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NASA NTRS1986

Vibration of a large space beam under gravity effect

Future space structures will have a low mass density and high flexibility, with ground test dynamic behavior differing significantly from that in zero-G orbit. Attention is presently given to the vibration behavior of a beam deformed by its own weight; the results obtained by the differential equations for both the static and dynamic responses of a large, simply supported beam, which are derived and solved analytically, allow ground test experiment measurements to be used for orbital dynamic characteristics verification efforts.

Shih, C.-F.↗

NASA NTRS1983

Technology requirements for large flexible space structures

Research, test, and demonstration experiments necessary for establishing a data base that will permit construction of large, lightweight flexible space structures meeting on-orbit pointing and surface precesion criteria are discussed. Attention is focused on the wrap-rib proof-of-concept antenna structures developed from technology used on the ATS-6 satellite. The target structure will be up to 150 m in diameter or smaller, operate at RF levels, be amenable to packaging for carriage in the Shuttle bay, be capable of being ground-tested, and permit on-orbit deployment and retraction. Graphite/epoxy has been chosen as the antenna ribs material, and the antenna mesh will be gold-plated Mo wire. A 55-m diam reflector was built as proof-of-concept with ground-test capability. Tests will proceed on components, a model, the entire structure, and in-flight. An analytical model has been formulated to characterize the antenna's thermal behavior. The flight test of the 55-m prototype in-orbit offers the chance to validate the analytical model and characterize the control, mechanical, and thermal characteristics of the antenna configuration.

Wada, B. K.↗

NASA NTRS2016

Manufacturing Challenges and Benefits when Scaling the HIAD Stacked-Torus Aeroshell to a 15m Class System

Over a decade of work has been conducted in the development of NASA's Hypersonic Inflatable Aerodynamic Decelerator (HIAD) deployable aeroshell technology. This effort has included multiple ground test campaigns and flight tests culminating in the HIAD project's second generation (Gen-2) aeroshell system. The HIAD project team has developed, fabricated, and tested stacked-torus inflatable structures (IS) with flexible thermal protection systems (F-TPS) ranging in diameters from 3-6m, with cone angles of 60 and 70 deg. To meet NASA and commercial near term objectives, the HIAD team must scale the current technology up to 12-15m in diameter. The HIAD project's experience in scaling the technology has reached a critical juncture. Growing from a 6m to a 15m class system will introduce many new structural and logistical challenges to an already complicated manufacturing process. Although the general architecture and key aspects of the HIAD design scale well to larger vehicles, details of the technology will need to be reevaluated and possibly redesigned for use in a 15m-class HIAD system. These include: layout and size of the structural webbing that transfers load throughout the IS, inflatable gas barrier design, torus diameter and braid construction, internal pressure and inflation line routing, adhesives used for coating and bonding, and F-TPS gore design and seam fabrication. The logistics of fabricating and testing the IS and the F-TPS also become more challenging with increased scale. Compared to the 6m aeroshell (the largest HIAD built to date), a 12m aeroshell has four times the cross-sectional area, and a 15m one has over six times the area. This means that fabrication and test procedures will need to be reexamined to account for the sheer size and weight of the aeroshell components. This will affect a variety of steps in the manufacturing process, such as: stacking the tori during assembly, stitching the structural webbing, initial inflation of tori, and stitching of F-TPS gores. Additionally, new approaches and hardware will be required for handling and ground testing of both individual tori and the fully assembled HIADs. There are also noteworthy benefits of scaling up the HIAD aeroshell to 15m-class system. Two complications in working with handmade textiles structures are the non-linearity of the materials and the role of human accuracy during fabrication. Larger, more capable, HIAD structures should see much larger operational loads, potentially bringing the structural response of the materials out of the non-linear regime and into the preferred linear response range. Also, making the reasonable assumption that the magnitude of fabrication accuracy remains constant as the structures grow, the relative effect of fabrication errors should decrease as a percentage of the textile component size. Combined, these two effects improve the predictive capability and the uniformity of the structural response for a 12-15m class HIAD. In this paper, the challenges and associated mitigation plans related to scaling up the HIAD stacked-torus aeroshell to a 15m class system will be discussed. In addition, the benefits of enlarging the structure will be further explored.

Decellerators↗

NASA NTRS2016

Manufacturing Challenges and Benefits When Scaling the HIAD Stacked-Torus Aeroshell to a 15 Meter Class System

Over a decade of work has been conducted in the development of NASA's Hypersonic Inflatable Aerodynamic Decelerator (HIAD) deployable aeroshell technology. This effort has included multiple ground test campaigns and flight tests culminating in the HIAD project's second generation (Gen-2) aeroshell system. The HIAD project team has developed, fabricated, and tested stacked-torus inflatable structures (IS) with flexible thermal protection systems (F-TPS) ranging in diameters from 3-6 meters, with cone angles of 60 and 70 degrees. To meet NASA and commercial near-term objectives, the HIAD team must scale the current technology up to 12-15 meters in diameter. Therefore, the HIAD project's experience in scaling the technology has reached a critical juncture. Growing from a 6-meter to a 15-meter class system will introduce many new structural and logistical challenges to an already complicated manufacturing process. Although the general architecture and key aspects of the HIAD design scale well to larger vehicles, details of the technology will need to be reevaluated and possibly redesigned for use in a 15-meter-class HIAD system. These include: layout and size of the structural webbing that transfers load throughout the IS, inflatable gas barrier design, torus diameter and braid construction, internal pressure and inflation line routing, adhesives used for coating and bonding, and F-TPS gore design and seam fabrication. The logistics of fabricating and testing the IS and the F-TPS also become more challenging with increased scale. Compared to the 6-meter aeroshell (the largest HIAD built to date), a 12-meter aeroshell has four times the cross-sectional area, and a 15-meter one has over six times the area. This means that fabrication and test procedures will need to be reexamined to account for the sheer size and weight of the aeroshell components. This will affect a variety of steps in the manufacturing process, such as: stacking the tori during assembly, stitching the structural webbing, initial inflation of tori, and stitching of F-TPS gores. Additionally, new approaches and hardware will be required for handling and ground testing of both individual tori and the fully assembled HIADs. There are also noteworthy benefits of scaling up the HIAD aeroshell to a 15m-class system. Two complications in working with handmade textile structures are the non-linearity of the material components and the role of human accuracy during fabrication. Larger, more capable, HIAD structures should see much larger operational loads, potentially bringing the structural response of the material components out of the non-linear regime and into the preferred linear response range. Also, making the reasonable assumption that the magnitude of fabrication accuracy remains constant as the structures grow, the relative effect of fabrication errors should decrease as a percentage of the textile component size. Combined, these two effects improve the predictive capability and the uniformity of the structural response for a 12-15-meter HIAD. In this presentation, a handful of the challenges and associated mitigation plans will be discussed, as well as an update on current manufacturing and testing that addressing these challenges.

Infatables↗

NASA NTRS2014

Technology Maturation in Preparation for the Cryogenic Propellant Storage and Transfer (CPST) Technology Demonstration Mission (TDM)

In support of its goal to find an innovative path for human space exploration, NASA embarked on the Cryogenic Propellant Storage and Transfer (CPST) Project, a Technology Demonstration Mission (TDM) to test and validate key cryogenic capabilities and technologies required for future exploration elements, opening up the architecture for large in-space cryogenic propulsion stages and propellant depots. Recognizing that key Cryogenic Fluid Management (CFM) technologies anticipated for on-orbit (flight) demonstration would benefit from additional maturation to a readiness level appropriate for infusion into the design of the flight demonstration, the NASA Headquarters Space Technology Mission Directorate (STMD) authorized funding for a one-year technology maturation phase of the CPST project. The strategy, proposed by the CPST Project Manager, focused on maturation through modeling, concept studies, and ground tests of the storage and fluid transfer of CFM technology sub-elements and components that were lower than a Technology Readiness Level (TRL) of 5. A technology maturation plan (TMP) was subsequently approved which described: the CFM technologies selected for maturation, the ground testing approach to be used, quantified success criteria of the technologies, hardware and data deliverables, and a deliverable to provide an assessment of the technology readiness after completion of the test, study or modeling activity. The specific technologies selected were grouped into five major categories: thick multilayer insulation, tank applied active thermal control, cryogenic fluid transfer, propellant gauging, and analytical tool development. Based on the success of the technology maturation efforts, the CPST project was approved to proceed to flight system development.

propellants↗

NASA NTRS2002

A Desorbed Gas Molecular Ionization Mechanism for Arcing Onset in Solar Arrays Immersed in a Low-Density Plasma

Previous experimental studies have hypothesized that the onset of Solar Array Arc (SAA) initiation in low-density space plasmas is caused by a desorbed gas molecular ionization mechanism. Indeed past investigations performed at the NASA Glenn Plasma Interaction Facility tend to not only support the desorbed gas molecular ionization mechanism, but have gone as far as identifying the crucial molecular species that must be present for molecular ion dominated process to occur. When electrical breakdown occurs at a triple junction site on a solar array panel, a quasi-neutral plasma cloud is ejected. Assuming the main component of the expelled plasma cloud by weight is due to water vapor, the fastest process available is due to HO molecules and OH(+) ions, or more succinctly, dissociative molecular-ion dominated recombination processes: H2O(+) + e(-) yields H* + OH*. Recently published spectroscopic observations of solar array arc spectra in ground tests have revealed the well-known molecular OH band (302 to 309nm), as well as the molecular SiH band (387nm peak), and the molecular CH band (432nm peak). Note that the OH band is observed in emission arcs where water vapor is present. Strong atomic lines were also observed for H(sub beta) at 486nm and H(sub alpha) at 656.3nm in prior ground testing. Independent supporting evidence of desorbed gas molecular ionization mechanisms also come from measurements of arc current pulse widths at different capacitances. We will revisit an earlier first order approximation demonstrating the dependence of arc current pulse widths on the square root of the capacitance. The simple arc current pulse width model will be then be used to estimate the temperature of the arc plasma (currently believed to be somewhere in the range of 3 to 5 eV). The current paper then seeks to extend the outlined work by including numerous vacuum chamber measurements obtained with a quadrupole mass spectrometer. A small solar array was mounted inside the vacuum chamber. A plasma source, also mounted inside the vacuum chamber, is used to simulate a low-density plasma environment. The solar array is then biased to a high negative potential and allowed to arc while a mass spectrometer is used to record the partial pressure of H2O and to track other significant changes in mass (1 to 150) AMU.

Galofaro, J.↗

NASA NTRS2016

Manufacturing Challenges and Benefits when Scaling the HIAD Stacked-Torus Aeroshell to a 15m-Class System

Over a decade of work has been conducted in the development of NASAs Hypersonic Inflatable Aerodynamic Decelerator (HIAD) deployable aeroshell technology. This effort has included multiple ground test campaigns and flight tests culminating in the HIAD projects second generation (Gen-2) aeroshell system. The HIAD project team has developed, fabricated, and tested stacked-torus inflatable structures (IS) with flexible thermal protection systems (F-TPS) ranging in diameters from 3-6m, with cone angles of 60 and 70 deg. To meet NASA and commercial near term objectives, the HIAD team must scale the current technology up to 12-15m in diameter. Therefore, the HIAD projects experience in scaling the technology has reached a critical juncture. Growing from a 6m to a 15m-class system will introduce many new structural and logistical challenges to an already complicated manufacturing process.Although the general architecture and key aspects of the HIAD design scale well to larger vehicles, details of the technology will need to be reevaluated and possibly redesigned for use in a 15m-class HIAD system. These include: layout and size of the structural webbing that transfers load throughout the IS, inflatable gas barrier design, torus diameter and braid construction, internal pressure and inflation line routing, adhesives used for coating and bonding, and F-TPS gore design and seam fabrication. The logistics of fabricating and testing the IS and the F-TPS also become more challenging with increased scale. Compared to the 6m aeroshell (the largest HIAD built to date), a 12m aeroshell has four times the cross-sectional area, and a 15m one has over six times the area. This means that fabrication and test procedures will need to be reexamined to ac-count for the sheer size and weight of the aeroshell components. This will affect a variety of steps in the manufacturing process, such as: stacking the tori during assembly, stitching the structural webbing, initial inflation of tori, and stitching of F-TPS gores. Additionally, new approaches and hardware will be required for handling and ground testing of both individual tori and the fully assembled HIADs.There are also noteworthy benefits of scaling up the HIAD aeroshell to a 15m-class system. Two complications in working with handmade textile structures are the non-linearity of the material components and the role of human accuracy during fabrication. Larger, more capable, HIAD structures should see much larger operational loads, potentially bringing the structural response of the material components out of the non-linear regime and into the preferred linear response range. Also, making the reasonable assumption that the magnitude of fabrication accuracy remains constant as the structures grow, the relative effect of fabrication errors should decrease as a percentage of the textile component size. Combined, these two effects improve the predictive capability and the uniformity of the structural response for a 12-15m HIAD.In this presentation, a handful of the challenges and associated mitigation plans will be discussed, as well as an update on current 12m aeroshell manufacturing and testing that is addressing these challenges

Infatables↗

NASA NTRS1974

An occultation satellite system for determining pressure levels in the atmosphere

A two-satellite microwave occultation system is described that will fix, as an absolute function of altitude, the pressure-temperature profile generated by a passive infrared sounder. The 300 mb pressure level is determined to within 24 m rms, assuming the temperture errors produced by the infrared sensor are not greater than 2 K rms. Error caused by water vapor in the radio path is corrected by climatological adjustments. A ground test of the proposed system is described. A microwave signal propagating between two mountain tops was found to be subject to periods of intense fading. Computer analysis of the raypath between the transmitting and receiving stations indicates that multipath and defocusing were responsible for this fading. It is unlikely that an operational pressure-reference-level system will be subject to the deep fades observed in the ground test, because the phenomena are associated with lower altitudes than the closest approach altitude of an occultation-system raypath.

Morrison, A. R.↗

NASA NTRSDate not supplied

Continuation of Solar Absorptivity Degradation of Spacecraft Materials Due to UV and Charged Particles in the Gateway Environment

Gateway is a lunar orbiting platform that will be assembled in space similar to the International Space Station. Modules will arrive following different transit orbits, some of which have an extended duration in the Van Allen Belts. This means the spacecraft is exposed to significant charged particle radiation, after which it must operate in cislunar space for 15 years. Charged particle and UV degradation of potential Gateway materials was identified as a knowledge gap by the Passive Thermal Control System (PTCS) Group as end-of-life solar absorptivity greatly impacts thermal performance. Multiple rounds of ground testing have been completed at NASA Goddard Space Flight Center in which potential Gateway materials were subjected to expected worst-case transit charged particle fluence (9.86x1015 protons/cm2 at 2.5 keV and 3.1x1016 electrons/cm2 at 10 keV) and up to 5093 Equivalent Solar Hours (ESH) of UV exposure. This report builds on previous data from Zinecker et al. 2022 which presented the first round of ground testing. Additional materials have been tested and solar absorptivity degradation along with lessons learned are shared. Comparisons to previous results are also presented.

Brandon Hoffmann↗

NASA NTRS2007

Lunar Dust on Heat Rejection System Surfaces: Problems and Prospects

Heat rejection from power systems will be necessary for human and robotic activity on the lunar surface. Functional operation of such heat rejection systems is at risk of degradation as a consequence of dust accumulation. The Apollo astronauts encountered marked degradation of performance in heat rejection systems for the lunar roving vehicle, science packages, and other components. Although ground testing of dust mitigation concepts in support of the Apollo mission identified mitigation tools, the brush concept adopted by the Apollo astronauts proved essentially ineffective. A better understanding of the issues associated with the impact of lunar dust on the functional performance of heat rejection systems and its removal is needed as planning gets underway for human and robotic missions to the Moon. Renewed emphasis must also be placed on ground testing of pristine and dust-covered heat rejection system surfaces to quantify degradation and address mitigation concepts. This paper presents a review of the degradation in performance of heat rejection systems encountered on the lunar surface to-date, and will discuss current activities underway to evaluate the durability of candidate heat rejection system surfaces and current dust mitigation concepts.

Gaier, James R.↗

NASA NTRS2013

Cryogenic Propellant Storage and Transfer (CPST) Technology Maturation: Establishing a Foundation for a Technology Demonstration Mission (TDM)

As part of U.S. National Space Policy, NASA is seeking an innovative path for human space exploration, which strengthens the capability to extend human and robotic presence throughout the solar system. NASA is laying the groundwork to enable humans to safely reach multiple potential destinations, including asteroids, Lagrange points, the Moon and Mars. In support of this, NASA is embarking on the Technology Demonstration Mission Cryogenic Propellant Storage and Transfer (TDM CPST) Project to test and validate key cryogenic capabilities and technologies required for future exploration elements, opening up the architecture for large cryogenic propulsion stages (CPS) and propellant depots. The TDM CPST project will provide an on-orbit demonstration of the capability to store, transfer, and measure cryogenic propellants for a duration which is relevant to enable long term human space exploration missions beyond low Earth orbit (LEO). Recognizing that key cryogenic fluid management technologies anticipated for on-orbit (flight) demonstration needed to be matured to a readiness level appropriate for infusion into the design of the flight demonstration, the NASA Headquarters Space Technology Mission Directorate authorized funding for a one-year (FY12) ground based technology maturation program. The strategy, proposed by the CPST Project Manager, focused on maturation through modeling, studies, and ground tests of the storage and fluid transfer Cryogenic Fluid Management (CFM) technology sub-elements and components that were not already at a Technology Readiness Level (TRL) of 5. A technology maturation plan (TMP) was subsequently approved which described: the CFM technologies selected for maturation, the ground testing approach to be used, quantified success criteria of the technologies, hardware and data deliverables, and a deliverable to provide an assessment of the technology readiness after completion of the test, study or modeling activity. This paper will present the testing, studies, and modeling that occurred in FY12 to mature cryogenic fluid management technologies for propellant storage, transfer, and supply, to examine extensibility to full scale, long duration missions, and to develop and validate analytical models. Finally, the paper will briefly describe an upcoming test to demonstrate Liquid Oxygen (LO2) Zero Boil- Off (ZBO).

Doherty, Michael P.↗

NASA NTRS2014

Cryogenic Propellant Storage and Transfer (CPST) Technology Maturation: Establishing a Foundation for a Technology Demonstration Mission (TDM)

Doherty, Michael P.↗

NASA NTRS1952

The Calculated and Measured Performance Characteristics of a Heated-Wire Liquid-Water-Content Meter for Measuring Icing Severity

Ground tests have been made of an instrument which, when assembled in a more compact form for flight installation, could be used to obtain statistical flight data on the liquid-water content of icing clouds and to provide an indication of icing severity. The sensing element of the instrument consists of an electrically heated wire which is mounted in the air stream. The degree of cooling of the wire resulting from evaporation of the impinging water droplets is a measure. of the liquid-water content of the cloud. Determination of the value of the liquid-water content from the wire temperature at any instant requires a knowledge of the airspeed, altitude, and air temperature. An analysis was made of the temperature response of a heated wire exposed to an air stream containing water drops. Comparisons were made of the liquid-water content as measured with several heated wires and absorbent cylinders in an artificially produced cloud. For one of the wires, comparative tests were made with a rotating-disk icing-rate meter in an icing wind tunnel. From the test results, it was shown that an instrument for measuring the concentration of liquid water in an air stream can be built using an electrically heated wire of known temperatureresistance characteristics, and that the performance of such a device can be predicted using appropriate theory. Although an instrument in a form suitable for gathering statistical data in flight was not built, the practicability of constructing such an instrument was illustrated. The ground-test results indicated that a flight heated-wire instrument would be simple and durable, would respond rapidly to variations in liquid-water content, and could be used for the measurement of water content in clouds which are above freezing temperature, as well as in icing clouds.

Neel, Carr B.↗

NASA NTRS2007

Ground-Based Testing of TiB2 and Al2O3/TiB2 Response to Space Environment

Two materials, titanium diboride and an alumina/titanium diboride composite, exhibit characteristics favorable for use in multiple space applications. These characteristics include low mass (4.52 gm/cc), high strain rate impact resistance, high temperature use (3000oC M.P.), thermal and electrical conductivity, thermal shock resistance, and high visible-range reflectivity. Additionally, the presence of boron in these materials gives them the potential to shield against neutron radiation as well as charged radiation. These materials are flying on MISSE 6 to assess material changes resulting from exposure to the space environment. This study provides a preliminary, ground-based examination of these materials' interactions with individual components of the space environment, in particular atomic oxygen (AO) and neutron radiation, in order to better predict and understand post-flight results. Individual specimens are exposed to ground state AO and surface oxidation is measured. Equivalent exposures of up to 13 months show no rapid oxidation, however evidence indicates some surface oxidation occurring. Other samples are placed near a polyethylene moderated, one Ci Am/Be neutron source to determine their shielding capability. Comparisons between exposed and shielded indium foil, which is activated by transmitted neutrons, measure each material's ability to shield neutrons. Preliminary results indicate a significant shielding benefit provided by both materials.

Jefferies, Sharon A.↗

NASA NTRS1990

Solar array arcing in plasmas

Solar cells in space plasma conditions are known to arc into the plasma when the interconnects are at a negative potential of a few hundred volts, relative to plasma potential. For cells with silver-coated interconnects, a threshold voltage for arcing exists at about -230 V, as found in both ground and LEO experiments. The arc rate beyond the threshold voltage depends nearly linearly on plasma density, but has a strong power-law dependence on voltage, such that for small increments in operating voltage there is a large increment in arc rate. The arcs generate broadband radio interference and visible light. In ground tests, interconnects have been damaged by arcs in cells having insufficient isolation from a source of high current. Models for the arcs are highly dependent on the choice of interconnect conductor material exposed to the plasma and possibly on the geometry and choice of adjacent insulator material. Finally, new technology solar cells use copper for the cell interconnects, a material which may have a lower arcing threshold voltage than silver. It is expected, from ground tests of simulated solar cells, that any junction of conductor and insulator exposed to space plasma conditions will arc into the plasma at a few hundred volts negative potential, relative to the local plasma.

Ferguson, Dale C.↗

NASA NTRS2001

Satellite Contamination and Materials Outgassing Knowledgebase - An Interactive Database Reference

The goal of this program is to collect at one site much of the knowledge accumulated about the outgassing properties of aerospace materials based on ground testing, the effects of this outgassing observed on spacecraft in flight, and the broader contamination environment measured by instruments on-orbit. We believe that this Web site will help move contamination a step forward, away from anecdotal folklore toward engineering discipline. Our hope is that once operational, this site will form a nucleus for information exchange, that users will not only take information from our knowledge base, but also provide new information from ground testing and space missions, expanding and increasing the value of this site to all. We urge Government and industry users to endorse this approach that will reduce redundant testing, reduce unnecessary delays, permit uniform comparisons, and permit informed decisions.

Green, D. B.↗

NASA NTRS1990

Laboratory Simulation of the Effect of Rocket Thrust on a Precessing Space Vehicle

Ground tests of solid propellant rocket motors have shown that metal-containing propellants produce various amounts of slag (primarily aluminum oxide) which is trapped in the motor case, causing a loss of specific impulse. Although not yet definitely established, the presence of a liquid pool of slag also may contribute to nutational instabilities that have been observed with certain spin-stabilized, upper-stage vehicles. Because of the rocket's axial acceleration, absent in the ground tests, estimates of in-flight slag mass have been very uncertain. Yet such estimates are needed to determine the magnitude of the control authority of the systems required for eliminating the instability. A test rig with an eccentrically mounted hemispherical bowl was designed and built which incorporates a follower force that properly aligns the thrust vector along the axis of spin. A program that computes the motion of a point mass in the spinning and precessing bowl was written. Using various RPMs, friction factors, and initial starting conditions, plots were generated showing the trace of the point mass around the inside of the fuel tank. The apparatus will incorporate future design features such as a variable nutation angle and a film height measuring instrument. Data obtained on the nutational instability characteristics will be used to determine order of magnitude estimates of control authority needed to minimize the sloshing effect.

Alvarez, Oscar↗

NASA NTRS1990

Laboratory simulation of the rocket motor thrust as a follower force

Ground tests of solid propellant rocket motors have shown that metal-containing propellants produce various amounts of slag (primarily aluminum oxide), which is trapped in the motor case causing a loss of specific impulse. Although not yet definitely established, the presence of a liquid pool of slag also may contribute to nutational instabilities that have been observed with certain spin-stabilized, upper-stage vehicles. Because of the rocket's axial acceleration - absent in the ground tests - estimates of in-flight slag mass have been very uncertain. Yet such estimates are needed to determine the magnitude of the control authority of the systems required for eliminating the instability. A test rig with an eccentrically mounted hemispherical bowl was designed and built that incorporates a follower force that properly aligns the thrust vector along the axis of spin. A program that computes the motion of a point mass in the spinning and precessing bowl was written. Using various rpm, friction factors, and initial starting conditions, plots were generated showing the trace of the point mass around the inside of the fuel tank. The apparatus will be used extensively during the 1990 to 1991 academic year and incorporate future design features such as a variable nutation angle and a film height measuring instrument. Data obtained on the nutational instability characteristics will be used to determine order-of-magnitude estimates of control authority needed to minimize the sloshing effect.

Source record↗