Packaging the LEM rendezvous radar and transponder.
Packaging technique for producing rendezvous radar and transponder noting structural design, reliability, cooling, outgassing, humidity, etc
SPACECRAFT INSTRUMENTATION↗
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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
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Ballistic Performance of Porous Ceramic Thermal Protection Systems at 9 km/s
Porous-ceramic, thermal-protection-systems are used heavily in current reentry vehicles like the Orbiter, and they are currently being proposed for the next generation of manned spacecraft, Orion. These materials insulate the structural components and sensitive electronic components of a spacecraft against the intense thermal environments of atmospheric reentry. Furthermore, these materials are also highly exposed to space environmental hazards like meteoroid and orbital debris impacts. This paper discusses recent impact testing up to 9 km/s on ceramic tiles similar to those used on the Orbiter. These tiles have a porous-batting of nominally 8 lb/cubic ft alumina-fiber-enhanced-thermal-barrier (AETB8) insulating material coated with a damage-resistant, toughened-unipiece-fibrous-insulation (TUFI) layer.
Miller, Joshua E.↗
Structural interaction with control systems
A monograph which assesses the state of the art of space vehicle design and development is presented. The monograph presents criteria and recommended practices for determining the structural data and a mathematical structural model of the vehicle needed for accurate prediction of structure and control-system interaction; for design to minimize undesirable interactions between the structure and the control system; and for determining techniques to achieve the maximum desirable interactions and associated structural design benefits. All space vehicles are treated, including launch vehicles, spacecraft, and entry vehicles. Important structural characteristics which affect the structural model used for structural and control-system interaction analysis are given.
Noll, R. B.↗
Docking and retrieval mechanism
An engineering prototype docking and retrieval mechanism (DRM) which enables two spacecraft to dock and be structurally joined on-orbit is described. The joining of two spacecraft or payloads on-orbit supports future planned space activities such as payload servicing, deployment and retrieval, and assembly or large space systems. Advantages of the DRM include: it is a nonimpact docking mechanism; does not require impact absorbing mechanisms or attitude stabilization on the target spacecraft; is capable of docking to a spinning spacecraft; and can spin up and deploy a spinning spacecraft or payload.
Tewell, J. R.↗
Rationale for an experimental test for flexible space structure attitude control
The problems of large flexible spacecraft control are characterized by the infinite bandwidth of structural vibrations, which cannot be accounted for in the dynamic design model. This may lead to instability even, if ideal control hardware is assumed, which can be concluded from preceding numerical investigations. Additional performance limitations are expected to occur due to hardware constraints. A laboratory experiment is proposed to investigate the key problems in more detail. The test setup requirements being defined by the idealized control system are extremely high demanding a high speed processor and special hardware component developments. The test element is a wire suspended plate being controlled by an array processor via high performance sensors and actuators. First tests on component level indicate the feasibility of the system presently being developed.
Lange, T.↗
Multi-Mission Suitability of the NASA Ames Modular Common Bus
The obvious advantages of small spacecraft - their lower cost structure and the rapid development schedule - have enabled a large number of missions in the past. However, most of these missions have been focused on Earth observation from low Earth orbits. In 2006, the Small Spacecraft Division at the NASA Ames Research Center began the development of the Modular Common Bus, a spacecraft capable of delivering scientifically and technically useful payloads to a variety of destinations within 0.1 AU around the Earth. The core technologies used in the Common Bus design are a composite structure with body-mounted solar cells, an integrated avionics unit, and a high performance bipropellant propulsion system. Due to its modular approach, the Common Bus can be adapted to fit specific mission needs while still using a standardized and qualified set of components. Additionally a number of low cost launch vehicles are supported, resulting in overall mission costs of around $150M including the launch vehicle but excluding the science payloads. This significant reduction in cost and the shorter development time would enable NASA to conduct more frequent exploration missions within its budget and timeframe constraints, compared to the status quo. In this paper the suitability of the Common Spacecraft Bus for four different exploration scenarios is analyzed. These scenarios include a lunar orbiter, a lunar lander, a mission to a Sun-Earth Libration Point, and a rendezvous mission to a Near Earth Object. For each scenario, a preliminary design reference mission is developed and key design parameters for the spacecraft are determined.
Tietz, Sascha↗
ISEE-1 and -2 observations of magnetotail flux ropes - FTEs in the plasma sheet?
ISEE-1 and 2 observations from about 20 Re down the near-earth magnetotail indicate the presence of magnetic flux ropes in the neutral sheet. Magnetic and electric field and fast plasma data show that these structures convect across the spacecraft at speeds of 200-600-km/s, and have scale sizes of roughly 3 5-Re. The rope axis orientation is approximately cross-tail. Their magnetic structure is similar to Venus ionospheric flux ropes, and to flux transfer events at the dayside magnetopause. These structures may arise from patchy reconnection or tearing mode reconnection within the plasma sheet.
Elphic, R. C.↗
Earth Observatory Satellite system definition study. Report no. 3: Design/cost tradeoff studies. Appendix D: EOS configuration design data. Part 1: Spacecraft configuration
The results of structural studies of the Earth Observatory Satellite (EOS) which define the member sizes to meet the vehicle design requirements are presented. The most significant requirements in sizing the members are the stiffness required to meet the launch vehicle design frequencies both in the late al and in the longitudinal directions. The selected configurations, both baseline and preferred, for the Delta and Titan launch vehicles were evaluated for stiffness requirements. The structural idealization used to estimate the stiffness of each structural arrangement, was based on an evaluation of primary loads paths, effectivity of structural members, and estimated sizes for the preferred configurations. The study included an evaluation of the following structural materials: (1) aluminum alloys, (2) titanium alloys, (3) beryllium, (4) beryllium/aluminum alloy, and (5) composite materials.
Source record↗
Coordinate transformation and truncation for rotating spacecraft with flexible appendages
The analytical structures of alternative coordinate transformations for the variables that characterize the deformations of finite element models of flexible appendages on rotating spacecraft are examined. Particular emphasis is given to the truncation operations that are essential for efficient numerical simulations of flexible spacecraft, due to the necessarily large number of coordinates employed in a finite element description. A theorem is developed which establishes necessary conditions for the commutativity of the required truncation and inversion operations. The recommended alternative for the rotating elastic appendage is a complex transformation with previously published orthogonality properties, which permit the derivation in this paper of a relatively simple explicit set of transformed and truncated equations of modal vibration in real variables.
Likins, P.↗
Practical approaches to the design of control systems for large space structures
Future space systems, flexible spacecraft, modular control, design-test-revise cycle, and control technique verification are discussed.
Buchanan, H. J.↗
A comparative study of distant magnetotail structure at Venus and earth
Spacecraft data on the magnetotails of earth and Venus are discussed with regards to interactions between the two planets and the solar wind. Consideration is given to Pioneer Venus Orbiter plasma analyzer, magnetometer and plasma wave observations around Venus and equivalent ISEE-3 data from the terrestrial region. The discussion is constrained to clear lobe and plasmasheet intervals greater than 5 and 30 min for Venus and earth, respectively. The orientation of the Venusian plasma sheet was normal to that of earth's, while both regions displayed distinct mantle, lobe, plasmasheet boundary layer, and plasmasheet regions. Additionally, the plasmasheet magnetic field distributions in distant regions of the tails exhibited characteristics of disconnection from the planets and a return to the solar wind. The lobe field magnitudes were consistent with the interaction of the planets with the solar wind. It is noted that the Venus plasmasheet is hotter and/or denser than the terrestrial magnetosheath.
Slavin, J. A.↗
A steerable radiator for spacecraft application
Future large space structures such as the Space Station will have high dissipation and long life requirements which dictate the requirements for steerable radiators. Several rotary coupling concepts were considered to accomplish heat transport across the steerable radiator system's rotating interface. Rotary fluid couplings were chosen over rotary contact couplings or flexible lines because of low temperature gradients and operational flexibility. A rotary fluid coupling has been designed for use on initial operating capability (IOC) and growth Space Station. Its internal seals have been demonstrated to a 10 year life with minimal (0.055 lbm NH3/year) leakage between internal passages. Materials compatibility tests have been conducted to demonstrate compatibility with ammonia. A seal leakage management system has been designed to: (1) monitor pressures between the internal seals, (2) remove leakage, and (3) store the leakage fluid in a reservoir.
Heizer, B. L.↗
Multidisciplinary analysis of actively controlled large flexible spacecraft
The control of Flexible Structures (COFS) program has supported the development of an analysis capability at the Langley Research Center called the Integrated Multidisciplinary Analysis Tool (IMAT) which provides an efficient data storage and transfer capability among commercial computer codes to aid in the dynamic analysis of actively controlled structures. IMAT is a system of computer programs which transfers Computer-Aided-Design (CAD) configurations, structural finite element models, material property and stress information, structural and rigid-body dynamic model information, and linear system matrices for control law formulation among various commercial applications programs through a common database. Although general in its formulation, IMAT was developed specifically to aid in the evaluation of the structures. A description of the IMAT system and results of an application of the system are given.
Cooper, Paul A.↗
Spatial structure of high energy photon sources in solar flares
Stereoscopic observations of high energy (greater than about 100 keV) photon emission from five solar flares have been made with the X-ray spectrometers aboard the ISEE-3 and Pioneer Venus Orbiter spacecraft. The observed altitude structure of the photon source and its dependence on the photon energy and time during a flare are compared with the predictions of thermal and non-thermal models of the hard X-ray source. In the case of the impulsive source, it is found that (1) the thermal model with adiabatic compression and expansion of a magnetically confined plasma and the thin target (non-thermal) model are not consistent with the observations; (2) the thick target (non-thermal) model and the dissipative thermal model are partially in agreement with the observations; (3) the emission probably originates in many individual non-thermal sources distributed in altitude, the lower altitude sources being brighter than those at higher altitude. In the case of the gradual source, it is found that (1) models with purely coronal sources are not consistent with the observations; (2) a partial precipitation model with trapped as well as precipitating electrons is consistent with the observations.
Kane, S. R.↗
Analytical Approach for Estimating Preliminary Mass of ARES I Crew Launch Vehicle Upper Stage Structural Components
In January 2004, President Bush gave the National Aeronautics and Space Administration (NASA) a vision for Space Exploration by setting our sight on a bold new path to go back to the Moon, then to Mars and beyond. In response to this vision, NASA started the Constellation Program, which is a new exploration launch vehicle program. The primary mission for the Constellation Program is to carry out a series of human expeditions ranging from Low Earth Orbit to the surface of Mars and beyond for the purposes of conducting human exploration of space, as specified by the Vision for Space Exploration (VSE). The intent is that the information and technology developed by this program will provide the foundation for broader exploration activities as our operational experience grows. The ARES I Crew Launch Vehicle (CLV) has been designated as the launch vehicle that will be developed as a "first step" to facilitate the aforementioned human expeditions. The CLV Project is broken into four major elements: First Stage, Upper Stage Engine, Upper Stage (US), and the Crew Exploration Vehicle (CEV). NASA's Marshall Space Flight Center (MSFC) is responsible for the design of the CLV and has the prime responsibility to design the upper stage of the vehicle. The US is the second propulsive stage of the CLV and provides CEV insertion into low Earth orbit (LEO) after separation from the First Stage of the Crew Launch Vehicle. The fully integrated Upper Stage is a mix of modified existing heritage hardware (J-2X Engine) and new development (primary structure, subsystems, and avionics). The Upper Stage assembly is a structurally stabilized cylindrical structure, which is powered by a single J-2X engine which is developed as a separate Element of the CLV. The primary structure includes the load bearing liquid hydrogen (LH2) and liquid oxygen (LOX) propellant tanks, a Forward Skirt, the Intertank structure, the Aft Skirt and the Thrust Structure. A Systems Tunnel, which carries fluid and electrical power functions to other Elements of the CLV, is included as secondary structure. The MSFC has an overall responsibility for the integrated US element as well as structural design an thermal control of the fuel tanks, intertank, interstage, avionics, main propulsion system, Reaction Control System (RCS) for both the Upper Stage and the First Stage. MSFC's Spacecraft and Vehicle Department, Structural and Analysis Design Division is developing a set of predicted mass of these elements. This paper details the methodology, criterion and tools used for the preliminary mass predictions of the upper stage structural assembly components. In general, weight of the cylindrical barrel sections are estimated using the commercial code Hypersizer, whereas, weight of the domes are developed using classical solutions. HyperSizer is software that performs automated structural analysis and sizing optimization based on aerospace methods for strength, stability, and stiffness. Analysis methods range from closed form, traditional hand calculations repeated every day in industry to more advanced panel buckling algorithms. Margin-of-safety reporting for every potential failure provides the engineer with a powerful insight into the structural problem. Optimization capabilities include finding minimum weight panel or beam concepts, material selections, cross sectional dimensions, thicknesses, and lay-ups from a library of 40 different stiffened and sandwich designs and a database of composite, metallic, honeycomb, and foam materials. Multiple different concepts (orthogrid, isogrid, and skin stiffener) were run for multiple loading combinations of ascent design load with and with out tank pressure as well as proof pressure condition. Subsequently, selected optimized concept obtained from Hypersizer runs was translated into a computer aid design (CAD) model to account for the wall thickness tolerance, weld land etc for developing the most probable weight of the components. The flow diram summarizes the analysis steps used in developing these predicted mass.
Aggarwal, Pravin↗
The effects of the sterilization environment on the design of structural systems for planetary landers.
Sterilization environment effects on structural systems design for interplanetary spacecraft
ENVIRONMENTAL CONTROL↗
ISEE-1 and 2 observations of magnetic flux ropes in the magnetotail - FTE's in the plasma sheet?
Magnetic field observations on ISEE-1 and 2 in and near the neutral sheet about 20 Re down the near-earth magnetotail reveal the occurrence of structures resembling magnetic flux ropes. Both electric field and fast plasma data show that these structures convect across the spacecraft at speeds of 200 - 600 km/s, and that they have scale sizes of roughly 3 5 Re. The rope axis orientation is across the tail, approximately in the -Y GSM direction. Their magnetic structure is strikingly similar to magnetic flux ropes observed in the Venus ionosphere, and to flux transfer events observed at the dayside magnetopause. The total field-aligned current within these ropes may approach a million amps. These structures may arise because of patchy reconnection within the plasma sheet, or may be tearing islands formed when the plasma sheet magnetic field has a cross-tail component. Plasma sheet flux ropes are not a common feature at ISEE orbital altitudes; this suggests that near-earth neutral line formation within ISEE apogee (22 Re) may be equally rare.
Elphic, R. C.↗
A design technique for determining actuator gains in spacecraft vibration control
A design procedure is described which determines the gains of a diagonal damping matrix to control the vibrations of a flexible structure with application to orbiting spacecraft. The procedure is based on minimizing the energy dissipated by control actuators using nonlinear mathematical programming. Each damping gain is assumed to be an active viscous damper and the design process is formulated so that the force or torque output of the actuator does not exceed a specified value. The response of the structure at some specified time after the termination of the disturbance is constrained to be less than some prescribed value based upon spacecraft mission performance requirements. A grillage example is used to demonstrate the design process for determining gains for two representative cases. Resulting designs are verified by a finite element analysis of the structure augmented by the control actuators.
Horner, G. C.↗