Current LISA Spacecraft Design S. M. Merkowitz, K. E. Castellucci, S. V. Depalo, J. A. Generie, P. G. Maghami, and H. L. Peabody NASA Goddard Space Flight Center, Greenbelt, MD 2077 1. USA E-mail: [email protected] Abstract. The Laser Interferometer Space Antenna (LISA) mission. a space based gravitational wave detector. uses laser metrology to measure distance fluctuations between proof maases aboard three spacecraft. LISA is unique from a mission design perspective in that the three spacecraft and their associated operations form one distributed science instrument. unlike more conventional missions where an instrument is a component of an individual spacecraft. The design of the LISA spacecraft is also tightly coupled to the design and requirements of the scientific payload; for this reason it is often referred to as a "sciencccraft." Here \ve describe some of the unique features of the LISA spacecraft design that help create the quiet environment necessary for gravitational wave observations. 1. Introduction The Laser Interferometer Space Antenna (LISA) mission will be sensitive to gravitational waves from astrophysical sources such as black hole. white dwarf, and neutron star binaries. LISA is con~prisedo f three identical spacecraft separated by 5 million kilometers nominally forming an equilateral triangle. Each spacecraft (SIC) encompasses two freely floating proof masses that are the references for the gravitational wave measurement. Each leg of the triangle acts as a single arm of an interferometer that is used to measure any change in the distance between the distant proof masses. The current design of LISA requires that all spurious accelerations of the proof masses be reduced to a level below 3x 1 0-15(1+ (f/8m~z)~+(0.lm~z!f)')n'~ /s'It!~zw ithin the measurement band~vid th. A detailed noise budget was dekeloped that set the required level of all known disturbance effects [1.2]. There are two types of contributors to this noise: effects that act directly on a proof mass (such as magnetic field fluctuations. voltage and charge effects, and thermally driven gas and radiation effects). and effects that couple SIC motion (caused by tluctuatiolls in the forces acting on the spacecmft. such as thruster noise and variations in the solar photon press~~reto) a proof mass through spatial gradients of the static force fields (gravitational. electric, etc.). The clesigii of the LISA SIC is driven by the need to keep the proof masses as u~ldisturbeda s possible by providing a quiet and stable enviro~~menAt.1 1 the elements used for this purpose are grouped together into \\hat is called the Disturbance Reduction System IDRS). The DRS consists of a set of sensors. actuators. and the cotltrol laws that bind them together to meet the disturbance and pointing requireine~~tdsu ring LISA science operations. It also includes design features. such as thrnnnl and magnetic shields. that are implemented to keep disturbances from reaching the proof masses. 111 this paper, we discuss how the c~ure~dlets ign of the LISA SIC s~~pportthse req~~irenienotsf the DRS by creating the necessar3 environment for the gravitational wave observations. 2. hlechanical Design The primary SIC structure consists of honeycomb panel upper and lower decks with aluminum alloy exterior sidewalls and a thermally isolated honeycomb panel solar array deck. Each of the three SIC will be nested inside of a propulsion module with the three propulsion modules stacked into a column inside the launch vehicle payload fairing. The stack design was chosen to carry the launch loads through the propulsion modules' outer shells. thereby isolating the SIC from the direct launch loads. A separation system, located on the bottom deck will serve to jettison the Sic from the propulsion module once the SIC has reached its operational orbit. Figure 1. Top view of the LISA Spacecraft. The SI:C bus structure is designed seamlessly with the scientific payload. The optical assemblies are contained within thermal shrouds and incorporate the use of flex-pivots for rotational motion. The electronic boxes are mounted directly onto the bottom deck panel and the exterior wall. The current design philosophy to isolate the payload and bus electronics from solar heat input and to reject extraneous electrical waste heat into space means that mounting of any electronic components to the top deck panel is avoided. In addition to housing the payload, the SIC bus structure must also provide mounting accommodations for the high gain and omni antennas. sun sensors. star trackers, and micronewton thrusters on the outside of the main structure as seen in fig~iresI and 9. All initial integration activities Ivill occur with the aolar array and top deck components removed. including the panels themselves. The solar array and top deck components will be installed in the last steps of assembly. In order to accolnmodate removal and replacement of components. rnairltenance or repairs after the SIC is completely assembled. access panels are provided at six locations around the circumference of thc bus structure. Figure 2. Side view of the LISA Spacecraft 3. Thermal Design The LISA thermal design benefits greatly from the stable thermal environment provided by the operational orbits, as the SIC operate far from the earth where eclipse. albedo, and planetary heating effects are negligible. The orientations of the SIC are held at a constant 30" angle between the SIC normal and the solar vector, providing for a stable solar heat input. All electrical components are power stabilized to provide constant power dissipations. with switching operations minimized, particularly \+,ithint he LISA measurement bandwidth. The shape and design of the SIC also provides for further thermal stability (31. The LISA SIC thermal design has two primary functions: protect the payload and mission critical components from the harsh thermal environ~nento f space throughout the mission. and maintain thermal stability within the LISA measurement bandwidth during the science phase of the mission. Surbival of the instrumentation during the cruise phase is achieved through the use of heaters and multi-layer-insulation on the SIC and propulsion module. Active thermal control is applied using thermostats and thermistors located inside the electronics boxes during the cruise phase. During the science phase. the survival heaters are turned off, being supplemented by electronics box waste heat. Current thermal analysis indicates that the required thermal stability during the science phase can be achieved without the use of active temperature control through the use of successive thermal shields ["i The first layer of' thermal isolation is provided by the solar array deck. The deck will be made of alurninum or composite honeycomb. The sun facing facesheet of the deck contains either gold or vapor deposited aluminum coated optical surface reflectors and solar cells. The solar array deck is isolated from the bus structure sing low conductivity flexures. The top surface of the top deck panel is gold coated to minimize absorption of any radiation from the bottom surface of the solar array deck. A high emissivity coating around the bottom outer edge of the solar array will facilitate heat rejection to space. This isolation results in less than 1% of incident solar energy being transferred to the structure. A second layer of isolation is provided by the payload shield. which is gold coated on the inside and outside surfaces in order to niinimize the absorptioli of any radiated energy (e.g. electronic box radiated heat). Low conducticity standoffs are also placed betjveen the payload shield and the bottom plat< to i~dititt,h e c~)~lclilclipuiit~th . The last layer of payload thermal isolatio~is~ probidcd by an internal shield. which is ;~lso gold coated on both sides and i~lcludesl ow conductivity mounts. This helps to shield the optical bench from any disturbances or gradients in the payload shield. Finally. the cylindrical exterior wall and bottom deck of the SIC function as radiators to reject electronics waste heat. The area and coatings of these surfaces are optimized to keep the temperature inside the SIC within the operating range of the electronics without the need for additional heaters. 4. Self-Gravity The gravitational field at the proof masses must be kept as stnall, flat, and as stable as possible. The static gravitational field at the proof masses must be kept below Sx 10-"' m!s2 along the measurement axes. In meeting this requirement, the amount of static force that must be compensated for by the gravitational reference sensor electrodes will be minimized. This is important for two reasons: first, the force fluctuations generated by the applied compensating electric field are proportional to the total force applied by that field. Fields in both the measurement axes and the other degrees of freedom are important as some of the force fluctuations from other directions will leak into the measurement axis through cross-couplings. Second. the compensating electric field creates a virtual spring between the Sic and the proof mass. The residual motion of the SIC will couple through this stiffness causing an acceleration disturbance to the proof mass [2]. The gradient of the gravitational field at the proof mass locations must be kept below 3x10-' s-'. The gravity gradient creates a virtual apring between the Sic and the proof mass. The residual motion of the SIC will couple through this stiffness causing an acceleration disturbance to the proof mass. Finally. fluctuating distortions of the S/C will change the gravitational field. These distortions must be minimized such that their acceleration disturbance to the proof masses is kept below 5x10-'" m/s2i'd~z. These self-gravity requirements are met through careful mass balancing of the spacecraft components and the minimization of moving parts. The use of items such as deployed solar arrays are ruled out due to both their uncertain positioning and the gravitational field fluctuations due to thermal distortions. The SIC structure is also designed to minimize thermal distortions and excitable vibrational modes. The self-gravity requirements flow down to set requirements on the knowledge of the mass properties and placement of all hardware in the LISA S!C. The SIC can be divided into zones where all items within a zone have the same knowledge requirements. This does not set the accuracy needed in producing each part; rather it sets the accuracy needed in measuring and identifying the part after it is manufactured. In other words. it defines how well the part must be weighed, measured. and placed within the Sic. A detailed discussion of self-gravity analysis and balancing can be found in references [S-71. 5. Magnetic Environment The magnetic properties of the proof mass and characteristics of the magnetic field contribute to several effects throughout the DRS error budget. The leading effects are: the interaction between the fluctuating interplanetary magnetic field with the SIC magnetic gradient. fluctuations from the magnetic gradient induced current dissipations (Eddy current damping), and a fluctuating Sic magnetic gradient [3].T he current budget value for the magnetic gradient at the proof mass is 5x10." T;m. Magnetic gradient fluctuations must be kept below 2.5x10-' T/m:\/kz at 0.1 mHz at the proof Inass locations. The magnetic field inside the spacecraft is driven by the magnetically hot components. Magnetic parts are generally avoided in the LISA design. however it is not cost effective to completely eliminate thein all. .411 magnetic itenis are tracked so that a masnetic budget can be maintained. The budget magnetic field should be easily met by placing all the magnetic co~iipo~~efnart sf rorn the proof masses and including a modest amc.un? of n~agnetics hielding ax! ~~~l?pei?sati~n. A number of design rules arc available to minimize magnetic fieliis. Thc obvious solution is to minimize the use of magnetic material. Unavoidable permanent magnets can be compensated for with oppositely oriented duplicates such as a cold spare. The remaining permanent magnets can be shielded using high permeability foil magnetic materials such as METGLAS or VITROVAC. Twisted pair is used for all wiring. Current loops in the power system is eliminated by using a star or single point grounding. No primary supply currents are allowed to flow through the SIC structure. The solar array can be a very significant source of stray magnetic fields due to the large currents. however, its linear geometry makes it straightforward to cancel out by correct placement of forward and return interconnections sing a technique called "backwiring." In backwiring the return wire from each string of solar cell modules is returned directly underneath the modules in that particular string and carefully routed along a line just behind the centerline of the modules. Each string and module of the string is self-canceling and does not depend on the magnetic field of an adjacent module or string for cancellation. If a inod~~flaei ls in flight the current in both the string and the return drop to zero simirltaneously leaving no uncompensated currents in the array. 6. Attitude and Drag Free Control The LISA attitude and drag-free control system is responsible for ensuring that the residual acceleration of the two proof masses fall below the LISA sensitivity requirements by providing tight pointing and translational control of the LISA spacecraft and its proof masses. Stringent requirements are placed on the rotational (8 nradIGHz) and translational dynamics (1.5~(1+(f/8mHz)~)n'"m id~zo)f each spacecraft to ensure that the proper sensitivity for science measurements can be achieved. This is implemented by performing a number of control functions [8,9]. Along the two sensitive axes (telescopes' lines of sight), the SIC is controlled around the proof- masses such that all residual accelerations in the sensitive axes are minimized. In addition, the SIC follows the average out-of-plane motion of the proof masses in order to compensate for the solar radiation pressure. The control error signals are obtained from the optical and electrostatic readouts of the proof masses. The very sensitive optical proof mass metrology is used on the two sensitive axes in order to reduce the sensor noise in the generated force signals to the micronewton propulsion system. The drag-free control covers all 3 translational degrees of freedom of the S/C. During science mode, the SIC attitude control is performed by feeding back the information from inertial wavefront sensing of the incoming laser light to the micronewton propulsion system. Since two telescopes are on-board and mounted at an angle of nominally 60 deg with respect to each other, inertial wavefront sensing provides a total of 4 tip and tilt error angles. By applying the corresponding geometric relations. these angles can be used to determine the complete SIC attitude error to align the telescopes with respect to the incoming laser beams as well as the angular error between the two telescopes. Due to orbital mechanics, the anglc between the two interferometer arms is constantly changing. therefore. this angle must be controlled as well. The information from the inertial wavefront sensing is used again to determine the angular error between the two telescopes. The measured angular error is used to generate a feedback signal for the optical assembly actuators. Note that one of the two optical assemblies will remain in a fixed position lvhile the second one will be constantly actuatcd at 10 Hz (either assembly can be actuated, providing redundancy). The constellation of three S!C rotates about the constellation center on a once per year period. Consequently. the three SIC are always moving relative to each other. Because of this motion and the 5 million km separation bet~veenS ic, the outgoing beam nlust be pointed ahead of the incoming beam in order to be directed towards where the distant SIC will be when the light arrives. This is accomplished by a point-ahead mirror that can be actuated. The mirror is actuated out-of-plane only, since that component of the point-ahead angle exhibits large variations throughout the year. The in- pl~uiec omponent shows a small \ariation about a fixed bias. which niay be iiccornmodated by a prc- fixed tilt of the ~r~i[~Tohre. mirror pointing can be pertormed by a simple lookup table as the angles itre slo\vly varying and are sufficiently predictable. LISA requires micronewton thrusters to provide the fine spacecraft attitude and position control for drag free flight and beam pointing to the distant spacecrafts. The thrusters are operated continuouslq during scicnce operations with their thrust le\els set by the disturbance reduction system control loops. Three different thruster technologies are nearing flight readiness for LISA Pathfinder and have demonstrated the LISA thrust and thrust noise requirements: the colloid micronewton thruster (CMNT) made by Busek Co. in Boston [lo]. the indium needle ficld emission electric propulsion (In- FEEP) thruster made by ARC Seibersdorf in Austria [l I], and cesium slit FEEP (Cs-FEEP) made by XLTA S .p.X. in Italy [ 121. At least six micronewton thrustcrs on each spacccraft must be operating continually during science operations for the entire LISA mission. Enough consumables are carried for thc entire extended mission. With sunlight photon pressure as the largest external force acting on the spacecraft. the micronewton thrusters must produce on the order of 10 yN of thrust with better than a 0.1 ,uh' rcsolution during science measurements. Furthermore, over the LISA science measurement bandwidth, thrust and thrust noise must be stable and within the error limitations of the DRS over the entire mission, < 0.1 pNIu'Hz (open loop) at the high end of the measurement band. Brief periods of higher thrust, >30 jrN, may be required during tip-off recovery. constellation acquisition. and safe- mode operations. 7. Summary The LISA SIC and scientific payload are tightly coupled into one seamless design in order to optimize the sensitivity to gravitational waves. While some of the LISA requirements are unique among spacecraft design (such as self gravity), the design principles and methodologies are familiar. Modeling and analysis of the current LISA SIC design indicates that these requirements can be met through standard engineering practices without the need for new technologies. For example, the self- gravity requircments can be met through careful tracking of the mass of all components. The level of precision may be more than a traditional space mission, but not beyond standard metrological tcchniques. With careful design and strong systems engineering. the LISA SIC will be ready to support the exciting astrophysical measurements from the first space based gravitational wave observations. References R. T. Stebbins cr 01.. Class. Quant. Grav. 21, S653 (2004). S. M. Merkowitz. in Proceedings of the 6th International LISA Symposiuln, edited by S. M. Merkowitz and J. C. Livas (American Institute of Physics Conference Proceedings, New York. 2006). H. Peabody and S. ?vlerko\vitz, Class. Quant. Grav. 22, S403 (2005). H. Peabody and S. M. hlerko\vitz, in Proceedings of the 6th International LISA S>nlposium. edited by S. M. Merkocvitz and J. C. Livas (American Institute of Physics Conference Proceedings. New York, 2006). A. M. Gopstein et a/., in Proceedings of the 6th International LISA Syniposiu~n,e dited by S. M. Merkowitz and J. C. 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