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An Overview of Spacecraft Particulate Contamination Phenomena
Published in K. L. Mittal, Particles On Surfaces, 2020
Michael C. Fong, Aleck L. Lee, Paul T. Ma
This paper presents an overview of selected topics on spacecraft (S/C) particulate contamination phenomena. Three topics are discussed: (a) characterization of surface particulate contamination in terms of obscuration ratio (OR), or particle area coverage, as well as correlation between OR and the MIL-STD-1246B cleanliness level, (b) spacecraft contamination due to vibroacoustically induced particle suspension and redistribution in the space shuttle payload bay (PLB) or in the payload fairing (PLF) cavity of an expendable launch vehicle (ELV) during ascent flight, and (c) thermo-optical surface performance degradation due to particulate contamination. These particulate contamination topics are important to spacecraft system design with stringent requirements to maintain thermal and optical system performance aboard the S/C during mission flight. Sample prediction results are included in this paper to demonstrate the analytical models for these topics.
Commercial Space Technologies
Published in Mohammad Razani, Commercial Space Technologies and Applications, 2018
Once the launch vehicle is fully integrated, it is then joined with its payload. This process is called payload integration. The payload will arrive at the launch site from the manufacturing or checkout site to a specialized facility designed to handle the unique needs of the payload. For example, payloads may require fueling, last-minute integration with components, or final testing and checkout. It is then attached to a payload adapter. The payload adapter is the physical connection between the payload and the launch vehicle, and can be integrated with the launch vehicle either horizontally or vertically depending on the vehicle. Once integrated, the payload fairing is installed. The vehicle and payload then make their way to the launch pad, where the combination continues to be monitored during a technical checklist called a countdown. Fueling of a vehicle using liquid propellants takes place at the pad, usually immediately prior to launch.
Large Mirror Design
Published in Paul Yoder, Daniel Vukobratovich, Opto-Mechanical Systems Design, 2017
In 1996 and 1997, proposals were submitted to NASA for a Next Generation Space Telescope (NGST) as a follow on to the HST (see Next Generation Space Telescope, 1997). The proposed NGST would be an IR telescope that would be diffraction limited at wavelengths of 2 μm and longer. The primary mirror diameter was to be between 6 and 8 m. Suppression of instrumental IR background required cryogenic telescope operation at <70 K. Because the maximum payload fairing diameter of existing launch vehicles (Ariane) was about 4.5 m, a segmented primary mirror, deployed on orbit, would be necessary. Launch vehicle payload capacity limited primary mirror weight to about 1000 kg, which meant that the maximum areal density would be about 20 kg/m2 for an 8 m aperture. According to Coulter and Jacobson (2000), NASA sought to reduce the areal density to about 15 kg/m2 to provide design margin. Both aperture and areal density were well beyond the technology of the HST; the aperture being 2.5–3.3 times bigger and the areal density 12 times smaller. Development of mirror technology for the NGST led ultimately to the JWST primary mirror. Although the JWST design is not yet confirmed by on-orbit performance, it represents the current state-of-the-art in lightweight mirror technology.
Design of a tuned vibration absorber for a slender hollow cylindrical structure
Published in Mechanics Based Design of Structures and Machines, 2020
Tuğrul Aksoy, Gökhan Osman Özgen, Bülent Acar
Harris (2003) designed a two DOF TVA to solve the resonant vibration problem of a payload fairing (resembling a massive hollow cylinder structure) sitting on ground, targeting FRF amplitude reduction at first and second resonant frequencies of the structure. The structure is experimentally identified and modeled. Target resonant frequencies are obtained from these experiments. In order to simulate the effect of the TVA on the structure, TVA is modeled using lumped mass, stiffness and damping elements and it is coupled to the experimentally obtained FRFs of the structure to simulate the effect of the added TVA. A physical design for the TVA is also proposed comprising a foam block as spring-damper element (with viscoelastic damping mechanism) and steel plates as mass elements where two single DOF TVAs are stacked together to form a two DOF TVA. Performance of the TVA design is also tested on a small-scale representative payload fairing structure through measured FRFs of the structure with and without the multiple two DOF TVAs. TVA designs are also validated experimentally as standalone systems using shaker tests (through transmissibility measurements).
LEU NTP Engine System Trades and Mission Options
Published in Nuclear Technology, 2020
C. Russell Joyner, Michael Eades, James Horton, Tyler Jennings, Timothy Kokan, Daniel J. H. Levack, Brian J. Muzek, Christopher B. Reynolds
Many of the benefits for using the LEU NTP on Mars crew mission architectures have been reported in previous AR papers. The trade studies from 2016 and 2017 showed that using a cis-lunar aggregation orbit and lunar-distant highly elliptical orbit (LDHEO)–type orbit for Earth departure (at perigee) and return was beneficial to reducing the NTP Mars crew vehicle mass while still providing a space transportation system with 5 to 6 months of transit capability from Earth to Mars. This Mars NTP vehicle can be assembled in four to five NASA SLS Block 2 launches using the 8.4-m payload fairing (PLF), including the crew habitat. The NTP Mars transfer vehicle (MTV) is a robust solution for a Mars vehicle and using the various stage elements in different approaches can provide an even more versatile exploration system approach.
Decoding Mission Design Problem for NTP Systems for Outer Planet Robotic Missions
Published in Nuclear Technology, 2022
Saroj Kumar, L. Dale Thomas, Jason T. Cassibry
The problem tackles multiple areas such as spacecraft and NTP system design based on mission objectives. The NTP system and spacecraft design parameters are also evaluated for launch vehicle constraints. This approach makes sure that the overall design is within the limits specified by each system.24 The problem statement begins with the science mission objectives, which determines whether a mission will be a flyby, rendezvous, or round trip. This information is used toward the spacecraft and NTP system design. For the scope of this study, the spacecraft design will be considered only with respect to its total mass and dimensions. The NTP system configuration is then addressed, which is based on the expendable or nonexpendable nature of the mission along with ∆V requirements, engine thrust class, and liquid hydrogen (LH2) propellant tank sizing. The spacecraft and NTP system configurations such as Initial Mass in Low Earth Orbit (IMLEO) and payload fairing encapsulation are evaluated based on commercial launch vehicle limitations. Because this exercise is multidisciplinary in nature, the issues during the design of one system do not exist in isolation but feed upon other systems as well. This problem is solved iteratively until multiple cross-dependent parameters starting from spacecraft design to NTP systems and launch vehicle requirements are satisfied. The initial calculations on ∆V requirements and trip-time estimates are based on the patched conic analysis performed on Matlab, and finally end-to-end high-fidelity trajectory design and optimization are performed using the Systems Tool Kit Astrogator module.