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High Temperature Superconducting Space Experiment II (HTSSE II) cryogenic design
Cryogenics 36 (1996) 741-752 Published by Elsevier Science Limited Printed in Great Britain 001l-2275/96/$15.00
SOOll-2275(96)00036-7
ELSEVIER
High Temperature Superconducting Space Experiment II (HTSSE II) cryogenic design T.G. Kawecki,
S.S. Chappie and D.R. Mahony*
Naval Research Laboratory, Naval Center for Space Technology, Washington, DC 20375, USA *Swales & Associates Inc., Beltsville, MD 20705, USA
Code 8223,
At 60 to 80 K large performance gains are possible from high temperature superconducting (HTS) microwave devices for communications applications. The High Temperature Superconducting Space Experiment II (HTSSE II) will demonstrate eight HTS experiments in space for up to 3 years of operation. HTSSE II is the first application of HTS technology to space. In addition to demonstrating HTS devices, an important secondary goal is to demonstrate the cryogenic technologies required for long life HTS space applications. HTSSE II utilizes a British Aerospace 80 K Stirling cycle cryocooler to refrigerate a central cryogenic bus of seven HTS experiments and has an additional stand-alone TRW HTS experiment cooled by a TRW Stirling cycle cryocooler. The HTSSE II flight unit has been assembled and has successfully passed vibration and thermal vacuum environmental tests. HTSSE II was developed on a fixed budget and a fast track schedule of 24 months and is due to launch in March 1997 on the ARGOS spacecraft. This paper presents the design and test results of the cryogenic subsystem, cryocooler integration and a cryogenic coaxial cable I/O assembly. Published by Elsevier Science Limited Keywords: high temperature wave devices
superconductivity;
The High Temperature Superconducting Space Experiment (HTSSE) programme was initiated by the US Naval Research Laboratory (NRL) to promote and demonstrate the potentially significant benefits of utilizing high temperature superconducting (HTS) materials in satellite electronic components. At 60 to 80 K large performance gains are possible in thallium and yttrium based HTS microwave devices used for communications applications. The programme goal of HTSSE is to provide a focal point for the HTS research and development community to develop and demonstrate space qualified HTS technology. HTSSE I contained 15 HTS experiments which were predominantly passive microwave circuit devices cooled to 77 K by a British Aerospace Stirling cycle cryocooler. HTSSE I was launched in 1993 and was lost before on-orbit start-up when the host spacecraft mission failed. In March 1993, HTSSE II was manifested on the Air Force Space Test Program Advanced Research and Global Observation Satellite (ARGOS) and the flight hardware design portion of HTSSE II began. HTSSE II experiments had already been under development. The primary HTSSE II mission requirement is to demonstrate HTS devices on orbit for one year. It is desired to demonstrate three years’ operation if spacecraft operations funding is available. HTS devices have evolved significantly from HTSSE I
space cryogenics;
cryocooler;
micro-
to HTSSE II by demonstrating higher device complexity and maturity in experiments that integrate HTS devices together or combine with semiconductor devices. Eleven HTS experiments were funded at the start of the HTSSE II programme. Eight experiments were selected for flight on the basis of technical merit, ability to integrate and the ability to survive the qualification test programme. HTSSE II flight microwave experiments include the following device types: channellized receiver, cueing receiver, digital instantaneous frequency measurement unit, delay lines, RF multiplexer and a digital multiplexer. In addition to demonstrating HTS devices, an important goal of HTSSE is to develop and demonstrate the critical cryogenic subsystems required to support HTS space systems. Two key cryogenic technologies required for space HTS applications were identified in the HTSSE effort: long life, high reliability mechanical refrigerators for cryocooling and low RF loss, and high thermal resistance coaxial cables to handle a microwave input and output to cryogenic HTS devices. To date, space applications of HTS devices have been largely oriented toward communications applications. Future cryogenic HTS commercial communication payloads are likely to have large volumes and significant I/O thermal parasitics. This suggests the need for cryogenic
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cooling capacities of the order of 0.5 W or greater in the 60-80 K temperature range. Communications applications typically require lifetimes of 5 years or greater. Relatively large thermal loads over long time periods preclude the practical use of solid cryogens for primary cooling. Cryogenic radiators for primary cooling are undesirable for communications applications in this temperature range because they greatly constrain orbit options and have a large impact on the overall spacecraft design. Long life, high reliability mechanical refrigerators appear to be the best cryocooling source for operational HTS payloads. These coolers offer a relatively large cooling capacity and the potential for lifetimes greater than 5 years. In concert with the HTSSE programme goal of demonstrating HTS support technologies for space, a Stirling cycle British Aerospace 80 K cryocooler is the basis of the HTSSE II cold bus cryogenic system. It maintains seven HTS experiments at 77 K on a continuous basis once on orbit. The eighth HTS experiment on HTSSE II is the TRW 2: 1 Digital Multiplexer which is mounted to a TRW Stirling cycle refrigerator operating in the 60-65 K range. At the time of the 1995 Space Cryogenics Workshop, HTSSE II had successfully completed assembly and environmental testing of the qualification unit and subsequently the flight unit. HTSSE II was then awaiting integration on the ARGOS host spacecraft in February 1996 and launch in March 1997.
Mechanical
configuration
Schedule and budget constraints were significant design drivers from the beginning. The budget had been planned in advance of the ARGOS manifest and was fixed. The schedule was determined by a short notice manifest on ARGOS. At the time of the manifest, HTSSE II was required to progress from concept to flight hardware integration on ARGOS in a 20-month period. HTSSE II HTS experiments were being developed in parallel with the rest of HTSSE II, further complicating the design effort. Due to ARGOS development delays since the start of the programme, the HTSSE II delivery date for ARGOS integration moved from September 94 to February 96. HTSSE II has used this schedule slip to reduce risk by expanding the test plan to increase the original 20 month development effort to 24 months. ARGOS is a three-axis stabilized, 2500 kg, 1 kW satellite to be launched on a Delta 2 launch vehicle in March 1997. HTSSE II is launched at ambient temperatures with only survival heater circuits enabled. HTSSE II starts cooling to cryogenic temperatures after an initial 6-week spacecraft checkout and test phase. ARGOS operates in a 450 nautical mile circular orbit at a sun synchronous inclination. HTSSE II is located on the zenith end of the spacecraft on a 1.02 m x 1.27 m honeycomb deck with volume available on both sides. The HTSSE II mass, volume and power allocations from ARGOS were generous. Next to strict budget and schedule constraints, the cryogenic considerations became the biggest design configuration driver. The heart of HTSSE II is the cryogenic subsystem which consists of a cold bus with seven HTS experiments cooled by a British Aerospace 80 K Stirling cycle cryocooler. The cold bus and experiments are supported by glass epoxy thermal isolation tension straps that are fastened to an exoskeletal cold bus support structure (CBSS) as shown in Fig-
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II: T.G. Kawecki et al. COLD BUS SUPPORT STRUCTURE
COAX i/O INTERFACE PLATE
Figure 1 (CBSS)
HTSSE
II cold
bus and cold
bus
support structure
we I. The CBSS is supported on the HTSSE II honeycomb deck with thermally insulating kinematic mounts. Figure 2 shows the ARGOS spacecraft with HTSSE II installed and a larger exploded view of the HTSSE II deck above without blankets and with some structure removed for clarity. The CBSS with radiators is located just behind the BAe 80 K cryocooler in this view. The cold bus is shown with cryogenic blankets installed. The exterior blanket dimensions are approximately 280 mm diameter by 305 mm long. Ambient temperature electronics are located on both sides of the HTSSE II deck. The TRW stand-alone experiment and cryocooler are located as shown. The overall deck layout of HTSSE II evolved into the HTS Experiment Cold Bus \
Cold Bus Structure +Z Radiator /
Cavity & +Y Radiators
Figure2 load
deck
ARGOS
host
spacecraft
with
detail
of HTSSE
II pay-
High Temperature
Superconducting
present configuration primarily to take advantage of the ARGOS sun synchronous orbit which provides secondary passive cooling to reduce cryogenic thermal parasitics. Radiators at cryogenic temperatures were not possible even in this favourable sun synchronous orbit due to the hot (60°C) +Y solar panel that rotates in a coning motion once each orbit about the Y axis. There were two additional important physical configuration drivers. One was to minimize the potential for electromechanical noise from the cryocooler on HTSSE experiments and instrumentation, and on the ARGOS spacecraft. The cryocooler was located as far as possible from the ARGOS magnetometers to avoid violating magnetic field interface control agreements. The second configuration driver was to locate the cryocooler cold finger in a vertical orientation for all ground thermal vacuum testing to minimize the chance of clearance seal problems during l-g environment operation. Clearance seals are the key to cooler longevity, so much design effort was focused on the cryocooler cold bus interface to maintain clearance seal integrity.
Thermal
design overview
The HTSSE II structure can be divided into the five temperature-controlled zones listed in Table I. The focus of this paper is on the first two zones that are directly related to cryogenic performance: the cold bus and its support structure. Two strategies were employed in the thermal design of HTSSE II: providing high resistance thermal paths between the cold bus and its external environment, and lowering the external temperatures as much as possible. The cold bus is suspended inside a triangular truss with low conductance isolation straps and is enclosed in a 25mm thick cryogenic multi-layer insulation (MLI) blanket. The RF and electrical connections for the HTS experiments on the cold bus are a major heat short through the cyro-blanket. Therefore, low conductance cables and wires are used from the cold bus to the interface plate on the +X end of the CBSS. The suspension straps, I/O cables and all the internal faces of the CBSS that view the cold bus cryo-blanket are covered with 15layer thermal blankets. The CBSS assembly is mounted to the electronics deck (zenith panel) using thermal isolators and functions as a thermal guard around the cold bus. Aluminium closure panels, attached to the -Y and +Z faces of the CBSS, act as radiation guards. The I/O interface plate is mounted to the +X face and is thermally sunk to the CBSS. The radiation guards and the I/O plate are designed to intercept heat coming through the ML1 blankets external to the CBSS. The I/O plate also intercepts heat from the external RF cabling connected to the warmer RF electronics. The interTable 1
HTSSE II
temperature
controllled
zones
Zone
Nominal temperature
Cryogenic bus (cold bus) Cold bus support structure (CBSS) BAe cryocooler TRW cryocooler Support electronics
77 K -50°C o-10°C 0-27°C 0-40x
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cepted heat is rejected to space by radiators mounted to the CBSS. Three radiators are mounted to the CBSS, two for cooling the CBSS structure and one for cooling the CBSS thermal blankets that border the ‘cavity’ around the cold bus cryo-blanket (see Figure I). The cavity radiator and one CBSS radiator are mounted to the +Y face of the CBSS. The other CBSS radiator is attached to the zenith apex of the CBSS. A small extension to the zenith radiator acts as a sun shade for the +Y radiator (see Figures 3 and 4). The +Y face receives almost no direct solar flux due to the sun shade and the nature of the host vehicle’s orbital attitude. The +Y face does have a view of the hot (60°C) +Y solar panel which severely limits its ability to achieve near-cryogenic temperatures (see Figure 5). Positioning the CBSS to the -Y side of ARGOS shields the +Y side of the CBSS from most of the albedo and planetary fluxes and from some of the +Y solar panel IR flux. The CBSS thermal guard operates at about -48°C. Heat from the CBSS passes through the CBSS ML1 and into the cavity between the CBSS and the cold bus cryo-blanket. The cavity radiator positioned on the +Y side of the CBSS has a high emissive radiation coupling to this cavity and is thermally isolated from the CBSS to provide the lowest temperature possible. It is essentially a second-stage radiator. The British Aerospace 80 K Cryocooler is mounted to a bracket, independent of the CBSS and thermally isolated from the zenith panel. The cold finger of the cooler is thermally connected to the -X end of the cold bus with a flexible conductive strap. Lockheed cryocooler drive electronics control the cooler. A PID control circuit within these electronics monitors the cold bus temperature and adjusts the cryocooler compressor stroke to obtain the desired temperature. In addition to the cooler drive electronics, Lockheed Research Laboratory has been a key contributor to the evolution of the HTSSE II cryosystem. Lockheed designed and fabricated the cold bus cryogenic thermal blankets. The cryocooler flex strap assembly is an NRL/Lockheed design and was fabricated and thermally tested at Lockheed.
Cold bus design To reduce risk and meet the ambitious schedule, a cold bus mechanical design similar to the demonstrated HTSSE I design was adopted. The design goal for HTSSE II was to include as many experiments as possible within the limited volume of the cold bus. Due to the increased complexity of the HTS devices being developed for HTSSE II, an effort was made to give the experiment providers considerable flexibility in the packaging of their experiments. As a result, the seven HTS experiments selected for flight on the HTSSE II cold bus are of various shapes, sizes and masses. An exploded view of the cold bus structure on which the experiments are mounted is shown in Figure 6. The three devises and the end plates are the primary load-carrying structure of the cold bus. Axial loads are transmitted through the A286 alloy clevises and radial loads are carried through the two 9-mm thick aluminium alloy end plates. With this load-carrying configuration, the experiment mounting plates were free to take on any configuration convenient to accommodate the experiments, since the plates need only support their own weight and the weight of the
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Figure 3
Superconducting
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+X view of HTSSE II flight unit (external
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MLI blankets removed)
experiments. This design flexibility was important because the experiments were being developed concurrently with the cold bus structure. Careful consideration was given to the thermal performance of the cold bus. The goal was to have a structure that was nearly isothermal at the operational temperature. To ensure good conduction, 9-mm thick aluminium alloy plate was machined to form the experiment mounting plates. A thin layer of epoxy was used at all mechanical joints on the cold bus as a conformal thermal interface to maximize contact area. Temperature sensors located on opposite ends of the qualification and flight cold buses indicated AT values of the order of tenths of a kelvin. The cold bus is suspended inside the CBSS with six Sglass epoxy thermal isolation straps. The six-strap design, based on HTSSE I, minimizes ML1 penetrations to minimize the cold bus thermal parasitics. The negative impact is that rotational modes of the cold bus are not constrained and can be excited if the cold bus is not mass balanced. Excitation of cold bus rotational modes (and coupling between translation and rotation modes) can lead to significant increases in strap loading. The maximum allowable centre of gravity offset was defined (by analysis) to be +12.7 mm axially and It6.4 mm laterally. Any centre of gravity deviation beyond this would require correction with balance weights. The cold bus design did include balance
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weight mounting provisions, but careful placement of experiments on the cold bus structure avoided the need for balance weights on either the qualification or the flight cold bus. All of the RF and d.c. If0 lines on the cold bus were routed to one end of the cold bus and brought through the cryo-blanket in one 6-mm diameter bundle. To reduce the potential for cyrocooler-induced electrical noise in the cryogenic I/O d.c. lines, the cold bus I/O was located as far from the cryocooler as possible. A cylindrical beryllium cold stem, located at the opposite end of the cold bus from the I/O bundle, serves as the thermal link to the cryocooler. The end of the beryllium cold stem sticks out through a penetration in the cryo-blanket surrounding the cold bus and interfaces with a flex strap copper collar. Figure 7 is a photograph of the assembled flight cold bus before the installation of the cryo-blanket and mandrel. Four of the six thermal isolation straps and four of the seven experiments are visible in this view. The thermal link to the cryocooler, the beryllium cold stem, is visible at centre left. Semi-rigid coaxial cables that connect the I-ITS experiments to a coaxial cable interface ring are visible in the centre and at the right side of the figure. The cable bundle at centre right is all of the cold bus I/O brought together to penetrate the cryo-blanket.
High Temperature
Figure 4 -X-Y removed)
view of HTSSE II flight unit; note external
Co;ng
Superconducting
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RF cables and cold bus cavity (cavity radiator and external
MLI blankets
i
3 Nadir Figure 5
+X view of external thermal
environment
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Coldbus Clevis (1 of 3) \
End Plate
Experiment Mounting Plate
Figure 6
Cold bus structure,
Figure 7
HTSSE II flight cold bus prior to cryo-blanketing
Thermal
isolation
exploded
view
straps
As in the HTSSE I design, six S-glass epoxy straps (manufactured by SCI) are used to structurally support and thermally isolate the HTSSE II cold bus from the CBSS. Several strap configurations were considered; the final strap
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design selected is 1.0 mm thick, 5.3 mm wide and has a centre-to-centre length of 102 mm. The ends of the straps fit into grooves cut into stainless-steel spools. The bore of each spool contains a spherical bearing which is included to prevent torsional or bending loads on the straps. A bolt through each spherical bearing attaches the strap/spool assembly to clevises as seen in Figure 8.
High Temperature
Superconducting
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Spool Strap . _
/
, r’,,
II: T.G. Kawecki et al.
Cold Bus Support Structure
(1 of 6) \
‘---_
Outer Clevis
__
Strap
_J
:“I; -
__-
1
Cold Bus Clevis
Clevis ’
ItIf . . .
’ . . .
Figure 8
Cold bus tension strap assembly
design
A high-fidelity engineering model of the CBSS and cold bus was assembled early in the design process to qualify the strap design and to demonstrate fatigue performance of the straps. The fatigue performance of the first lot of straps was lower than expected, but the fatigue life capability was proved through a series of four qualification level random vibration tests. The fatigue testing was performed with the worst case cold bus mass and the worst case centre of gravity offset.
Cold bus support
structure
(CBSS)
As a significant facet of the HTSSE II thermal design, the CBSS was designed to operate at temperatures significantly below the ambient deck temperature. Approximately 0.4 m* of thermal radiator surface area cools the CBSS to reduce thermal parasitics into the cold bus. The CBSS is fastened to the HTSSE 11 deck with kinematic mounts, which prevent the deck from loading the CBSS during launch, allow the CBSS to contract during cooling and help to thermally isolate the CBSS from the warm deck temperatures. Torlon spacers are used on both sides of the kinematic mounts, and fibreglass washers are used under the heads of the mounting bolts for thermal isolation. The CBSS structure consists of three longerons which capture the outer ends of the tension strap assemblies, six angle braces separating the longerons and three ‘X’ shaped cross braces that stabilize the structure. The outer clevis of the tension strap assembly has a square shaft that fits into a square channel on the ends of the CBSS longerons. This configuration minimizes any twisting of the strap assembly during the preloading of the straps. The ends of the outer clevises are threaded and a large locking nut riding on the thread is used to apply and maintain the strap preload. Strain gauges installed on all six strap assemblies facilitated proper preloading of the thermal isolation straps. Positioning gauges were used to centre the cold bus in the CBSS. Relatively accurate positioning of the cold bus at the centre of the CBSS is important due to the limited axial and radial alignment mismatch that can be accommodated by the flex link between the cold stem and the cryocooler. The CBSS is mounted in a rotating fixture that held the CBSS prior to integration on the deck, and greatly simpli-
fied the installation of the cold bus cryo-blanket. The photograph also gives a good indication of the relative size of the HTSSE II cryogenic package.
Cryogenic
multi-layer
insulation
(MLI)
The cryogenic thermal load risk was minimized early in the HTSSE II programme by utilizing the proven HTSSE I cold bus geometry and ML1 design. The cold bus is enclosed in an aluminium sheet metal blanket mandrel approximately 230 mm in diameter by 240 mm long. It is mounted to the cold bus with thermal isolators and provides a smooth and symmetrical structure for supporting the cryoblanket. The 25-mm thick multi-layer insulation cryoblanket is constructed of approximately 40 layers of 6 pm double-aluminized Mylar with silk net separators between each layer. The layers were applied one at a time by Lockheed technicians. Each layer covers the entire external surface, including the gap between the legs of each support strap. Before applying the cold bus MLI, each leg of the support straps, the I/O cable bundle and the flex strap were individually wrapped several times with silk net to preclude any of these items from shorting out the Mylar layers within the cryo-blanket. Minimizing the number of cold bus cryo-blanket penetrations not only helped make the cryoblanket more effective, but also simplified the design and installation of the blanket. All surfaces inside the CBSS that view the cryo-blanket are covered with 15layer thermal blankets. These blankets are made of perforated 7.5 Frn double-aluminized Mylar with Dacron mesh separators. Similar blankets are used for thermal protection throughout the HTSSE II structure.
Sensors and I/O There are six cryogenic temperature sensors on HTSSE II for feedback control of the cold bus temperature and for monitoring the cold bus and the cooler performance. Two Lakeshore Cryogenics DT-470 silicon-diode temperature sensors are mounted to the cold bus and connected to the Lockheed cooler drive electronics (CDE) for closed-loop control of the cold bus temperature. Two Rosemount plati-
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num resistance thermometers (PRTs) are attached to the cold bus and connected to HTSSE II electronics for monitoring the thermal health of the cold bus. These sensors will also be used as back-up to the silicon-diode sensors for control of the cold bus temperature. Two PRTs are mounted to the flex strap, one at each end of the copper braid. The sensor wires are external to the main cryo-blanket and are routed out along the cold finger and into the wire harness. The temperature difference across the braid can be used to estimate the heat being drawn out of the cold bus by the cryocooler. This value is checked against the cryocooler performance curves to provide a health check of the cryocooler. All sensors use four-wire measurement techniques. The HTS experiments require 27 coaxial cables to handle the various RF input and output signals. To reduce the thermal losses associated with these cables, low-conductivity cables manufactured by Gore Industries are used. These stainless-steel coaxial cables are 0.5 mm in diameter and each has a thermal conductance equivalent to that of a 0.4mm diameter stainless-steel wire. The RF attenuation of these cables is 6 dB per 305 mm at 10 GHz, acceptable for HTSSE characterization testing but too large for an operational system. A total of 30 coaxial cables, including three spares, were used in building the flight cold bus. Spare cables were included because the cables are installed in an I/O subassembly and replacing failed cables would be very difficult after subassembly integration. The cables are relatively rugged and passed all environmental tests, but they are susceptible to handling damage. The weakest point on the coaxial cables is the solder joint between the conductors and the connector at the end of the cable. To protect this joint, the ends of the cables are secured with epoxy near the solder joints. This epoxy prevents loads placed on the cables from being transmitted to the solder joints. The I/O subassembly is shown in Figure 9. To save time, the I/O subassembly was built in parallel with the cold bus subassembly. The completed I/O and cold bus subassemblies are mated before installation into the cold bus support structure. In addition to the RF cables, 32 non-RF I/O wires are also needed. Quad-twist, 0.13 mm phosphor-bronze wires are used for this purpose. Eight wires are for one cryogenic experiment’s I/O and are in a stainless-steel braided shield. The other 24 wires are connected to cryogenic temperature sensors and are not shielded. All cryogenic I/O are bundled together to create a single 6-mm diameter ML1 penetration. The external end of the I/O bundle is terminated at the I/O interface plate mounted to the +X end of the CBSS. Stainless-steel jacketed 3.6 mm coaxial cables are used to connect this interface plate to the ambient RF electronics. The outboard wire harness connects to the I/O plate for nonRF connections. HTSSE II has been instrumented to monitor and diagnose cryocooler and cryogenic payload health while on orbit. The cryocooler mounting bracket has four accelerometers to provide axial and lateral vibration characterization data for the compressor and displacer. The cold bus and cold finger ends of the flex strap have platinum resistance temperature sensors. The flex strap resistance has been calibrated to a known thermal load to provide a flight heat flow meter to monitor cryocooler cooling capacity. Lockheed cryocooler drive electronics provide the compressor and displacer r.m.s. motor current. The cryogenic and non-cryogenic environments are fully characterized with temperature sensors.
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Cooler and flex strap The British Aerospace (BAe) 80 K cryocooler was selected for cryocooling the HTSSE II cold bus for three reasons: the 80 K cooler design has demonstrated the greatest reliability in space to date, it was the lowest cost option because NRL could utilize the HTSSE I qualification cooler for HTSSE II qualification testing and it offered the least schedule risk to the HTSSE II programme. The limited experience of cryocoolers in space and the single point failure influence of the cryocooler on HTSSE II was a large factor in the HTSSE programme decision to pursue a full qualification and flight unit approach to the cryogenic system. The qualification unit did pass all environmental testing successfully. The qualification unit demonstrated good design margins and reduced the risk associated with the uncertainties of cryocooler integration on HTSSE II. A large issue in the utilization of mechanical cryocoolers in space has been the effect of their mechanism-induced vibration on instruments and payloads. Unlike imaging applications or precision pointing applications, the operation of HTSSE II HTS experiments is not sensitive to the vibration environment of a single unbalanced BAe cooler. The vibration levels of a single cooler were also acceptable to the ARGOS spacecraft. The cryocooler is installed in a bracket which is bolted on thermally insulating mounts to the HTSSE II deck. A 0.26 m* radiator, mounted to the bracket, rejects heat from the cooler to space. A copper braid flex strap conductively couples the cryocooler cold finger to the cold bus. HTSSE II utilizes the three-link flex strap design as shown in Figure 10. The beryllium stem conducts heat from the cold bus through a penetration in the cryogenic ML1 to the copper collar. The copper collar has a precision bore that just fits over the beryllium stem at room temperature. It shrinks down to provide high contact pressure and low thermal interface resistance at cryogenic temperatures due to the differing coefficients of thermal expansion. This shrink fit interface is repeatable, efficient and provides an easily demated joint. The long life potential of the British Aerospace 80 K cryocooler is primarily due to the use of clearance seals instead of elastimer seals to eliminate contacting wear mechanisms that limit life. Much design priority was focused on maintaining clearance seal integrity in order to ensure long life. There are two sources of dangers to the cryocooler clearance seals which can be summarized as outside forces and temperature extremes. Outside forces include launch environment accelerations, gravity and cryogenic payload static and dynamic forces structurally transferred to the cold finger. The HTSSE II cryogenic system is launched in a warm ambient non-operating state which requires less stringent force environments during launch. Gravity-induced forces are avoided by orienting the cryocooler vertically in the HTSSE II configuration for all terrestrial testing. Launch environment accelerations were analysed and proved in qualification testing. The cryocooler cold finger is sensitive to transverse forces in both operating and non-operating states. Cryogenic payload forces on the cold finger are dependent on the flexible strap stiffness and mass characteristics. Flexibility is needed during launch to decouple a predicted ti.8 mm displacement of the cold bus relative to the cold finger. In addition, flexibility is needed to allow for the mismatch that occurs from tolerance build-up during
High Temperature
Superconducting
Space Experiment
Cable Bundle
II: T. G. Kawecki et al.
Assembly Fixture
_ (1 of 3) \
90” Coax Feed Throughs
0.5mmCoax Cables
\
\
1 6wII
v Section Inside Cryogenic MLI
/
Protective Cover / (1 of 3) Section Thermally Sunk to CSSS Figure 9
XTSSE II cryogenic
input/output
assembly
design
assembly and the thermal contraction that occurs as structures cool down. Conflicting requirements are inherent in thermal link design. Flexibility is not always compatible with the high thermal conductivity desired to operate a cold finger at the highest possible temperature relative to the payload to maximize the cooling capacity. The HTSSE II flex strap is a compromise between flexibility and conductivity. Thermal testing conducted at Lockheed determined a favourable 1.8 K W-’ thermal resistance at 77 K. Mechanical testing performed at NRL on the flex strap as a free rate of cantilever determined a transverse spring 0.14 N mm-’ which was linear in behaviour up to deflections of 5.1 mm. When adjusted for the more representative condition of moment-constrained ends, the transverse spring rate is calculated to be 0.56 N mm-’ from Roark’s beam equations, and the EI (section modulus) determined from free cantilever testing.
Stiction testing, a non-destructive test technique developed by the Thermal and Cryogenics Technology Group at the Jet Propulsion Laboratory’, can indicate when clearance seals are compromised in a cold finger and the piston is touching the inside bore. This technique was employed on the HTSSE II qualification and flight cryocoolers which were stiction mapped at 45” radial increments around the finger by adding progressively larger transverse forces to the finger until stiction occurred. Stiction occurred at forces as small as 75 f 25 g and as large as 375 f 25 g on the flight cooler depending on the radial angle of force application. Stiction testing was performed after cryocooler integration and all vibration tests. No indications of piston contact or stiction were observed with the fully integrated coolers. It is desirable to reject heat from a cryocooler at as low a temperature as possible to maximize cooler performance.
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Figure 10
Cryocooler
Superconducting
to cold bus flexible
Space Experiment
conductive
II: T.G. Kawecki et al.
strap assembly
Cryocooler ambient temperature extremes can cause piston contacting in the British Aerospace cryocooler. Stiction tests by the Jet Propulsion Laboratory on their own cooler indicated contacting of the displacer piston at temperatures below -20°C and at temperatures above +2O”C in some cases. The HTSSE II cryocooler is designed to reject at 0°C during normal operation to minimize seal risk. The compressor to displacer helium transfer tube is also thermally coupled to the cryocooler bracket to enhance cryocooler performance.
Deck 2oc
Guard 1 -57c
Guard 2 -82C
Wires 0.105
Wires 2.5
W Straps 0.056
l
Mount 8.3
) Mount 2.0 l
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Thermal model Geometric computer models of the HTSSE II on the host vehicle were constructed using the computer code TRASYS. Internal and external radiation conductors and external solar, albedo and planetary absorbed fluxes were computed for worst case hot and cold orbits and sun angles. Orbital average external radiation couplings between HTSSE II and the host vehicle’s rotating solar arrays were also computed using TRASYS. The TRASYS conductors and fluxes were combined into a SINDAFLUINT model for predicting temperatures and heat loads. All heat Daths through ML1 are modelled with parallel radiation aid conductron couplings. Figure iI shows the predicted temperatures and heat flows for the first-year hot design case as defined in Table 2.
Cold Bus -196C
l
w
MLI 0.2
MLI 0.6
I
II
I
MLI 0.190
I
I Flex Strap
Total 2.8
Total 11.0 CBSS Radiator 8.2
) Total 0.404 Cavity Radiator ”
9*
L.7
Cooler 0.404 Figure 11
Thermal
model
heat flow
predictions
(W)
High Temperature Table 2
First-year
Superconducting
hot case definition
Parameter
450 nmi. circular orbit, +Z nadir Beta = 24” (angle between the sun vector and the orbit plane) Hot environmental constants (solar, albedo, planetary) Degraded external optical properties (alpha = BOL + 0.08) Hot host vehicle (10°C) Maximum orbital average electronics waste heat (40 W)
Test history To mitigate risk, the HTSSE II project philosophy was to use proven or conservative designs where possible and test engineering development models and or qualification models as soon as possible to prove high risk elements of the design. The 20-month delivery schedule would not allow much time for recovery if a failure occurred late in the programme. During the conceptual development and project planning phase of HTSSE 11, the highest technical risk areas were considered to be tension strap fatigue strength, cryocooler integration, cryogenic thermal load, HTS experiment robustness and the cryogenic coaxial I/O strength. Limited experience of cryocoolers in space and the uniqueness of the HTSSE 11 cryogenic payload were the principal factors in the HTSSE programme’s decision to take a qualification unit and flight unit build approach to the cryogenic system. Electronic and RF boxes were considered to carry less risk and a protoflight approach was taken with them to stay within the budget. Mass simulators were substituted for electronics and RF boxes during qualification vibration testing. The test flow described below was developed and implemented as planned with only minor anomalies. Two early engineering development models (EDM) of high risk components were fabricated. The fatigue environment and the geometry-dependent fatigue performance of the HTSSE II composite straps were significant unknowns. The strap procurement was considered a long lead procurement to the HTSSE II 20-month schedule. Early testing was essential to give confidence in our structure and cryogenic thermal load estimates. A tension strap structure (the CBSS) and a worst case weighted cold bus were fabricated and random vibration tested. A worst case cold bus was used due to the high uncertainty of the final size and mix of HTS experiments, and HTSSE II hardware development was concurrent with HTS experiment development. The composite tension straps survived four random vibration tests at qualification levels without failure. An early EDM of the cryocooler cold bus flex link was also fabricated to investigate the transverse and axial spring constants. These two EDM tests allowed HTSSE II to finalize the configuration early and push toward more comprehensive full system qualification tests with confidence. Qualification testing focused on testing a complete functional cryogenic system. Full functionals can only be performed on HTSSE II when it is under high vacuum to allow HTS experiments to operate at 77 K. The test sequence was to perform thermal vacuum testing for pre-vibration functionals, followed by vibration testing, and finally thermal vacuum testing again for post-vibration functional testing. These three qualification tests took a significant period of time to complete, largely due to the seven day cooldown
Space Experiment
II: T.G. Kawecki et al.
period required for the 8.3 kg cold bus to reach 77 K. Qualification testing was desirable to demonstrate the design margin of the complex high risk cryogenic connections of the coaxial assembly and the cryocooler flex strap. Postand pre-vibration cryogenic thermal characterization also provided data indicating that the launch environment would not degrade cryogenic thermal performance. Once the HTSSE II cryogenic design was demonstrated successfully in qualification testing, the flight unit construction began. The margin demonstrated in qualification testing prevented the need for pre-vibration thermal vacuum testing of the flight unit. Environmental vibration tests were completed and then thermal vacuum testing performed to demonstrate post-vibration functionality. One failure occurred on the final flight unit. An HTS experiment on the cold bus appears to have failed in vibration. Post-failure RF reflectometry tests indicated a through path to the device and that the failure was not due to a coaxial I/O failure.
Test levels The flight level component vibration test spectrum for HTSSE II is shown in Table 3 at a 7.4 G r.m.s. level. Qualification component vibration levels are 6 dB higher and 30 s longer at an overall 14.8 G r.m.s. level. Qualification and flight unit system vibration tests were designed for acoustic input from 100 to 2000 Hz with simultaneous oneaxis random vibration input from 10 to 100 Hz to simulate a low-level structure-borne vibration input and to demonstrate workmanship. The lo- 100 Hz random vibration test is performed on each of the other axes without acoustic input. The goal of cryogenic thermal vacuum testing was to demonstrate the cryogenic margin in a worst case environment. The HTSSE requirement is to demonstrate a 30% cryocooling reserve under end of life, one-year hot case conditions. The one-year hot case was simulated by controlling the deck temperature and all the radiators to the one-year worst case temperatures as predicted by the HTSSE II thermal model.
Thermal
test results
The cryogenic thermal requirements were met during testing: to hold the cold bus at 77 + 1 K and to do it with at least a 30% reserve in cooling capacity. There were three thermal vacuum tests to verify cryogenic performance, each lasting about 2 weeks of around the clock operation: the pre- and post-vibration qualification tests and the postvibration flight acceptance thermal vacuum test. The heat load on the cryocooler, determined during these tests from Table3 trum
HTSSE II component
Frequency
acceptance
(Hz)
spec-
Level (G* Hz-‘) 0.01 +3 dB octave-’ 0.08 -3 dB octaves’ 0.01
20 20-160 160-250 250-2000 2000
Cryogenics
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High Temperature Table 4
Cryogenic
Superconducting performance
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et al.
summary
Parameter
Model predictions (W)
Qualification test 1 w
Qualification test 2 NW
Flight test w
L = heat load C = capacity of coolera R = reserve (C- L) Duty cycle (L/C) Cryogenic margin (R/L)
0.404 0.780 0.376 52% 93%
0.360 0.780 0.420 46% 117%
0.382 0.780 0.398 49% 104%
0.380 0.640 0.260 59% 68%
“At 75 K cold-finger a”
I-
temperature
and 0°C rejection temperature
CoolerElectronics
Crcpgenic
Thermometer
17.1
11.0 c
I 76.3 2 76.8 76.7’ +r*95
’
I
’ ’ 9
f
I
Time/Dare
’ I 10
’
’
’
’ ’ A$95
(days)
Figure 12 Cold bus and cryocooler versus time stability plots
electronics
temperatures
the cooler performance curves, agrees very closely with the model prediction. Table 4 summarizes the test results. The heat load was estimated from the cooler performance curves at the measured cooler rejection temperature, cold finger temperature and compressor stroke. The temperature drop across the flex strap was another estimate of the heat load. Heat loads determined by this method agreed with the cooler performance curves within 5%. The HTSSE II design goal is to have a tested duty cycle of no more than 70% on the flight cryocooler. The flight test demonstrated an equivalent duty cycle of 59%. At this duty cycle, the cold bus heat load could grow as much as 68% and the cooler could still handle the load. This number is lower than the qualification test margin because the flight cooler has a lower cooling capacity than the qualification cooler. The steady-state cold bus temperatures for the first-year hot case from the flight test are shown in Figure 12. The cooler drive electronics controlled the cold bus temperature to within + 0.2 K. This is tighter than the + 1.0 K requirement. Note that the cold bus temperature drift was affected by temperature-induced cooler drive electronics (CDE) instrumentation error. At a steady-state CDE temperature, cold bus stability was + 0.1 K. Tighter temperature control tolerances may be possible by optimizing the PID control constants in the cooler drive electronics.
HTS experiments are included in the HTSSE II payload. Unique aspects of the cryogenic system include the longterm space demonstration of two different long life Stirlingcycle cryocoolers and a cryogenic coaxial cable I/O interface. HTSSE II utilizes secondary passive cooling to reduce cryocooler cryogenic loading. The main cryogenic cold bus has a low (0.3 K) temperature gradient and is temperature controlled by the cryocooler feedback electronics to 77 f 0.1 K in the simulated one-year hot case. Under the one-year worst case thermal environment, the British Aerospace cryocooler operates at 60% of full capacity, meeting the HTSSE II margin requirement. The performance of the cryogenic design was repeatedly demonstrated in a total of three thermal vacuum tests including both the qualification and flight units. The project was completed in 24 months on a fixed cost budget basis. The cryogenic design was completed concurrently with the HTS experiment development. Planned HTSSE II integration on the ARGOS host spacecraft was in February 1996 and launch in March 1997. HTS subsystems offer potentially large benefits over conventional components but at the cost of a cryogenic system to support the low temperatures. The overhead cryosystem penalty on the spacecraft is cost, power, weight, volume, reliability and potentially vibration. The cryogenic overhead penalty must be reduced or the HTS system benefit must be very high before HTS systems are included in future operational spacecraft. Efficient cryogenic systems are essential to the utilization of HTS devices in space.
Acknowledgements Sponsorship and funding for HTSSE are provided by the US Navy’s Space and Naval Warfare Systems Command (SPAWAR). Several other US government agencies provided funding for the development of some HTS devices used in HTSSE such as DARPA, NASA and the Office of Naval Research (ONR). Several industrial groups provided HTS devices developed with their own IR&D funds.
References 1
Summary
2
HTSSE II has successfully demonstrated its design performance in qualification unit and flight unit testing. Eight
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Ross, R.G. Jr, Johnson, D.L. and Kotsubo, V.Y. BAe 80K Stirling cooler performance characterization, JPL AIRS Project Report, Pasadena, CA (May 1992) Kaweckl, T.G. High temperature superconducting space experiment II (HTSSE II): overview and preliminary cryocooler integration experience Proc 8th Int Cryocooler ConfVail, Colorado, June 1994, Plenum Publishing Corp., New York (1994)
SOOll-2275(96)00036-7
ELSEVIER
High Temperature Superconducting Space Experiment II (HTSSE II) cryogenic design T.G. Kawecki,
S.S. Chappie and D.R. Mahony*
Naval Research Laboratory, Naval Center for Space Technology, Washington, DC 20375, USA *Swales & Associates Inc., Beltsville, MD 20705, USA
Code 8223,
At 60 to 80 K large performance gains are possible from high temperature superconducting (HTS) microwave devices for communications applications. The High Temperature Superconducting Space Experiment II (HTSSE II) will demonstrate eight HTS experiments in space for up to 3 years of operation. HTSSE II is the first application of HTS technology to space. In addition to demonstrating HTS devices, an important secondary goal is to demonstrate the cryogenic technologies required for long life HTS space applications. HTSSE II utilizes a British Aerospace 80 K Stirling cycle cryocooler to refrigerate a central cryogenic bus of seven HTS experiments and has an additional stand-alone TRW HTS experiment cooled by a TRW Stirling cycle cryocooler. The HTSSE II flight unit has been assembled and has successfully passed vibration and thermal vacuum environmental tests. HTSSE II was developed on a fixed budget and a fast track schedule of 24 months and is due to launch in March 1997 on the ARGOS spacecraft. This paper presents the design and test results of the cryogenic subsystem, cryocooler integration and a cryogenic coaxial cable I/O assembly. Published by Elsevier Science Limited Keywords: high temperature wave devices
superconductivity;
The High Temperature Superconducting Space Experiment (HTSSE) programme was initiated by the US Naval Research Laboratory (NRL) to promote and demonstrate the potentially significant benefits of utilizing high temperature superconducting (HTS) materials in satellite electronic components. At 60 to 80 K large performance gains are possible in thallium and yttrium based HTS microwave devices used for communications applications. The programme goal of HTSSE is to provide a focal point for the HTS research and development community to develop and demonstrate space qualified HTS technology. HTSSE I contained 15 HTS experiments which were predominantly passive microwave circuit devices cooled to 77 K by a British Aerospace Stirling cycle cryocooler. HTSSE I was launched in 1993 and was lost before on-orbit start-up when the host spacecraft mission failed. In March 1993, HTSSE II was manifested on the Air Force Space Test Program Advanced Research and Global Observation Satellite (ARGOS) and the flight hardware design portion of HTSSE II began. HTSSE II experiments had already been under development. The primary HTSSE II mission requirement is to demonstrate HTS devices on orbit for one year. It is desired to demonstrate three years’ operation if spacecraft operations funding is available. HTS devices have evolved significantly from HTSSE I
space cryogenics;
cryocooler;
micro-
to HTSSE II by demonstrating higher device complexity and maturity in experiments that integrate HTS devices together or combine with semiconductor devices. Eleven HTS experiments were funded at the start of the HTSSE II programme. Eight experiments were selected for flight on the basis of technical merit, ability to integrate and the ability to survive the qualification test programme. HTSSE II flight microwave experiments include the following device types: channellized receiver, cueing receiver, digital instantaneous frequency measurement unit, delay lines, RF multiplexer and a digital multiplexer. In addition to demonstrating HTS devices, an important goal of HTSSE is to develop and demonstrate the critical cryogenic subsystems required to support HTS space systems. Two key cryogenic technologies required for space HTS applications were identified in the HTSSE effort: long life, high reliability mechanical refrigerators for cryocooling and low RF loss, and high thermal resistance coaxial cables to handle a microwave input and output to cryogenic HTS devices. To date, space applications of HTS devices have been largely oriented toward communications applications. Future cryogenic HTS commercial communication payloads are likely to have large volumes and significant I/O thermal parasitics. This suggests the need for cryogenic
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cooling capacities of the order of 0.5 W or greater in the 60-80 K temperature range. Communications applications typically require lifetimes of 5 years or greater. Relatively large thermal loads over long time periods preclude the practical use of solid cryogens for primary cooling. Cryogenic radiators for primary cooling are undesirable for communications applications in this temperature range because they greatly constrain orbit options and have a large impact on the overall spacecraft design. Long life, high reliability mechanical refrigerators appear to be the best cryocooling source for operational HTS payloads. These coolers offer a relatively large cooling capacity and the potential for lifetimes greater than 5 years. In concert with the HTSSE programme goal of demonstrating HTS support technologies for space, a Stirling cycle British Aerospace 80 K cryocooler is the basis of the HTSSE II cold bus cryogenic system. It maintains seven HTS experiments at 77 K on a continuous basis once on orbit. The eighth HTS experiment on HTSSE II is the TRW 2: 1 Digital Multiplexer which is mounted to a TRW Stirling cycle refrigerator operating in the 60-65 K range. At the time of the 1995 Space Cryogenics Workshop, HTSSE II had successfully completed assembly and environmental testing of the qualification unit and subsequently the flight unit. HTSSE II was then awaiting integration on the ARGOS host spacecraft in February 1996 and launch in March 1997.
Mechanical
configuration
Schedule and budget constraints were significant design drivers from the beginning. The budget had been planned in advance of the ARGOS manifest and was fixed. The schedule was determined by a short notice manifest on ARGOS. At the time of the manifest, HTSSE II was required to progress from concept to flight hardware integration on ARGOS in a 20-month period. HTSSE II HTS experiments were being developed in parallel with the rest of HTSSE II, further complicating the design effort. Due to ARGOS development delays since the start of the programme, the HTSSE II delivery date for ARGOS integration moved from September 94 to February 96. HTSSE II has used this schedule slip to reduce risk by expanding the test plan to increase the original 20 month development effort to 24 months. ARGOS is a three-axis stabilized, 2500 kg, 1 kW satellite to be launched on a Delta 2 launch vehicle in March 1997. HTSSE II is launched at ambient temperatures with only survival heater circuits enabled. HTSSE II starts cooling to cryogenic temperatures after an initial 6-week spacecraft checkout and test phase. ARGOS operates in a 450 nautical mile circular orbit at a sun synchronous inclination. HTSSE II is located on the zenith end of the spacecraft on a 1.02 m x 1.27 m honeycomb deck with volume available on both sides. The HTSSE II mass, volume and power allocations from ARGOS were generous. Next to strict budget and schedule constraints, the cryogenic considerations became the biggest design configuration driver. The heart of HTSSE II is the cryogenic subsystem which consists of a cold bus with seven HTS experiments cooled by a British Aerospace 80 K Stirling cycle cryocooler. The cold bus and experiments are supported by glass epoxy thermal isolation tension straps that are fastened to an exoskeletal cold bus support structure (CBSS) as shown in Fig-
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II: T.G. Kawecki et al. COLD BUS SUPPORT STRUCTURE
COAX i/O INTERFACE PLATE
Figure 1 (CBSS)
HTSSE
II cold
bus and cold
bus
support structure
we I. The CBSS is supported on the HTSSE II honeycomb deck with thermally insulating kinematic mounts. Figure 2 shows the ARGOS spacecraft with HTSSE II installed and a larger exploded view of the HTSSE II deck above without blankets and with some structure removed for clarity. The CBSS with radiators is located just behind the BAe 80 K cryocooler in this view. The cold bus is shown with cryogenic blankets installed. The exterior blanket dimensions are approximately 280 mm diameter by 305 mm long. Ambient temperature electronics are located on both sides of the HTSSE II deck. The TRW stand-alone experiment and cryocooler are located as shown. The overall deck layout of HTSSE II evolved into the HTS Experiment Cold Bus \
Cold Bus Structure +Z Radiator /
Cavity & +Y Radiators
Figure2 load
deck
ARGOS
host
spacecraft
with
detail
of HTSSE
II pay-
High Temperature
Superconducting
present configuration primarily to take advantage of the ARGOS sun synchronous orbit which provides secondary passive cooling to reduce cryogenic thermal parasitics. Radiators at cryogenic temperatures were not possible even in this favourable sun synchronous orbit due to the hot (60°C) +Y solar panel that rotates in a coning motion once each orbit about the Y axis. There were two additional important physical configuration drivers. One was to minimize the potential for electromechanical noise from the cryocooler on HTSSE experiments and instrumentation, and on the ARGOS spacecraft. The cryocooler was located as far as possible from the ARGOS magnetometers to avoid violating magnetic field interface control agreements. The second configuration driver was to locate the cryocooler cold finger in a vertical orientation for all ground thermal vacuum testing to minimize the chance of clearance seal problems during l-g environment operation. Clearance seals are the key to cooler longevity, so much design effort was focused on the cryocooler cold bus interface to maintain clearance seal integrity.
Thermal
design overview
The HTSSE II structure can be divided into the five temperature-controlled zones listed in Table I. The focus of this paper is on the first two zones that are directly related to cryogenic performance: the cold bus and its support structure. Two strategies were employed in the thermal design of HTSSE II: providing high resistance thermal paths between the cold bus and its external environment, and lowering the external temperatures as much as possible. The cold bus is suspended inside a triangular truss with low conductance isolation straps and is enclosed in a 25mm thick cryogenic multi-layer insulation (MLI) blanket. The RF and electrical connections for the HTS experiments on the cold bus are a major heat short through the cyro-blanket. Therefore, low conductance cables and wires are used from the cold bus to the interface plate on the +X end of the CBSS. The suspension straps, I/O cables and all the internal faces of the CBSS that view the cold bus cryo-blanket are covered with 15layer thermal blankets. The CBSS assembly is mounted to the electronics deck (zenith panel) using thermal isolators and functions as a thermal guard around the cold bus. Aluminium closure panels, attached to the -Y and +Z faces of the CBSS, act as radiation guards. The I/O interface plate is mounted to the +X face and is thermally sunk to the CBSS. The radiation guards and the I/O plate are designed to intercept heat coming through the ML1 blankets external to the CBSS. The I/O plate also intercepts heat from the external RF cabling connected to the warmer RF electronics. The interTable 1
HTSSE II
temperature
controllled
zones
Zone
Nominal temperature
Cryogenic bus (cold bus) Cold bus support structure (CBSS) BAe cryocooler TRW cryocooler Support electronics
77 K -50°C o-10°C 0-27°C 0-40x
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II: T.G. Kawecki et al.
cepted heat is rejected to space by radiators mounted to the CBSS. Three radiators are mounted to the CBSS, two for cooling the CBSS structure and one for cooling the CBSS thermal blankets that border the ‘cavity’ around the cold bus cryo-blanket (see Figure I). The cavity radiator and one CBSS radiator are mounted to the +Y face of the CBSS. The other CBSS radiator is attached to the zenith apex of the CBSS. A small extension to the zenith radiator acts as a sun shade for the +Y radiator (see Figures 3 and 4). The +Y face receives almost no direct solar flux due to the sun shade and the nature of the host vehicle’s orbital attitude. The +Y face does have a view of the hot (60°C) +Y solar panel which severely limits its ability to achieve near-cryogenic temperatures (see Figure 5). Positioning the CBSS to the -Y side of ARGOS shields the +Y side of the CBSS from most of the albedo and planetary fluxes and from some of the +Y solar panel IR flux. The CBSS thermal guard operates at about -48°C. Heat from the CBSS passes through the CBSS ML1 and into the cavity between the CBSS and the cold bus cryo-blanket. The cavity radiator positioned on the +Y side of the CBSS has a high emissive radiation coupling to this cavity and is thermally isolated from the CBSS to provide the lowest temperature possible. It is essentially a second-stage radiator. The British Aerospace 80 K Cryocooler is mounted to a bracket, independent of the CBSS and thermally isolated from the zenith panel. The cold finger of the cooler is thermally connected to the -X end of the cold bus with a flexible conductive strap. Lockheed cryocooler drive electronics control the cooler. A PID control circuit within these electronics monitors the cold bus temperature and adjusts the cryocooler compressor stroke to obtain the desired temperature. In addition to the cooler drive electronics, Lockheed Research Laboratory has been a key contributor to the evolution of the HTSSE II cryosystem. Lockheed designed and fabricated the cold bus cryogenic thermal blankets. The cryocooler flex strap assembly is an NRL/Lockheed design and was fabricated and thermally tested at Lockheed.
Cold bus design To reduce risk and meet the ambitious schedule, a cold bus mechanical design similar to the demonstrated HTSSE I design was adopted. The design goal for HTSSE II was to include as many experiments as possible within the limited volume of the cold bus. Due to the increased complexity of the HTS devices being developed for HTSSE II, an effort was made to give the experiment providers considerable flexibility in the packaging of their experiments. As a result, the seven HTS experiments selected for flight on the HTSSE II cold bus are of various shapes, sizes and masses. An exploded view of the cold bus structure on which the experiments are mounted is shown in Figure 6. The three devises and the end plates are the primary load-carrying structure of the cold bus. Axial loads are transmitted through the A286 alloy clevises and radial loads are carried through the two 9-mm thick aluminium alloy end plates. With this load-carrying configuration, the experiment mounting plates were free to take on any configuration convenient to accommodate the experiments, since the plates need only support their own weight and the weight of the
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Figure 3
Superconducting
Space Experiment
+X view of HTSSE II flight unit (external
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MLI blankets removed)
experiments. This design flexibility was important because the experiments were being developed concurrently with the cold bus structure. Careful consideration was given to the thermal performance of the cold bus. The goal was to have a structure that was nearly isothermal at the operational temperature. To ensure good conduction, 9-mm thick aluminium alloy plate was machined to form the experiment mounting plates. A thin layer of epoxy was used at all mechanical joints on the cold bus as a conformal thermal interface to maximize contact area. Temperature sensors located on opposite ends of the qualification and flight cold buses indicated AT values of the order of tenths of a kelvin. The cold bus is suspended inside the CBSS with six Sglass epoxy thermal isolation straps. The six-strap design, based on HTSSE I, minimizes ML1 penetrations to minimize the cold bus thermal parasitics. The negative impact is that rotational modes of the cold bus are not constrained and can be excited if the cold bus is not mass balanced. Excitation of cold bus rotational modes (and coupling between translation and rotation modes) can lead to significant increases in strap loading. The maximum allowable centre of gravity offset was defined (by analysis) to be +12.7 mm axially and It6.4 mm laterally. Any centre of gravity deviation beyond this would require correction with balance weights. The cold bus design did include balance
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weight mounting provisions, but careful placement of experiments on the cold bus structure avoided the need for balance weights on either the qualification or the flight cold bus. All of the RF and d.c. If0 lines on the cold bus were routed to one end of the cold bus and brought through the cryo-blanket in one 6-mm diameter bundle. To reduce the potential for cyrocooler-induced electrical noise in the cryogenic I/O d.c. lines, the cold bus I/O was located as far from the cryocooler as possible. A cylindrical beryllium cold stem, located at the opposite end of the cold bus from the I/O bundle, serves as the thermal link to the cryocooler. The end of the beryllium cold stem sticks out through a penetration in the cryo-blanket surrounding the cold bus and interfaces with a flex strap copper collar. Figure 7 is a photograph of the assembled flight cold bus before the installation of the cryo-blanket and mandrel. Four of the six thermal isolation straps and four of the seven experiments are visible in this view. The thermal link to the cryocooler, the beryllium cold stem, is visible at centre left. Semi-rigid coaxial cables that connect the I-ITS experiments to a coaxial cable interface ring are visible in the centre and at the right side of the figure. The cable bundle at centre right is all of the cold bus I/O brought together to penetrate the cryo-blanket.
High Temperature
Figure 4 -X-Y removed)
view of HTSSE II flight unit; note external
Co;ng
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RF cables and cold bus cavity (cavity radiator and external
MLI blankets
i
3 Nadir Figure 5
+X view of external thermal
environment
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Coldbus Clevis (1 of 3) \
End Plate
Experiment Mounting Plate
Figure 6
Cold bus structure,
Figure 7
HTSSE II flight cold bus prior to cryo-blanketing
Thermal
isolation
exploded
view
straps
As in the HTSSE I design, six S-glass epoxy straps (manufactured by SCI) are used to structurally support and thermally isolate the HTSSE II cold bus from the CBSS. Several strap configurations were considered; the final strap
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design selected is 1.0 mm thick, 5.3 mm wide and has a centre-to-centre length of 102 mm. The ends of the straps fit into grooves cut into stainless-steel spools. The bore of each spool contains a spherical bearing which is included to prevent torsional or bending loads on the straps. A bolt through each spherical bearing attaches the strap/spool assembly to clevises as seen in Figure 8.
High Temperature
Superconducting
Space Experiment
Spool Strap . _
/
, r’,,
II: T.G. Kawecki et al.
Cold Bus Support Structure
(1 of 6) \
‘---_
Outer Clevis
__
Strap
_J
:“I; -
__-
1
Cold Bus Clevis
Clevis ’
ItIf . . .
’ . . .
Figure 8
Cold bus tension strap assembly
design
A high-fidelity engineering model of the CBSS and cold bus was assembled early in the design process to qualify the strap design and to demonstrate fatigue performance of the straps. The fatigue performance of the first lot of straps was lower than expected, but the fatigue life capability was proved through a series of four qualification level random vibration tests. The fatigue testing was performed with the worst case cold bus mass and the worst case centre of gravity offset.
Cold bus support
structure
(CBSS)
As a significant facet of the HTSSE II thermal design, the CBSS was designed to operate at temperatures significantly below the ambient deck temperature. Approximately 0.4 m* of thermal radiator surface area cools the CBSS to reduce thermal parasitics into the cold bus. The CBSS is fastened to the HTSSE 11 deck with kinematic mounts, which prevent the deck from loading the CBSS during launch, allow the CBSS to contract during cooling and help to thermally isolate the CBSS from the warm deck temperatures. Torlon spacers are used on both sides of the kinematic mounts, and fibreglass washers are used under the heads of the mounting bolts for thermal isolation. The CBSS structure consists of three longerons which capture the outer ends of the tension strap assemblies, six angle braces separating the longerons and three ‘X’ shaped cross braces that stabilize the structure. The outer clevis of the tension strap assembly has a square shaft that fits into a square channel on the ends of the CBSS longerons. This configuration minimizes any twisting of the strap assembly during the preloading of the straps. The ends of the outer clevises are threaded and a large locking nut riding on the thread is used to apply and maintain the strap preload. Strain gauges installed on all six strap assemblies facilitated proper preloading of the thermal isolation straps. Positioning gauges were used to centre the cold bus in the CBSS. Relatively accurate positioning of the cold bus at the centre of the CBSS is important due to the limited axial and radial alignment mismatch that can be accommodated by the flex link between the cold stem and the cryocooler. The CBSS is mounted in a rotating fixture that held the CBSS prior to integration on the deck, and greatly simpli-
fied the installation of the cold bus cryo-blanket. The photograph also gives a good indication of the relative size of the HTSSE II cryogenic package.
Cryogenic
multi-layer
insulation
(MLI)
The cryogenic thermal load risk was minimized early in the HTSSE II programme by utilizing the proven HTSSE I cold bus geometry and ML1 design. The cold bus is enclosed in an aluminium sheet metal blanket mandrel approximately 230 mm in diameter by 240 mm long. It is mounted to the cold bus with thermal isolators and provides a smooth and symmetrical structure for supporting the cryoblanket. The 25-mm thick multi-layer insulation cryoblanket is constructed of approximately 40 layers of 6 pm double-aluminized Mylar with silk net separators between each layer. The layers were applied one at a time by Lockheed technicians. Each layer covers the entire external surface, including the gap between the legs of each support strap. Before applying the cold bus MLI, each leg of the support straps, the I/O cable bundle and the flex strap were individually wrapped several times with silk net to preclude any of these items from shorting out the Mylar layers within the cryo-blanket. Minimizing the number of cold bus cryo-blanket penetrations not only helped make the cryoblanket more effective, but also simplified the design and installation of the blanket. All surfaces inside the CBSS that view the cryo-blanket are covered with 15layer thermal blankets. These blankets are made of perforated 7.5 Frn double-aluminized Mylar with Dacron mesh separators. Similar blankets are used for thermal protection throughout the HTSSE II structure.
Sensors and I/O There are six cryogenic temperature sensors on HTSSE II for feedback control of the cold bus temperature and for monitoring the cold bus and the cooler performance. Two Lakeshore Cryogenics DT-470 silicon-diode temperature sensors are mounted to the cold bus and connected to the Lockheed cooler drive electronics (CDE) for closed-loop control of the cold bus temperature. Two Rosemount plati-
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num resistance thermometers (PRTs) are attached to the cold bus and connected to HTSSE II electronics for monitoring the thermal health of the cold bus. These sensors will also be used as back-up to the silicon-diode sensors for control of the cold bus temperature. Two PRTs are mounted to the flex strap, one at each end of the copper braid. The sensor wires are external to the main cryo-blanket and are routed out along the cold finger and into the wire harness. The temperature difference across the braid can be used to estimate the heat being drawn out of the cold bus by the cryocooler. This value is checked against the cryocooler performance curves to provide a health check of the cryocooler. All sensors use four-wire measurement techniques. The HTS experiments require 27 coaxial cables to handle the various RF input and output signals. To reduce the thermal losses associated with these cables, low-conductivity cables manufactured by Gore Industries are used. These stainless-steel coaxial cables are 0.5 mm in diameter and each has a thermal conductance equivalent to that of a 0.4mm diameter stainless-steel wire. The RF attenuation of these cables is 6 dB per 305 mm at 10 GHz, acceptable for HTSSE characterization testing but too large for an operational system. A total of 30 coaxial cables, including three spares, were used in building the flight cold bus. Spare cables were included because the cables are installed in an I/O subassembly and replacing failed cables would be very difficult after subassembly integration. The cables are relatively rugged and passed all environmental tests, but they are susceptible to handling damage. The weakest point on the coaxial cables is the solder joint between the conductors and the connector at the end of the cable. To protect this joint, the ends of the cables are secured with epoxy near the solder joints. This epoxy prevents loads placed on the cables from being transmitted to the solder joints. The I/O subassembly is shown in Figure 9. To save time, the I/O subassembly was built in parallel with the cold bus subassembly. The completed I/O and cold bus subassemblies are mated before installation into the cold bus support structure. In addition to the RF cables, 32 non-RF I/O wires are also needed. Quad-twist, 0.13 mm phosphor-bronze wires are used for this purpose. Eight wires are for one cryogenic experiment’s I/O and are in a stainless-steel braided shield. The other 24 wires are connected to cryogenic temperature sensors and are not shielded. All cryogenic I/O are bundled together to create a single 6-mm diameter ML1 penetration. The external end of the I/O bundle is terminated at the I/O interface plate mounted to the +X end of the CBSS. Stainless-steel jacketed 3.6 mm coaxial cables are used to connect this interface plate to the ambient RF electronics. The outboard wire harness connects to the I/O plate for nonRF connections. HTSSE II has been instrumented to monitor and diagnose cryocooler and cryogenic payload health while on orbit. The cryocooler mounting bracket has four accelerometers to provide axial and lateral vibration characterization data for the compressor and displacer. The cold bus and cold finger ends of the flex strap have platinum resistance temperature sensors. The flex strap resistance has been calibrated to a known thermal load to provide a flight heat flow meter to monitor cryocooler cooling capacity. Lockheed cryocooler drive electronics provide the compressor and displacer r.m.s. motor current. The cryogenic and non-cryogenic environments are fully characterized with temperature sensors.
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Cooler and flex strap The British Aerospace (BAe) 80 K cryocooler was selected for cryocooling the HTSSE II cold bus for three reasons: the 80 K cooler design has demonstrated the greatest reliability in space to date, it was the lowest cost option because NRL could utilize the HTSSE I qualification cooler for HTSSE II qualification testing and it offered the least schedule risk to the HTSSE II programme. The limited experience of cryocoolers in space and the single point failure influence of the cryocooler on HTSSE II was a large factor in the HTSSE programme decision to pursue a full qualification and flight unit approach to the cryogenic system. The qualification unit did pass all environmental testing successfully. The qualification unit demonstrated good design margins and reduced the risk associated with the uncertainties of cryocooler integration on HTSSE II. A large issue in the utilization of mechanical cryocoolers in space has been the effect of their mechanism-induced vibration on instruments and payloads. Unlike imaging applications or precision pointing applications, the operation of HTSSE II HTS experiments is not sensitive to the vibration environment of a single unbalanced BAe cooler. The vibration levels of a single cooler were also acceptable to the ARGOS spacecraft. The cryocooler is installed in a bracket which is bolted on thermally insulating mounts to the HTSSE II deck. A 0.26 m* radiator, mounted to the bracket, rejects heat from the cooler to space. A copper braid flex strap conductively couples the cryocooler cold finger to the cold bus. HTSSE II utilizes the three-link flex strap design as shown in Figure 10. The beryllium stem conducts heat from the cold bus through a penetration in the cryogenic ML1 to the copper collar. The copper collar has a precision bore that just fits over the beryllium stem at room temperature. It shrinks down to provide high contact pressure and low thermal interface resistance at cryogenic temperatures due to the differing coefficients of thermal expansion. This shrink fit interface is repeatable, efficient and provides an easily demated joint. The long life potential of the British Aerospace 80 K cryocooler is primarily due to the use of clearance seals instead of elastimer seals to eliminate contacting wear mechanisms that limit life. Much design priority was focused on maintaining clearance seal integrity in order to ensure long life. There are two sources of dangers to the cryocooler clearance seals which can be summarized as outside forces and temperature extremes. Outside forces include launch environment accelerations, gravity and cryogenic payload static and dynamic forces structurally transferred to the cold finger. The HTSSE II cryogenic system is launched in a warm ambient non-operating state which requires less stringent force environments during launch. Gravity-induced forces are avoided by orienting the cryocooler vertically in the HTSSE II configuration for all terrestrial testing. Launch environment accelerations were analysed and proved in qualification testing. The cryocooler cold finger is sensitive to transverse forces in both operating and non-operating states. Cryogenic payload forces on the cold finger are dependent on the flexible strap stiffness and mass characteristics. Flexibility is needed during launch to decouple a predicted ti.8 mm displacement of the cold bus relative to the cold finger. In addition, flexibility is needed to allow for the mismatch that occurs from tolerance build-up during
High Temperature
Superconducting
Space Experiment
Cable Bundle
II: T. G. Kawecki et al.
Assembly Fixture
_ (1 of 3) \
90” Coax Feed Throughs
0.5mmCoax Cables
\
\
1 6wII
v Section Inside Cryogenic MLI
/
Protective Cover / (1 of 3) Section Thermally Sunk to CSSS Figure 9
XTSSE II cryogenic
input/output
assembly
design
assembly and the thermal contraction that occurs as structures cool down. Conflicting requirements are inherent in thermal link design. Flexibility is not always compatible with the high thermal conductivity desired to operate a cold finger at the highest possible temperature relative to the payload to maximize the cooling capacity. The HTSSE II flex strap is a compromise between flexibility and conductivity. Thermal testing conducted at Lockheed determined a favourable 1.8 K W-’ thermal resistance at 77 K. Mechanical testing performed at NRL on the flex strap as a free rate of cantilever determined a transverse spring 0.14 N mm-’ which was linear in behaviour up to deflections of 5.1 mm. When adjusted for the more representative condition of moment-constrained ends, the transverse spring rate is calculated to be 0.56 N mm-’ from Roark’s beam equations, and the EI (section modulus) determined from free cantilever testing.
Stiction testing, a non-destructive test technique developed by the Thermal and Cryogenics Technology Group at the Jet Propulsion Laboratory’, can indicate when clearance seals are compromised in a cold finger and the piston is touching the inside bore. This technique was employed on the HTSSE II qualification and flight cryocoolers which were stiction mapped at 45” radial increments around the finger by adding progressively larger transverse forces to the finger until stiction occurred. Stiction occurred at forces as small as 75 f 25 g and as large as 375 f 25 g on the flight cooler depending on the radial angle of force application. Stiction testing was performed after cryocooler integration and all vibration tests. No indications of piston contact or stiction were observed with the fully integrated coolers. It is desirable to reject heat from a cryocooler at as low a temperature as possible to maximize cooler performance.
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High Temperature
Figure 10
Cryocooler
Superconducting
to cold bus flexible
Space Experiment
conductive
II: T.G. Kawecki et al.
strap assembly
Cryocooler ambient temperature extremes can cause piston contacting in the British Aerospace cryocooler. Stiction tests by the Jet Propulsion Laboratory on their own cooler indicated contacting of the displacer piston at temperatures below -20°C and at temperatures above +2O”C in some cases. The HTSSE II cryocooler is designed to reject at 0°C during normal operation to minimize seal risk. The compressor to displacer helium transfer tube is also thermally coupled to the cryocooler bracket to enhance cryocooler performance.
Deck 2oc
Guard 1 -57c
Guard 2 -82C
Wires 0.105
Wires 2.5
W Straps 0.056
l
Mount 8.3
) Mount 2.0 l
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b
W
Thermal model Geometric computer models of the HTSSE II on the host vehicle were constructed using the computer code TRASYS. Internal and external radiation conductors and external solar, albedo and planetary absorbed fluxes were computed for worst case hot and cold orbits and sun angles. Orbital average external radiation couplings between HTSSE II and the host vehicle’s rotating solar arrays were also computed using TRASYS. The TRASYS conductors and fluxes were combined into a SINDAFLUINT model for predicting temperatures and heat loads. All heat Daths through ML1 are modelled with parallel radiation aid conductron couplings. Figure iI shows the predicted temperatures and heat flows for the first-year hot design case as defined in Table 2.
Cold Bus -196C
l
w
MLI 0.2
MLI 0.6
I
II
I
MLI 0.190
I
I Flex Strap
Total 2.8
Total 11.0 CBSS Radiator 8.2
) Total 0.404 Cavity Radiator ”
9*
L.7
Cooler 0.404 Figure 11
Thermal
model
heat flow
predictions
(W)
High Temperature Table 2
First-year
Superconducting
hot case definition
Parameter
450 nmi. circular orbit, +Z nadir Beta = 24” (angle between the sun vector and the orbit plane) Hot environmental constants (solar, albedo, planetary) Degraded external optical properties (alpha = BOL + 0.08) Hot host vehicle (10°C) Maximum orbital average electronics waste heat (40 W)
Test history To mitigate risk, the HTSSE II project philosophy was to use proven or conservative designs where possible and test engineering development models and or qualification models as soon as possible to prove high risk elements of the design. The 20-month delivery schedule would not allow much time for recovery if a failure occurred late in the programme. During the conceptual development and project planning phase of HTSSE 11, the highest technical risk areas were considered to be tension strap fatigue strength, cryocooler integration, cryogenic thermal load, HTS experiment robustness and the cryogenic coaxial I/O strength. Limited experience of cryocoolers in space and the uniqueness of the HTSSE 11 cryogenic payload were the principal factors in the HTSSE programme’s decision to take a qualification unit and flight unit build approach to the cryogenic system. Electronic and RF boxes were considered to carry less risk and a protoflight approach was taken with them to stay within the budget. Mass simulators were substituted for electronics and RF boxes during qualification vibration testing. The test flow described below was developed and implemented as planned with only minor anomalies. Two early engineering development models (EDM) of high risk components were fabricated. The fatigue environment and the geometry-dependent fatigue performance of the HTSSE II composite straps were significant unknowns. The strap procurement was considered a long lead procurement to the HTSSE II 20-month schedule. Early testing was essential to give confidence in our structure and cryogenic thermal load estimates. A tension strap structure (the CBSS) and a worst case weighted cold bus were fabricated and random vibration tested. A worst case cold bus was used due to the high uncertainty of the final size and mix of HTS experiments, and HTSSE II hardware development was concurrent with HTS experiment development. The composite tension straps survived four random vibration tests at qualification levels without failure. An early EDM of the cryocooler cold bus flex link was also fabricated to investigate the transverse and axial spring constants. These two EDM tests allowed HTSSE II to finalize the configuration early and push toward more comprehensive full system qualification tests with confidence. Qualification testing focused on testing a complete functional cryogenic system. Full functionals can only be performed on HTSSE II when it is under high vacuum to allow HTS experiments to operate at 77 K. The test sequence was to perform thermal vacuum testing for pre-vibration functionals, followed by vibration testing, and finally thermal vacuum testing again for post-vibration functional testing. These three qualification tests took a significant period of time to complete, largely due to the seven day cooldown
Space Experiment
II: T.G. Kawecki et al.
period required for the 8.3 kg cold bus to reach 77 K. Qualification testing was desirable to demonstrate the design margin of the complex high risk cryogenic connections of the coaxial assembly and the cryocooler flex strap. Postand pre-vibration cryogenic thermal characterization also provided data indicating that the launch environment would not degrade cryogenic thermal performance. Once the HTSSE II cryogenic design was demonstrated successfully in qualification testing, the flight unit construction began. The margin demonstrated in qualification testing prevented the need for pre-vibration thermal vacuum testing of the flight unit. Environmental vibration tests were completed and then thermal vacuum testing performed to demonstrate post-vibration functionality. One failure occurred on the final flight unit. An HTS experiment on the cold bus appears to have failed in vibration. Post-failure RF reflectometry tests indicated a through path to the device and that the failure was not due to a coaxial I/O failure.
Test levels The flight level component vibration test spectrum for HTSSE II is shown in Table 3 at a 7.4 G r.m.s. level. Qualification component vibration levels are 6 dB higher and 30 s longer at an overall 14.8 G r.m.s. level. Qualification and flight unit system vibration tests were designed for acoustic input from 100 to 2000 Hz with simultaneous oneaxis random vibration input from 10 to 100 Hz to simulate a low-level structure-borne vibration input and to demonstrate workmanship. The lo- 100 Hz random vibration test is performed on each of the other axes without acoustic input. The goal of cryogenic thermal vacuum testing was to demonstrate the cryogenic margin in a worst case environment. The HTSSE requirement is to demonstrate a 30% cryocooling reserve under end of life, one-year hot case conditions. The one-year hot case was simulated by controlling the deck temperature and all the radiators to the one-year worst case temperatures as predicted by the HTSSE II thermal model.
Thermal
test results
The cryogenic thermal requirements were met during testing: to hold the cold bus at 77 + 1 K and to do it with at least a 30% reserve in cooling capacity. There were three thermal vacuum tests to verify cryogenic performance, each lasting about 2 weeks of around the clock operation: the pre- and post-vibration qualification tests and the postvibration flight acceptance thermal vacuum test. The heat load on the cryocooler, determined during these tests from Table3 trum
HTSSE II component
Frequency
acceptance
(Hz)
spec-
Level (G* Hz-‘) 0.01 +3 dB octave-’ 0.08 -3 dB octaves’ 0.01
20 20-160 160-250 250-2000 2000
Cryogenics
level vibration
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High Temperature Table 4
Cryogenic
Superconducting performance
Space Experiment
II: T.G. Kawecki
et al.
summary
Parameter
Model predictions (W)
Qualification test 1 w
Qualification test 2 NW
Flight test w
L = heat load C = capacity of coolera R = reserve (C- L) Duty cycle (L/C) Cryogenic margin (R/L)
0.404 0.780 0.376 52% 93%
0.360 0.780 0.420 46% 117%
0.382 0.780 0.398 49% 104%
0.380 0.640 0.260 59% 68%
“At 75 K cold-finger a”
I-
temperature
and 0°C rejection temperature
CoolerElectronics
Crcpgenic
Thermometer
17.1
11.0 c
I 76.3 2 76.8 76.7’ +r*95
’
I
’ ’ 9
f
I
Time/Dare
’ I 10
’
’
’
’ ’ A$95
(days)
Figure 12 Cold bus and cryocooler versus time stability plots
electronics
temperatures
the cooler performance curves, agrees very closely with the model prediction. Table 4 summarizes the test results. The heat load was estimated from the cooler performance curves at the measured cooler rejection temperature, cold finger temperature and compressor stroke. The temperature drop across the flex strap was another estimate of the heat load. Heat loads determined by this method agreed with the cooler performance curves within 5%. The HTSSE II design goal is to have a tested duty cycle of no more than 70% on the flight cryocooler. The flight test demonstrated an equivalent duty cycle of 59%. At this duty cycle, the cold bus heat load could grow as much as 68% and the cooler could still handle the load. This number is lower than the qualification test margin because the flight cooler has a lower cooling capacity than the qualification cooler. The steady-state cold bus temperatures for the first-year hot case from the flight test are shown in Figure 12. The cooler drive electronics controlled the cold bus temperature to within + 0.2 K. This is tighter than the + 1.0 K requirement. Note that the cold bus temperature drift was affected by temperature-induced cooler drive electronics (CDE) instrumentation error. At a steady-state CDE temperature, cold bus stability was + 0.1 K. Tighter temperature control tolerances may be possible by optimizing the PID control constants in the cooler drive electronics.
HTS experiments are included in the HTSSE II payload. Unique aspects of the cryogenic system include the longterm space demonstration of two different long life Stirlingcycle cryocoolers and a cryogenic coaxial cable I/O interface. HTSSE II utilizes secondary passive cooling to reduce cryocooler cryogenic loading. The main cryogenic cold bus has a low (0.3 K) temperature gradient and is temperature controlled by the cryocooler feedback electronics to 77 f 0.1 K in the simulated one-year hot case. Under the one-year worst case thermal environment, the British Aerospace cryocooler operates at 60% of full capacity, meeting the HTSSE II margin requirement. The performance of the cryogenic design was repeatedly demonstrated in a total of three thermal vacuum tests including both the qualification and flight units. The project was completed in 24 months on a fixed cost budget basis. The cryogenic design was completed concurrently with the HTS experiment development. Planned HTSSE II integration on the ARGOS host spacecraft was in February 1996 and launch in March 1997. HTS subsystems offer potentially large benefits over conventional components but at the cost of a cryogenic system to support the low temperatures. The overhead cryosystem penalty on the spacecraft is cost, power, weight, volume, reliability and potentially vibration. The cryogenic overhead penalty must be reduced or the HTS system benefit must be very high before HTS systems are included in future operational spacecraft. Efficient cryogenic systems are essential to the utilization of HTS devices in space.
Acknowledgements Sponsorship and funding for HTSSE are provided by the US Navy’s Space and Naval Warfare Systems Command (SPAWAR). Several other US government agencies provided funding for the development of some HTS devices used in HTSSE such as DARPA, NASA and the Office of Naval Research (ONR). Several industrial groups provided HTS devices developed with their own IR&D funds.
References 1
Summary
2
HTSSE II has successfully demonstrated its design performance in qualification unit and flight unit testing. Eight
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Ross, R.G. Jr, Johnson, D.L. and Kotsubo, V.Y. BAe 80K Stirling cooler performance characterization, JPL AIRS Project Report, Pasadena, CA (May 1992) Kaweckl, T.G. High temperature superconducting space experiment II (HTSSE II): overview and preliminary cryocooler integration experience Proc 8th Int Cryocooler ConfVail, Colorado, June 1994, Plenum Publishing Corp., New York (1994)