  above the piston, and the combustion chamber of the number 2 unit is the volume around the shaft. The  volumes of these two chambers are approximately e qual. The 3-D rendering on the right is shown at a  slight angle to better show  the piston shaft. The mounting rod, render ed in light green, is located near the  bottom protruding to the right. The electrochemical units are shown in dark blue. The igniters and argon fill valves are shown in brown.   

Figure 12.  Bidirectional Actuator  for the Rotor Blade Application  Figure 13 is a photograph of the actuator development hardware in a laboratory  test stand. The photo  shows black tubes and fittings connected to the argo n fill valves. These valves are used to fill the  electrochemical combustive units with argon and would be removed for normal operation in the blade. Note the electrical wires (white, yellow, blue) connec ted to each igniter and to the electrochemical areas.  

Electrochemical  area  Electrochemical  area Argon Fill ValveIgniter Igniter Argon Fill Valve Piston  Shaft Combustion  area  Combustion  area Electrochemical -  Combustive unit 1  Electrochemical -  Combustive unit 2  NASA/CP-2010-216272