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At least 1,171 records · Page 65

Lunar module alightment system

The lunar module (LM) landing phase is one of the most critical phases of the entire Apollo mission. Landing impact is of critical concern because at no other time during the LM mission is there less opportunity for vehicle control. To assure a successful landing, the alightment system must not only provide required shock attenuation for protection of crew and structure, but must also assure stability against overturning after touchdown. The paper discusses important design features of the alightment system, such as geometric arrangement, energy absorber characteristics, and deployment mechanisms. Performance characteristics of the vehicle at touchdown, as found analytically, are presented. Model and full-size testing of the landing system is briefly discussed.

LANDING SYSTEM↗

Structural members, method and apparatus

A method and apparatus for fabricating a structural member such as truss from flexible sheet material in compacted form are disclosed. A number of generally tubular columns are progressively formed from the sheet material and deployed generally parallel to one another. Adjacent pairs of the columns are interconnected by respective side members, each of which is comprised of a strip of the sheet material. The sheet material is fastened together by self-attaching fasteners integrally formed from the sheet material of the columns and side members themselves.

Kinzler, J. A.↗

Structures for large precision reflectors

The design and fabrication of a highly accurate large aperture infrared reflecting telescope is discussed. The structural concept is an off axis Cassegrainian design with a focal length equal to the aperture diameter. Thermal shielding, peak heat fluxes, cooling rates, active control tasks, and deployment sequence are discussed.

Hedgepeth, J. M.↗

Deployable beam flight experiment (MAST)

Improvement of technology of space systems control is discussed. Future space systems such as large antennas or a space station may have dimensions on the order of 30 m to 200 m, yet their basic structures may be relatively lightweight and flexible, making ground tests for loads, controls analyses, and design verifications questionable if not impossible. Abandoning the extensive ground test and analysis verification program that led to the success of previous spacecraft is not a sensible option; making it meaningful using current technology will require inefficient, ultraconservative structure and control designs. New test methods are outlined.

Hanks, B. R.↗

Development of the Toggle Deployment Mechanism

The Toggle Deployment Mechanism (TDM) is a two fault tolerant, single point, low shock pyro/mechanical releasing device. Many forms of releasing are single fault tolerant and involve breaking of primary structure. Other releasing mechanisms, that do not break primary structure, are only pyrotechnically redundant and not mechanically redundant. The TDM contains 3 independent pyro actuators, and only one of the 3 is required for release. The 2 separating members in the TDM are held together by a toggle that is a cylindrical stem with a larger diameter spherical shape on the top and flares out in a conical shape at the bottom. The spherical end of the toggle sits in a socket with the top assembly and the bottom is held down by 3 pins or hooks equally spaced around the conical shaped end. Each of the TDM's 3 independent actuators shares a third of the separating load and does not require as much pyrotechnic energy as many single fault tolerant actuators. Other single separating actuators, i.e., separating nuts or pin pullers, have the pyrotechnic energy releasing the entire preload holding the separating members together. Two types of TDM's ,described in this paper, release the toggle with pin pullers, and the third TDM releases the toggle with hooks. Each design has different advantages and disadvantages. This paper describes the TDM's construction and testing up to the summer of 1993.

Brown, Christopher W.↗

The Thermal Hogan - A Means of Surviving the Lunar Night

document describes the Thermal Hogan, a new shelter concept that would be used on the Moon to moderate the extreme nighttime temperatures, allowing survival of equipment with minimal heater power. It is lightweight, has few mechanical parts, and would be relatively easy to deploy on the Moon. The Lunar Hogan has two parts: an insulated shelter and a thermal mass. The shelter is constructed of multilayer insulation (MLI) draped over a structural framework. Entry and egress are accomplished either by raising the structure or via a door constructed of the same MLI material as the shelter. The thermal mass can be manufactured from locally available materials, either by piling substantially sized rocks to a depth of 0.25 meter, or by filling a 0.25-meter-deep conductive honeycomb-like structure with lunar dust. For ease of transport, the structural framework and honeycomb can be collapsible. The door can be opened by pushing on it in either direction. Gravity would cause it to close and it could be sealed via magnetic strips on the doorframe.

Fruitwala, Neelay↗

Mars Pathfinder Science Investigations and Objectives

Pathfinder's primary ojective is to demonstrate an inexpensive system for curise, entry, descent, and landing on mars. Additional objective include the deployent and operation of various science instruments: a stereoscopic imager with filters on a pop up mast, and alpha proton x-ray spectometer (APXS), and an atmospheric structure instrument/meteorology package

Mars↗

Near Earth Asteroid Scout - Mission Update

After its deployment from NASA’s Space Launch System (SLS), the Near-Earth Asteroid (NEA) Scout mission will travel to and image an asteroid during a close flyby using an 86m^2 solar sail as its primary propulsion. Solar sails are large, mirror-like structures made of a lightweight material that reflects sunlight to propel the spacecraft. The continuous solar photon pressure provides thrust with no need for the heavy, expendable propellants used by conventional chemical and electric propulsion systems. Developed by NASA’s Marshall Space Flight Center (MSFC) and Jet Propulsion Laboratory (JPL), the NEA Scout is based on the industry-standard CubeSat form factor. The spacecraft measures 11 cm x 24 cm x 36 cm and weighs less than 14 kilograms. Following deployment from the Space Launch System (SLS), the solar sail will deploy, and the spacecraft will begin its 2.0 – 2.5-year journey. About one month before the asteroid flyby, NEA Scout will search for the target and start its Approach Phase using a combination of radio tracking and optical navigation and perform a relatively slow flyby (10-20 m/s) of the target. A summary of the mission, sailcraft, mission design, and its first several months of deep space operation will be described.

Near Earth Asteroid↗

Fully Coupled Aeroelastic Stability Analysis of Adaptive Shape Memory Alloy Structural Technologies for Airframe Noise Reduction

This final report documents work performed by ATA Engineering, Inc., (ATA) to develop computational models and analyze the coupled fluid-structure response of two types of noise treatments applied to the leading-edge-slat component of a high-lift system typical of modern transport aircraft. The first treatment is a slat-gap filler (SGF), which closes the gap between the suction surfaces of a deployed slat and an aircraft main wing, and the second treatment is a slat-cove filler (SCF), which replaces the recirculating flow on the slat cove with a surface that promotes flow attachment. The representative airframe chosen for this work was NASA’s High-Lift Common Research Model (CRM-HL) in a baseline high-lift configuration. Superelastic shape memory alloys (SMAs) have been identified as enabling materials for these structural treatments. Since the technology elements rely upon having a highly reconfigurable structure, designs were assessed for their static aeroelastic deflection as well as their dynamic aeroelastic stability using coupled computational fluid dynamics (CFD) and nonlinear computational structural dynamics (NL-CSD) tools. Specifically, fluid-structure interaction (FSI) problems were solved computationally using the CFD solver Loci/CHEM and the NL-CSD solver Abaqus. As a part of the overall project, a similar capability was implemented using the CFD solver FUN3D coupled to Abaqus, although that work is documented in a separate report and that FSI framework was not used to analyze any of NASA’s SGF and SCF configurations. The technical approach consisted of solving for the flow field around the entire vehicle using a global CFD model, followed by extraction of relevant local subdomain data for CFD and NL-CSD co-simulations. The SGF design was analyzed using both 2D and 3D co-simulations to predict quasi-static aeroelastic deformations and to assess dynamic aeroelastic stability, whereas the SCF was analyzed in 2D only. SGF static aeroelastic response predictions focused on characterizing the deformed shape, with maximum displacements predicted to be on the order of magnitude of the technology element panel thickness. SGF dynamic aeroelastic response predictions used Partial Floquet analysis of the temporal evolution of selected nodal displacements to quantify the sign and magnitude of aeroelastic damping. Results suggest that the CRM-HL operating conditions would result in a dynamically stable response. The simulated dynamic pressure was also increased up to a factor of about four, and resulting responses suggest that predicted dynamic stability would be achieved with some margin.

Fluid Structure Interaction↗

Nonlinear dynamic analysis of deploying flexible space booms

The dynamics of the planar deployment and lock-up of two flexible boom-type appendages on a spacecraft are analyzed by establishing nonlinear equations of motion for the deployment phase and linear for the post-lock-up phase. Nondimensional parameters and threshold values are identified in terms of boom tip deflections, deployment time and root moments for appendages with tip masses attached to a central rigid body through a rotational spring; an account is taken of the nonlinear kinematic and structural terms. It was found that, in general, the solution is influenced only by two mass ratios and one nondimensional stiffness parameter.

Mcgowan, P. E.↗

Composite material technology requirements for large precision space structures

The development of dimensionally stable, precision composite structures has been recognized as a high risk technology driver in NASA's continuing large space structures research. Attempts are being made to understand the influences controlling thermal performance in such composites, and specifically in composite sandwich panels. The necessary tools for such composite panels' deployment, the experimental verification of analytical predictions, and the demonstration of technology in small scale hardware, are presently addressed.

Mcelroy, P. M.↗

Experiments of opportunity pay loads revisited

The design and operation of the Spartan program are discussed. The objective of the Spartan program is to utilize the Space Shuttle for conducting exoatmospheric science research. The Spartan 1 data storage and handling, attitude control, power, and thermal control systems and structure are described. The Spartan 1 was to study the structure of two prominent cosmic X-ray sources, the Perseus cluster of galaxies, and the center of the Milky Way. Spartan 1 was launched on June 17, 1985, deployed on June 20, and retrieved on June 22. The performance of the Spartan's system, and the collected data are analyzed; it was observed that the systems performed as planned. The methods used to minimize the Spartan program costs are considered.

Shrewsberry, D. J.↗

Programmatic and technical parameters associated with Shuttle/Centaur development and operations

The cost, schedule, technical, integration, and performance parameters associated with the development of the Shuttle/Centaur configurations, Centaur G and Centaur G-Prime, for compatibility with various elements of the STS are discussed. It is noted that the design concepts for these configurations involved the use of as much flight-proven hardware from the Atlas/Centaur program as possible. Forward and aft adapters were designed using graphite/epoxy composite structures for strength, weight, and thermal flow considerations. The Centaur integrated support system serves as the system interface for all systems between the Shuttle and Centaur from prelaunch through deployment of the Centaur in LEO. As a result of the Shuttle accident of January 28, 1986, the earliest possible flight for the Centaur mission appears to be December 1987.

Clark, Harry J.↗

The development of a complementary expendable launch vehicle interface for an STS deployable payload

The development is described of an interface, the Titan Payload Adapter (TPA), between a Space Transportation System (STS) deployable payload and an expendable launch vehicle (ELV). Separate ascent and separation constraint systems allow a payload with integral trunnions to retain its originally designed, boost-phase load structure, yet also allow the expendable booster vehicle to separate from the payload via retro-rockets. Design requirements as well as development problems and their solutions are discussed.

Eubanks, ED↗

Stroboscopic Interferometer for Measuring Mirror Vibrations

Stroboscopic interferometry is a technique for measuring the modes of vibration of mirrors that are lightweight and, therefore, unavoidably flexible. The technique was conceived especially for modal characterization of lightweight focusing mirror segments to be deployed in outer space; however, the technique can be applied to lightweight mirrors designed for use on Earth as well as the modal investigation of other optical and mechanical structures. To determine the modal structure of vibration of a mirror, it is necessary to excite the mirror by applying a force that varies periodically with time at a controllable frequency. The excitation can utilize sinusoidal, square, triangular, or even asynchronous waveforms. Because vibrational modes occur at specific resonant frequencies, it is necessary to perform synchronous measurements and sweep the frequency to locate the significant resonant modes. For a given mode it is possible to step the phase of data acquisition in order to capture the modal behavior over a single cycle of the resonant frequency. In order to measure interferometrically the vibrational response of the mirror at a given frequency, an interferometer must be suitably aligned with the mirror and adjustably phase-locked with the excitation signal. As in conventional stroboscopic photography, the basic idea in stroboscopic interferometry is to capture an image of the shape of a moving object (in this case, the vibrating mirror) at a specified instant of time in the vibration cycle. Adjusting the phase difference over a full cycle causes the interference fringes to vary over the full range of motion for the mode at the excitation frequency. The interference-fringe pattern is recorded as a function of the phase difference, and, from the resulting data, the surface shape of the mirror for the given mode is extracted. In addition to the interferometer and the mirror to be tested, the equipment needed for stroboscopic interferometry includes an arbitrary-function generator (that is, a signal generator), an oscilloscope, a trigger filter, and an advanced charge-coupled-device (CCD) camera. The optical components are positioned to form a pupil image of the mirror under test on the CCD chip, so that the interference pattern representative of the instantaneous mirror shape is imaged on the CCD chip.

Stahl, H. Philip↗

Technology Candidates for Air-to-Air and Air-to-Ground Data Exchange

Technology Candidates for Air-to-Air and Air-to-Ground Data Exchange is a two-year research effort to visualize the U. S. aviation industry at a point 50 years in the future, and to define potential communication solutions to meet those future data exchange needs. The research team, led by XCELAR, was tasked with identifying future National Airspace System (NAS) scenarios, determining requirements and functions (including gaps), investigating technical and business issues for air, ground, & air-to-ground interactions, and reporting on the results. The project was conducted under technical direction from NASA and in collaboration with XCELAR's partner, National Institute of Aerospace, and NASA technical representatives. Parallel efforts were initiated to define the information exchange functional needs of the future NAS, and specific communication link technologies to potentially serve those needs. Those efforts converged with the mapping of each identified future NAS function to potential enabling communication solutions; those solutions were then compared with, and ranked relative to, each other on a technical basis in a structured analysis process. The technical solutions emerging from that process were then assessed from a business case perspective to determine their viability from a real-world adoption and deployment standpoint. The results of that analysis produced a proposed set of future solutions and most promising candidate technologies. Gap analyses were conducted at two points in the process, the first examining technical factors, and the second as part of the business case analysis. In each case, no gaps or unmet needs were identified in applying the solutions evaluated to the requirements identified. The future communication solutions identified in the research comprise both specific link technologies and two enabling technologies that apply to most or all specific links. As a result, the research resulted in a new analysis approach, viewing the underlying architecture of ground-air and air-air communications as a whole, rather than as simple "link to function" paired solutions. For the business case analysis, a number of "reference architectures" were developed for both the future technologies and the current systems, based on three typical configurations of current aircraft. Current and future costs were assigned, and various comparisons made between the current and future architectures. In general, it was assumed that if a future architecture offers lower cost than the current typical architecture, while delivering equivalent or better performance, it is likely that the future solution will gain industry acceptance. Conversely, future architectures presenting higher costs than their current counterparts must present a compelling benefit case in other areas or risk a lack of industry acceptance. The business case analysis consistently indicated lower costs for the proposed future architectures, and in most cases, significantly so. The proposed future solutions were found to offer significantly greater functionality, flexibility, and growth potential over time, at lower cost, than current systems. This was true for overall, fleet-wide equipage for domestic and oceanic air carriers, as well as for single, General Aviation (GA) aircraft. The overall research results indicate that all identified requirements can be met by the proposed solutions with significant capacity for future growth. Results also illustrate that the majority of the future communication needs can be met using currently allocated aviation RF spectrum, if used in more effective ways than it is today. A combination of such optimized aviation-specific links and commercial communication systems meets all identified needs for the 50-year future and beyond, with the caveat that a new, overall function will be needed to manage all information exchange, individual links, security, cost, and other factors. This function was labeled "Delivery Manager" (DM) within this research. DM employs a distributed client/server architecture, for both airborne and ground communications architectures. Final research results included identifying the most promising candidate technologies for the future system, conclusions and recommendations, and identifying areas where further research should be considered.

Communications↗

Summary of 2006 to 2010 FPMU Measurements of International Space Station Frame Potential Variations

Electric potential variations on the International Space Station (ISS) structure in low Earth orbit are dominated by contributions from interactions of the United States (US) 160 volt solar arrays with the relatively high density, low temperature plasma environment and inductive potentials generated by motion of the large vehicle across the Earth?s magnetic field. The Floating Potential Measurement Unit (FPMU) instrument suite comprising two Langmuir probes, a plasma impedance probe, and a floating potential probe was deployed in August 2006 for use in characterizing variations in ISS potential, the state of the ionosphere along the ISS orbit and its effect on ISS charging, evaluating effects of payloads and visiting vehicles, and for supporting ISS plasma hazard assessments. This presentation summarizes observations of ISS frame potential variations obtained from the FPMU from deployment in 2006 through the current time. We first describe ISS potential variations due to current collection by solar arrays in the day time sector of the orbit including eclipse exit and entry charging events, potential variations due to plasma environment variations in the equatorial anomaly, and visiting vehicles docked to the ISS structure. Next, we discuss potential variations due to inductive electric fields generated by motion of the vehicle across the geomagnetic field and the effects of external electric fields in the ionosphere. Examples of night time potential variations at high latitudes and their possible relationship to auroral charging are described and, finally, we demonstrate effects on the ISS potential due to European Space Agency and US plasma contactor devices.

Minow, Joseph I.↗

Shuttle operational expectations

The results of orbital flight tests (OFT) of the Space Shuttle are reviewed, and modifications planned for upcoming operational flights are discussed. The performance of the solid rocket boosters, external tank, main engines, structural system, propulsion system, reaction control system, electric power system, heat rejection system, hydraulic system, avionics, and other systems is described and evaluated as generally highly satisfactory. Payload servicing and deployment were also successfully demonstrated by OFT. Additional facilities planned for the operational flights are briefly described, and improvements that will make the Challenger spacecraft lighter than Columbia, provide it with more thrust, and give it a larger payload are summarized. Some software modifications being introduced are also mentioned.

Abrahamson, J. A.↗