APPLYING FIBER SUPPORT TECHNOLOGY to SMALL SATELLITE SYSTEMS I Scott M
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I I APPLYING FIBER SUPPORT TECHNOLOGY TO SMALL SATELLITE SYSTEMS I Scott M. Jensen* J. Clair Batty * * I Dave McLain * * * Utah State University/Space Dynamics Laboratory I Logan, Utah 84341 I Abstract A support system designed, I Cryogenically cooled fabricated and tested for application components in infrared instruments on the SABER instrument utilizes designed by Utah State University's strands of Kevlar 49 in a bicycle I Space Dynamics Laboratory wheel-spoke like arrangement to traditionally have been mounted on support a focal plane assembly (FPA) glass-epoxy composite (G-10) cooled by a miniature pulse tube I cylinders for thermal isolation. refrigerator. The first resonant Ensuring adequate mechanical frequency of the Kevlar supported stiffness to withstand typical launch assembly has been measured to be I loads often compromises the desired greater than 600 Hz in all axes. This thermal isolation and conduction compares with typical values of 50 to parasitic heat loads become a 70 Hz for a similar assembly I concern. Beginning with a senior supported by concentric G-10 design project in the Mechanical and cylinders. In addition to an order of Aerospace Engineering Department, a magnitude increase in the mechanical I new approach to rigidly supporting stiffness, the Kevlar system reduced cold components using high parasitic conduction heat loads by performance fibers in tension was almost two orders of magnitude. I initiated. The development effort Calculated conduction loads through included consideration of several the Kevlar strands total less than 1 candidate fibers for tensile strength, mW as compared to 85 mW for the G- I shear strength, creep, and thermal 10 supports. This dramatic reduction conductivity as well as a technique for brought total parasitic heat loads to attaching and tensioning the support within the limited cooling capacity of I strands. the miniature refrigerator and made possible the SABER instrument with its very stringent power and mass I constraints. Further applications for * Graduate Student, Mechanical and Aerospace Engineenng this technology are currently being •• Professor, Mechanical and Aerospace Englneenng designed. I ,. 'Mechanical Designer, Space Dynamics Laboratory I I I Nomenclature view requires very massive and strong FIST Fiber Support Technology support structures. From a thermal I MLI Multi-Layer Insulation stand point, the structure should be MTS Mechanical Testing small with minimal cross-sectional System areas to reduce the parasitic heat I SOL Space Dynamics transfer. Laboratory Traditionally, cold components SABER Sounding of the developed by Utah State I Atmosphere using University/Space Dynamics Laboratory Broadband Emission (USU/SDL) have been supported by I Radiometry concentric composite G-1 0 USU Utah State University cylinders(See Figure 1). FPA Focal Plane Assembly This approach has two I MAE Mechanical and limitations that are becoming more Aerospace Engineering serious in the new "smaller, faster, A Area cheaper" environment of the 90's. I d Diameter The first resonant frequency of E Modulus of Elasticity these types of support systems, F Force depicted in Figure 1, is roughly 60 I k Thermal Conductivity Hertz. Higher values are desirable to K Dynamic Spring Constant avoid the risk of resonant response L Length during launch. I q Heat load e Strain I s Stress w Natural Circular Frequency I MU I. Introduction I A major challenge faced by those developing cryogenically cooled components for space applications is I to provide both adequate mechanical support and thermal isolation. Achieving the thermal isolation G-10 I objective is often in direct conflict MLlJ with the mechanical support requirements. Precisely locating and I rigidly supporting cooled components under harsh dynamic conditions, such as those experienced by a rocket ride I into space, from a mechanical point of Figure 1 Traditional G-10 Support Method I 2 I I I I Also, the parasitic heat load rather stringent mass constraints due to conduction through the G-1 0 precluded the use of an expendable support system has a significant cryogen as the heat sink. Power I impact. For systems using constraints also eliminated standard expendable cryogen as the heat sink, size mechanical refrigerators with mission life is limited. In systems roughly 100 W input power I using mechanical coolers, the cooling requirements. A miniature stirling capacity may be exceeded. For some Cycle refrigerator developed by TRW time cryogenic systems engineers at for the Brilliant Pebbles program was I USU and SOL have discussed the proposed provided a reduction in total possibilities of supporting cold parasitic heat loads to less than the components with tension strands in miniature refrigerators cooling I an arrangement something like the capacity of 250 mW at 72 K could be spokes in a bicycle wheel. A group at achieved. The major contributor to NASA Ames [1] reported development the FPA parasitic heat load was 85 I of a thermal isolating support bench mW conducted through the concentric that used Kevlar strands in tension. As G-10 cylinders supporting the focal an MAE senior design project in the plane. We then conceptualized a I spring of 1993, two student teams design to utilize FIST mounting on the accepted the task of further SABER FPA. developing Fiber Support Technology I (FIST) by employing ultra-low thermal conductivity fibers (e.g., Kevlar, Vectran) in tension to provide I adequate mechanical support and thermal isolation of system components I Two prototypes, very different in appearance from the NASA Ames fixture, were designed, constructed I and tested with promising results in I the Fall of 1993 and Winter of 1994. II. Application to SABER Epoxy Filled Bolt I A possible application for the USU FIST technology was already being discussed. SOL was proposing I to build an infrared limb scanning instrument called SABER in a teaming arrangement with I\JASA LaRC. I SABER's detectors had to be cooled to below 75 K. A desired mission life of 2 years or more together with I Figure 2 Fiber Support Technology Scheme I 3 I I III. Support Structure Ultimate Strength Test I The support structure consists Kevlar 49 strands were put of four basic components; A warm through a series of tension tests to I outer support mechanism, a cold inner verify the ultimate strength. Both support plate (disk), fiber fasteners, ends of the strand were wrapped and fiber woven into appropriate size numerous times around a large radius I strands. The support structure, base so as to minimize shear loads. lined for SABER, is shown in Figure 2. Multiple tests were performed on an Size and configuration constraints of MTS machine and data collected I the outer and inner support structure electronically by a computer. All are imposed by the system in general. specimens failed mid-strand thus verifying no significant shear loading I IV. Fiber Selection at the end connections was present. The backbone of FIST is the Figure 3 shows the ultimate strength fiber. It provides the necessary of the Kevlar 49 strands selected for I thermal isolation as well as the SABER. mechanical rigidity for the system. To date, Carbon, Graphite, Glass, Kevlar, I Vectran, Spectra, and Nomex fibers !l KEVLAR 49 COMPARISON have been considered. Metals are ULTIMATE STRENGTH TEST also of interest because of their I 140 ~----------------' excellent mechanical properties, and J may find application when slightly 130 larger heat loads can be tolerated. 120 I 110 For the SABER instrument, several fibers were considered but Kevlar 49 was chosen for supporting ! i I o I ' the cold FPA and the chilled telescope ~ 70 ~i I I within the spacecraft because of its ..J light weight, high tensile strength, and I low creep characteristics when compared with the other fibers, in 30 addition to its extremely low thermal 20 I 10 conductivity of (.04 W/m-K) [2,3]. o I Fiber Testing i 'I l _~~~_~ __~~~~_~= Ultimate strength, shear Figure 3 Kevlar 49 Ultimate Strength Test I strength, and creep were of the most Results concern, hence, tests were performed to verify data on these mechanical I properties as provided by the fiber manufacturers. I 4 I I I technology evolution by modifying feed housing I length and adding or subtracting standardized modules or custom snap-in elements as required. Micromachined Components I Like any filter, bandwidth, roll-off, and insertion loss are driving requirements. I Traditional strip line and waveguide filters are large, even at the millimeter wave frequencies. Evolution of semiconductor etching processes I over the last few years has yielded a host of micromachined components ranging from I electro-mechanical actuators to simple filters. Micromachined filters provide performance comparable to traditional counterparts with I reduced production cost and dramatic reductions in size and weight. Filter parameters are tightly 1.4 ~Lm·Thick controlled by the masking and etching process, Membrane I so variations in critical performance parameters over time and between filters of a given design I should be minimal. This stability enables automated .production and test with minimal tuning. Micromachined filters use proven silicon I etching and deposition techniques to reduce microwave filter size by up to 75%, resulting in I smaller electronics packages. A typical micromachined filter is shown in Figure 3. Micromachined Filters Figure 3. The planar structure combines I standard silicon etching and metal deposition is mounted on a nadir-facing panel of the techniques with emerging membrane technology ADEOS II spacecraft and consists of a rotating to produce filters with integral shielding cavities antenna subsystem, the Scatterometer I etched into the two laminated substrates. The Electronics (SES) and a Command and Data center silicon wafer is etched through. The System (CDS). The complete instrument weighs I resulting sandwich resembles a waveguide filter in at 176.8 KG and draws 234 Watts of prime in performance at 25% of the volume.