The Segmented Mirror Telescope (SMT) at the Naval Postgraduate School (NPS) in Monterey is a next-generation
deployable telescope, featuring a 3-meter 6-segment primary mirror and advanced wavefront sensing and correction
capabilities. In its stowed configuration, the SMT primary mirror segments collapse into a small volume; once on
location, these segments open to the full 3-meter diameter. The segments must be very accurately aligned after
deployment and the segment surfaces are actively controlled using numerous small, embedded actuators.
The SMT employs a passive damping system to complement the actuators and mitigate the effects of low-frequency
(<40 Hz) vibration modes of the primary mirror segments. Each of the six segments has three or more modes in this
bandwidth, and resonant vibration excited by acoustics or small disturbances on the structure can result in phase
mismatches between adjacent segments thereby degrading image quality. The damping system consists of two tuned
mass dampers (TMDs) for each of the mirror segments. An adjustable TMD with passive magnetic damping was
selected to minimize sensitivity to changes in temperature; both frequency and damping characteristics can be tuned for
optimal vibration mitigation.
Modal testing was performed with a laser vibrometry system to characterize the SMT segments with and without the
TMDs. Objectives of this test were to determine operating deflection shapes of the mirror and to quantify segment edge
displacements; relative alignment of λ/4 or better was desired. The TMDs attenuated the vibration amplitudes by 80%
and reduced adjacent segment phase mismatches to acceptable levels.
The Gemini Planet Imager (GPi) is an instrument that will mount to either of two nominally identical Telescopes,
Gemini North in Hawaii and Gemini South in Chile, to perform direct imaging and spectroscopy of extra-solar planets.
This 2,000-kg instrument has stringent mass, center-of-gravity, flexure, and power constraints. The Flexure Sensitive
Structure (FSS) supports the main opto-mechanical sub-systems of the GPi which work in series to process and analyse
the telescope optical beam.
The opto-mechanical sub-systems within the FSS are sensitive to mechanical vibrations, and passive damping strategies
were considered to mitigate image jitter. Based on analysis with the system finite element model (FEM) of the GPi, an
array of 1-kg tuned mass dampers (TMDs) was identified as an efficient approach to damp the first two FSS flexural
modes which are the main sources of jitter. It is estimated that 5% of critical damping can be added to each of these
modes with the addition of 23 kg of TMD mass. This estimate is based on installing TMD units on the FSS structural
members. TMD mass can be reduced by nearly 50% if the units can be installed on the opto-mechanical sub-systems
within the FSS with the highest modal displacements.
This paper describes the structural design and vibration response of the FSS, modal test results, and plans for
implementation of the TMDs. Modal measurements of the FSS structure were made to validate the FEM and to assess
the viability of TMDs for reducing jitter. The test configuration differed from the operational one because some
payloads were not present and the structure was mounted to a flexible base. However, this test was valuable for
understanding the primary modes that will be addressed with the TMDs and measuring the effective mass of these
modes.
This paper describes evaluation of an autonomous-material system tailored for free-layer vibration damping of structural
elements. The magnetostrictive particulate composite (MPC) material described has moderate stiffness and minimal
temperature and frequency dependence. The composite is created by curing Terfenol particles {Tb(1-x)Dy(x)Fe(2),0.2
Delivery of Orbital Replacement Units (ORUs) to the International Space Station (ISS) and other on-orbit destinations is an important component of the space program. ORUs are integrated on orbit with space assets to maintain and upgrade functionality. For ORUs comprised of sensitive equipment, the dynamic launch environment drives design and testing requirements, and high frequency random vibrations are generally the cause for failure. Vibration isolation can mitigate the structure-borne vibration environment during launch, and hardware has been developed that can provide a reduced environment for current and future launch environments.
Random vibration testing of one ORU to equivalent Space Shuttle launch levels revealed that its qualification and acceptance requirements were exceeded. An isolation system was designed to mitigate the structure-borne launch vibration environment. To protect this ORU, the random vibration levels at 50 Hz must be attenuated by a factor of two and those at higher frequencies even more. Design load factors for Shuttle launch are high, so a metallic load path is needed to maintain strength margins. Isolation system design was performed using a finite element model of the ORU on its carrier with representative disturbance inputs. Iterations on the model led to an optimized design based on flight-proven SoftRide MultiFlex isolators. Component testing has been performed on prototype isolators to validate analytical predictions.
KEYWORDS: Temperature metrology, Telescopes, Titanium, Inspection, Space telescopes, Hubble Space Telescope, Accelerated life testing, Ultrasonics, Aerospace engineering, Adhesives
During the March 2002 Servicing Mission by Space Shuttle (STS 109), the Hubble Space Telescope was refurbished with two new solar arrays that now provide all of its power. These arrays were built with viscoelastic/titanium dampers, integral to the supporting masts, which reduce the interaction of the wing bending modes with the Telescope. Damping of over 3% of critical was achieved. To assess the damper's ability to maintain nominal performance over the 10-year on-orbit design goal, material specimens were subjected to an accelerated life test. The test matrix consisted of scheduled events to expose the specimens to pre-determined combinations of temperatures, frequencies, displacement levels, and numbers of cycles. These exposure events were designed to replicate the life environment of the damper from fabrication through testing to launch and life on-orbit. To determine whether material degradation occurred during the exposure sequence, material performance was evaluated before and after the accelerated aging with complex stiffness measurements. Based on comparison of pre- and post-life-cycle measurements, the material is expected to maintain nominal performance through end of life on-orbit. Recent telemetry from the Telescope indicates that the dampers are performing flawlessly.
The Stratospheric Observatory For Infrared Astronomy, SOFIA is being developed by NASA and the German space agency, Deutschen Zentrum fur Luft- und Raumfahrt (DLR), with an international contractor team. The 2.5-meter reflecting telescope of SOFIA will be the world's largest airborne telescope. Flying in an open cavity on a modified 747 aircraft, SOFIA will perform infrared astronomy while cruising at 41,000 feet and while being buffeted by a 550- mile-per-hour slipstream. A primary system requirement of SOFIA is tracking stability of 0.2 arc-seconds, and a 3-axis pointing control model has been used to evaluate the feasibility of achieving this kind of stability. The pointing control model shows that increased levels of damping in certain elastic modes of the telescope assembly will help achieve the tracking stability goal and also expand the bandwidth of the attitude controller. This paper describes the preliminary work that has been done to approximate the reduction in image motion yielded by various structure configurations that use reaction masses to attenuate the flexible motions of the telescope structure. Three approaches are considered: passive tuned-mass dampers, active-mass dampers, and attitude control with reaction-mass actuators. Expected performance improvements for each approach, and practical advantages and disadvantages associated with each are presented.
Despite growing international interest in small satellites, high dedicated expendable launch vehicle costs and the lack of secondary launch opportunities continue to hinder the full exploitation of small satellite technology. In the United States, the Department of Defense (DoD), NASA, other government agencies, commercial companies, and many universities use small satellites to perform space experiments, demonstrate new technology, and test operational prototype hardware. In addition, the DoD continues to study the role of small satellites in fulfilling operational mission requirements. However, the US lacks sufficient small satellite launch capacity. Furthermore, US government agencies are restricted to the use of US launch vehicles, which eliminates many affordable launch opportunities. In an effort to increase the number of space experiments that can be flown with a small, fixed budget, the DoD Space Test Program (STP) has teamed with the Air Force Research Laboratory Space Vehicles Directorate (AFRL/VS) to develop a low-cost solution for the small satellite launch program. Our solution, which can be implemented on both Boeing and Lockheed-Martin Evolved Expendable Launch Vehicle-Medium (EELV-M) boosters, is called the EELV Secondary Payload Adaptor (ESPA). ESPA will increase the number of launch opportunities for 180kg-class (or smaller) satellites at prices highly competitive with other secondary launch services worldwide.
KEYWORDS: Satellites, Space operations, Optical isolators, Interfaces, Systems modeling, Vibration isolation, Manufacturing, Finite element methods, Analytical research, Defense and security
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.
KEYWORDS: Gyroscopes, Temperature metrology, Finite element methods, Hubble Space Telescope, Control systems design, Sensors, Titanium, Hardware testing, Control systems, Systems modeling
The Hubble Space Telescope (HST) is currently operating with two flexible solar arrays (or 'wings'), referred to as SA2, that were installed during Servicing Mission 1. These flexible solar arrays are to be replaced with two rigid solar arrays, SA3, during Servicing Mission 3B which is currently scheduled for May, 2001. The key requirements for these arrays are to: (1) increase long term power to support the HST mission, (2) improve the jitter performance while maintaining stability margin requirements, and (3) withstand re-boost loads without astronaut or ground intervention. Analysis of the original SA3 design showed that the Pointing Control System (PCS) stability margin requirements would be violated because of the modal characteristics of the SA3 fundamental bending modes. One of the options to regain the stability margins was to increase the damping of these modes. Damping of 1.5% of critical of the SA3 fundamental bending modes, at the HST system level, is needed to meet stability margin requirements. Therefore, the development of a discrete damping device was undertaken to provide adequate damping of the SA3 fundamental bending modes for all operational conditions.
KEYWORDS: Titanium, Temperature metrology, Finite element methods, Hubble Space Telescope, Telescopes, Space telescopes, Space operations, Solar energy, Aerospace engineering, Optical instrument design
This paper describes the design of a solar array damper that will be built into each of two new solar arrays to be installed on the Hubble Space Telescope (HST) during Servicing Mission 3. On this mission, currently scheduled for August 2000, two 'rigid' solar array wings will replace the 'flexible' wings currently providing power for HST. In addition to increased power, the new arrays will provide the capability for HST to survive re-boost to a higher orbit. The objective of the damper is to reduce the dynamic interaction of these new wings with the Telescope spacecraft. The damper, which is integral to the mast of the solar array, suppresses the fundamental bending modes of the deployed wings at 1.2 Hz (in-plane) and 1.6 Hz (out-of-plane). With the flight version of the damper, modal damping of 2.3% of critical is expected over the temperature range of -4 degrees Celsius to 23 degrees Celsius with a peak damping level of 3.9%. The unique damper design, a combination of titanium spring and viscoelastic damper, was developed using a system finite element model of the solar array wing and measured viscoelastic material properties. Direct complex stiffness (DCS) testing was performed to characterize the frequency- and temperature-dependent behavior of the damper prior to fixed- base modal testing of the wing at NASA/Goddard Space Flight Center (GSFC).
KEYWORDS: Composites, Foam, Finite element methods, Material characterization, Adhesives, Temperature metrology, Structural design, Manufacturing, Chemical elements, Epoxies
Conventional composite materials have high stiffness-to-weight ratios but exhibit little damping; many viscoelastic materials provide high levels of energy dissipation with minimal structural stiffness. The objective of this work was to combine these two material types to produce highly damped structural elements with favorable stiffness and weight characteristics. Cocuring refers to the inclusion of one or more layers of viscoelastic damping material sandwiched between composite plies prior to curing of the composite. Cocured viscoelastic/composite layups were studied experimentally at the material level, modeled analytically, and used to build optimized damped structural components. Measured cocured material properties were used in finite element models to design damped components which were built and tested individually and as part of a truss test structure. Load-carrying and highly damped struts and panels were fabricated. The curing process modified the viscoelastic behavior to some degree, but the materials retained significant, and predictable, damping capability.
KEYWORDS: Space operations, Finite element methods, Satellites, Systems modeling, Antennas, Aluminum, Epoxies, Performance modeling, Solar cells, Photography
FORTE is a small satellite being developed by Los Alamos National Laboratory (LANL) and Sandia National Laboratories Albuquerque (SNLA). It will be placed into orbit via a Pegasus launch in 1996. Testing a full-scale engineering model of the structure using the proto- qualification, system-level vibration spectrum indicated that acceleration levels caused by structural resonances exceed component levels to which certain sensitive components had previously been qualified. Viscoelastic struts were designed to reduce response levels associated with these resonances by increasing the level of damping in key structural modes of the spacecraft. Four identical shear-lap struts were fabricated and installed between the two primary equipment decks. The struts were designed using a system finite element model (FEM) of the spacecraft, a component FEM of the strut, and measured viscoelastic properties. Direct complex stiffness testing was performed to characterize the frequency-dependent behavior of the struts, and these measured properties (shear modulus and loss factor) were used to represent the struts in the spacecraft model. System-level tests were repeated with the struts installed and the response power spectral densities at critical component locations were reduced by as much as 10 dB in the frequency range of interest.
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