The soft x-ray spectrometer (SXS) onboard ASTRO-H (named Hitomi after launch) is a microcalorimeter-type spectrometer, installed in a dewar to be cooled at 50 mK. The energy resolution of the SXS engineering model suffered from microvibration from cryocoolers mounted on the dewar. This is mitigated for the flight model (FM) by introducing vibration isolation systems between the cryocoolers and the dewar. The detector performance of the FM was verified before launch of the spacecraft in both ambient condition and thermal-vacuum condition, showing no detectable degradation in energy resolution. The in-orbit detector spectral performance and cryocooler cooling performance were also consistent with that on ground, indicating that the cryocoolers were not damaged by launch environment. The design and performance of the vibration isolation system along with the mechanism of how the microvibration could degrade the cryogenic detector is shown. Lessons learned from the development to mitigate unexpected issues are also described.
Soft X-ray Spectrometer (SXS) onboard ASTRO-H (named Hitomi after launch) is a microcalorimeter-type spectrometer, installed in a dewar to be cooled at 50 mK. The energy resolution of the SXS engineering model suffered from micro-vibration from cryocoolers mounted on the dewar. This is mitigated for the flight model by introducing vibration isolation systems between the cryocoolers and the dewar. The detector performance of the flight model was verified before launch of the spacecraft in both ambient condition and thermal-vac condition, showing no detectable degradation in energy resolution. The in-orbit performance was also consistent with that on ground, indicating that the cryocoolers were not damaged by launch environment. The design and performance of the vibration isolation system along with the mechanism of how the micro-vibration could degrade the cryogenic detector is shown.
Spacecraft are subjected to shock loads in the several thousands of g's level during their trip to orbit. These high shock loads usually result from some separation event, such as staging, spacecraft separation, and fairing separation. Shock loads are very detrimental to spacecraft components, instruments and electronics. A new type of shock isolation system is discussed. This shock system, referred to as the SoftRide ShockRing, is a whole-spacecraft isolation system, i.e., it shock isolates the complete spacecraft from the launch vehicle. Seven whole-spacecraft vibration isolation systems (SoftRide) have flown to date and flight data confirms large reductions of the dynamic loads on the spacecraft. The standard SoftRide system is a lower frequency isolation system than the ShockRing, vibration isolating the spacecraft starting in the approximately 25 Hz range. The ShockRing is targeted at shock loads and is set to isolate above approximately 75 Hz. Component tests have been performed on the ShockRing using a specially built pneumatic gun that can generate 10,000 g's on the test article. Results from these tests demonstrate substantial reductions of the shock being transmitted to the payload. Results from a system test consisting of a spacecraft simulator, payload attachment fittings, avionics section, and shock plate are discussed. In the system tests, pyrotechnic devices were used to obtain the high levels of shock for the tests.
Small launch vehicles present an economically viable method for placing small satellites into orbit. These launch vehicles would be even more attractive to satellite customers if they could provide a softer ride to orbit. Passive whole-spacecraft vibration isolation systems have been developed for small launch vehicles to greatly reduce the dynamic launch loads. To date, two types of isolation systems have been designed.
Small launch vehicles historically provide a very rough ride to spacecraft during launch. This is particularly true of solid-fueled launch vehicles. In order for the spacecraft to survive such a trip to orbit, one of two choices must be made: (1) design all structure, payloads, and systems on the spacecraft to be strong enough to survive the high launch loads, or (2) reduce the magnitude of the high launch loads. The former is not a good choice because it typically requires additional cost, schedule, and weight. The latter is the preferred choice because it allows the focus of the spacecraft design to be primarily for on-orbit performance rather that for launch survival. Under a number of contrasts from the Air Force Research Laboratory, Space Vehicles Directorate, whole- spacecraft vibration isolation systems have been in development since 1993. This work has resulted in two whole- spacecraft isolation systems (SoftRide) that have been flown on Taurus launch vehicles, the first in February 1998 with the GFO spacecraft and the second in October 1998 with the STEX spacecraft. Both of these isolation systems were designed primarily to reduce axial dynamic responses on the spacecraft due to resonant burn excitations from the motors of the solid- fueled booster. Full coupled-loads analyses were used to predict the performance of the SoftRide systems. Using the isolation requirements derived from these analyses, hardware having the correct damping and stiffness was designed to implement the isolation system. All isolation system components were extensively tested and characterized. Typical results show 85% attenuation (i.e., only 15% of original) for the worst case resonant burn condition and 59% attenuation for a combination of static plus worst case resonant burn condition in the axial spacecraft c.g. location. No detrimental effects from the SoftRide system were observed. Limited flight data from the two flights agree with the predictions. SoftRide systems are now under development for the first and second OSP launches and for the Taurus/MTI launch. Additionally, isolation systems are being designed for larger liquid-fueled launch vehicles. This isolation system technology will greatly further the goal of better, faster, cheaper, and lighter spacecraft.
ESPA, the Secondary Payload Adapter for Evolved Expendable Launch Vehicles, addresses two of the major problems currently facing the launch industry: the vibration environment of launch vehicles, and the high cost of putting satellites into orbit. (1) During the 1990s, billions of dollars have been lost due to satellite malfunctions, resulting in total or partial mission failure, which can be directly attributed to vibration loads experienced by payloads during launch. Flight data from several recent launches have shown that whole- spacecraft launch isolation is an excellent solution to this problem. (2) Despite growing worldwide interest in small satellites, launch costs continue to hinder the full exploitation of small satellite technology. Many small satellite users are faced with shrinking budgets, limiting the scope of what can be considered an 'affordable' launch opportunity.
A whole-spacecraft isolation system for the GFO/Taurus mission was designed, fabricated, tested, and subsequently flown on February 10, 1998. This isolation system was designed to reduce dynamic responses on the GFO spacecraft caused by the resonant burn dynamic load introduced by the Castor 120 solid rocket motor. Longitudinal (flight direction) response of the GFO spacecraft center of gravity, due to the resonant burn load, was reduced by a factor of seven. The isolation system design was very nonintrusive to existing hardware, lightweight, and effective. Flight data indicates that the isolation system performed as designed. The GFO spacecraft had a successful launch and is currently operational on-orbit. A second flight of this type of isolation system occurred in October 1998. Similar isolation systems are planned for other flights in 1999 and 2000. This whole-spacecraft isolation technology was highly successful for the GFO/Taurus mission.
A U.S. Air Force-sponsored team consisting of Boeing (formerly McDonnell Douglas), Honeywell Satellite Systems, and CSA Engineering has developed technology to reduce the vibration felt by an isolated payload during launch. Spacecraft designers indicate that a launch vibration isolation system (LVIS) could provide significant cost benefits in payload design, testing, launch, and lifetime. This paper contains developments occurring since those reported previously. Simulations, which included models of a 6,500 pound spacecraft, an isolating payload attach fitting (PAF) to replace an existing PAF, and the Boeing Delta II launch vehicle, were used to generate PAF performance requirements for the desired levels of attenuation. Hardware was designed to meet the requirements. The isolating PAF concept replaces portions of a conventional metallic fitting with hydraulic- pneumatic struts featuring a unique hydraulic cross-link feature that stiffens under rotation to meet rocking restrictions. The pneumatics provide low-stiffness longitudinal support. Two demonstration isolating PAF struts were designed, fabricated and tested to determine their stiffness and damping characteristics and to verify the performance of the hydraulic crosslink concept. Measurements matched analytical predictions closely. An active closed-loop control system was simulated to assess its potential isolation performance. A factor of 100 performance increase over the passive case was achieved with minor weight addition and minimal power consumption.
One of the most difficult tasks in the structural control industry is providing linear, predictable, passive damping over a wide frequency range. This challenge has been worked around successfully in the past, but rarely has it been performed ideally. The subject matter of this paper takes a radical step toward attaining the goal of linear damping performance, while adding very low static stiffness to the system being damped.
A spacecraft is subjected to very large dynamic forces from its launch vehicle during its ascent into orbit. These large forces place stringent design requirements on the spacecraft and its components to assure that the trip to orbit will be survived. The severe launch environment accounts for much of the expense of designing, qualifying, and testing satellite components. Reduction of launch loads would allow more sensitive equipment to be included in missions, reduce risk of equipment or component failure, and possibly allow the mass of the spacecraft bus to be reduced. These benefits apply to military as well as commercial satellites. This paper reports the design and testing of a prototype whole-spacecraft isolation system which will replace current payload attach fittings, is passive-only in nature, and provides lateral isolation to a spacecraft which is mounted on it. This isolation system is being designed for a medium launch vehicle and a 6500 lb spacecraft, but the isolation technology is applicable to practically all launch vehicles and spacecraft, small and large. The feasibility of such a system on a small launch vehicle has been demonstrated with a system-level analysis which shows great improvements. The isolator significantly reduces the launch loads seen by the spacecraft. Follow-on contracts will produce isolating payload attach fittings for commercial and government launches.
Experiments performed aboard the space shuttle often utilize sensitive scientific equipment which cannot withstand high launch loads without damage. It would be highly advantageous to reduce the severity of the dynamic launch environment so that space-qualification of such equipment would be faster and less expensive. This is the goal of the Soft Ride to Orbit program. This program has identified passive damping as one technology which will reduce loads seen by equipment and thereby provide a softer ride. Finite element structural modeling was used to predict both undamped and damped responses to simulated launch loads. The modal strain energy method was used to design the passive damping treatment. This treatment was manufactured and applied to the drawer. All analyses were verified by modal and vibration testing. It was found that predicted and tested frequencies, damping values, and vibration response levels agreed reasonably well, thus showing that passive damping may be designed into future equipment drawers to reduce launch loads on sensitive equipment.
Instruments and machines requiring very high stability should be isolated from their normally less stable environment. Exact constraint mounting using six, single-constraint flexures provides a stiff connection between the instrument and its environment while isolating the instrument from low frequency deformations of the environment, such as thermal expansion. Higher frequency disturbances, however, transfer through the flexures and excite vibration modes of the instrument. Traditionally, passive or active vibration isolation is employed to attenuate environmental disturbances reaching the instrument. However, strict alignment requirements for the instrument preclude the use of low-frequency isolation, unless active methods are used. Therefore, the solution is to provide damping in parallel with the flexures to reduce the vibration amplitudes of the instrument. Flexures concentrate strain energy in blades making them excellent candidates for damping treatments. A properly designed damping treatment across the flexures can provide as much as 8% to 10% viscous damping to the isolation modes and will also help attenuate the instrument vibration modes. Thus, through the use of six damped single-constraint flexures the instrument's requirements for stability, alignment, stress, and vibration may be met. An application of this approach will be employed on the Reflection Grating Array (RGA) for the X-ray Multi-mirror Mission for the European Space Agency. The RGA is an array of 200 diffraction gratings aligned to sub-micron and sub-arc-second tolerances relative to each other. This produces a coherent wavefront for spectrum analysis. The launch vehicle will be an Ariane 5 scheduled for 1998.