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At least 127 records · Page 7
All sourcesNASA NTRSDOE OSTI
NASA NTRS2009

Effect of Changing Weight and Mass on Human Performance in a Lunar Prototype Spacesuit

Physical effort, compensation, and controllability in a spacesuit can be affected by suit mass and gravity level. Because of limitations in certain reduced-gravity simulators and the finite selection of lunar prototype suits, it is difficult to ascertain how a change in suit mass affects suited human performance. One method of simulating a change in mass is to vary the total gravity-adjusted weight (TGAW), which is defined as the sum of the suit mass and subject mass, multiplied by the gravity level. PURPOSE: To determine if two methods of changing TGAW during parabolic flight - changing suit mass or gravity level - affect subjective ratings of suited human performance equally. METHODS: A custom weight support structure was connected to a lunar prototype spacesuit, allowing the addition of mass to the suit while maintaining a near-constant center of mass. In the varied-weight (VW) series, suit mass (120 kg) was constant at 0.1 G, 0.17 G, and 0.3 G, yielding TGAWs of 196, 333, and 588 N, assuming an 80-kg subject. In the varied-mass (VM) series, gravity level was constant at 0.17 G and suit mass was 89, 120, and 181 kg, yielding TGAWs of 282, 333, and 435 N. The 333 N condition was common to both series. Direct comparison was not possible due to limited adjustability of suit mass and limited options for parabolic profiles. Five astronaut subjects (80.3 11.8 kg) completed 4 different tasks (walk, bag pickup, lunge, and shoveling) in all conditions and provided ratings of perceived exertion (RPE) and the gravity compensation and performance scale (GCPS) upon completion of each task. RESULTS: Where VM and VW series overlapped, RPE and GCPS trendlines were similar. Mean RPE and GCPS at 333 N was 8.4 and 3.7. Mean RPE and GCPS for VM was 7.8 and 3.8 for 282 N and 9.8 and 4.1 for 435 N. Extrapolation of the VM trend to match VW TGAWs 196 and 588 N predicts an RPE of 6.5 and 12.3 and GCPS of 4.4 and 5.9, whereas the measured VW values for RPE were 8.1 and 9.8 and GCPS were 4.4 and 3.7. CONCLUSION: Modeling a change in suit mass by altering weight alone may be an adequate simulation through a limited range when looking at gross metrics of subjective suited human performance. Whether altering weight alone will be sufficient for more precise metrics of human performance, and across a wider range of activities, still needs further study.

Norcross, Jason R.↗
NASA NTRS2010

Effect of Changing Weight and Mass on Human Performance in a Lunar Prototype Spacesuit

Physical effort, compensation, and controllability in a spacesuit can be affected by suit mass and gravity level. Because of limitations in certain reduced-gravity simulators and the finite selection of lunar prototype suits, it is difficult to ascertain how a change in suit mass affects suited human performance. One method of simulating a change in mass is to vary the total gravity-adjusted weight (TGAW), which is defined as the sum of the suit mass and subject mass, multiplied by the gravity level. PURPOSE: To determine if two methods of changing TGAW during parabolic flight - changing suit mass or gravity level - affect subjective ratings of suited human performance equally.METHODS: A custom weight support structure was connected to a lunar prototype spacesuit, allowing the addition of mass to the suit while maintaining a near-constant center of mass. In the varied-weight (VW) series, suit mass (120 kg) was constant at 0.1-g, 0.17-g, and 0.3-g, yielding TGAWs of 196, 333, and 588 N, assuming an 80-kg subject. In the varied-mass (VM) series, gravity level was constant at 0.17-g and suit mass was 89, 120, and 181 kg, yielding TGAWs of 282, 333, and 435 N. The 333 N condition was common to both series. Direct comparison was not possible due to limited adjustability of suit mass and limited options for parabolic profiles. Five astronaut subjects (80.311.8 kg) completed 4 different tasks (walk, bag pickup, lunge, and shoveling) in all conditions and provided ratings of perceived exertion (RPE) and the gravity compensation and performance scale (GCPS) upon completion of each task. RESULTS: Where VM and VW series overlapped, RPE and GCPS trend lines were similar. Mean RPE and GCPS at 333 N was 8.4 and 3.7. Mean RPE and GCPS for VM was 7.8 and 3.8 for 282 N and 9.8 and 4.1 for 435 N. Extrapolation of the VM trend to match VW TGAWs 196 and 588 N predicts an RPE of 6.5 and 12.3 and GCPS of 4.4 and 5.9, whereas the measured VW values for RPE were 8.1 and 9.8 and GCPS were 4.4 and 3.7. CONCLUSION: Modeling a change in suit mass by altering weight alone may be an adequate simulation through a limited range when looking at gross metrics of subjective suited human performance. Whether altering weight alone will be sufficient for more precise metrics of human performance, and across a wider range of activities, still needs further study

Chappell, Steve↗
NASA NTRS2007

Space Technology 7 : Micropropulsion and Mass Distribution

The NASA New Millennium Program Space Technology 7 (ST7) project will validate technology for precision spacecraft control. The ST7 disturbance reduction system (DRS) will contain new micropropulsion technology to be flown as part of the European Space Agency's LISA (laser interferometer space antenna) Pathfinder project. After launch into a low Earth orbit in early 2010, the LISA Pathfinder spacecraft will be maneuvered to a halo orbit about the Earth-Sun LI Lagrange point for operations. The DRS will control the position of the spacecraft relative to a reference to an accuracy of one nanometer over time scales of several thousand seconds. To perform the control the spacecraft will use a new colloid thruster technology. The thrusters will operate over the range of 5 to 30 micro-Newtons with precision of 0.1 micro-Newton. The thrust will be generated by using a high electric field to extract charged droplets of a conducting colloid fluid and accelerating them with a precisely adjustable voltage. The control position reference will be provided by the European LISA Technology Package, which will include two nearly free-floating test masses. The test mass position and attitude will be sensed and adjusted using electrostatic capacitance bridges. The DRS will control the spacecraft position with respect to one test mass while minimizing disturbances on the second test mass. The dynamic control system will cover eighteen degrees of freedom, six for each of the test masses and six for the spacecraft. In the absence of other disturbances, the test masses will slowly gravitate toward local concentrations of spacecraft mass. The test mass acceleration must be minimized to maintain the acceleration of the enclosing drag-free spacecraft within the control authority of the micropropulsion system. Therefore, test mass acceleration must be predicted by accurate measurement of mass distribution, then offset by the placement of specially shaped balance masses near each test mass. The - acceleration is characterized by calculating the gravitational effect of over ten million modeled points of a nearly 500-kg spacecraft. This paper provides an overview of the mission technology and the process of precision mass modeling of the DRS equipment.

ST-7↗
NASA NTRS2014

3D-HST + CANDELS: the Evolution of the Galaxy Size-mass Distribution Since Z=3

Spectroscopic and photometric redshifts, stellar mass estimates, and rest-frame colors from the 3D-HST survey are combined with structural parameter measurements from CANDELS imaging to determine the galaxy size-mass distribution over the redshift (z) range 0 < z < 3. Separating early- and late-type galaxies on the basis of star-formation activity, we confirm that early-type galaxies are on average smaller than late-type galaxies at all redshifts, and find a significantly different rate of average size evolution at fixed galaxy mass, with fast evolution for the early-type population, effective radius is in proportion to (1 + z) (sup −1.48), and moderate evolution for the late-type population, effective radius is in proportion to (1 + z) (sup −0.75). The large sample size and dynamic range in both galaxy mass and redshift, in combination with the high fidelity of our measurements due to the extensive use of spectroscopic data, not only fortify previous results, but also enable us to probe beyond simple average galaxy size measurements. At all redshifts the slope of the size-mass relation is shallow, effective radius in proportion to mass of a black hole (sup 0.22), for late-type galaxies with stellar mass > 3 x 10 (sup 9) solar masses, and steep, effective radius in proportion to mass of a black hole (sup 0.75), for early-type galaxies with stellar mass > 2 x 10 (sup 10) solar masses. The intrinsic scatter is approximately or less than 0.2 decimal exponents for all galaxy types and redshifts. For late-type galaxies, the logarithmic size distribution is not symmetric, but skewed toward small sizes: at all redshifts and masses a tail of small late-type galaxies exists that overlaps in size with the early-type galaxy population. The number density of massive (approximately 10 (sup 11) solar masses), compact (effective radius less than 2 kiloparsecs) early-type galaxies increases from z = 3 to z = 1.5 − 2 and then strongly decreases at later cosmic times.

EVOLUTION↗
NASA NTRS1979

Stellar evolution at high mass including the effect of a stellar wind

The effect of a stellar wind on the evolution of stars in the mass range from 15 to 120 solar masses is investigated. All the stellar models are constructed with the use of Cox-Stewart opacities. Four possible cases of mass loss are considered: (1) no mass loss at all; (2) substantial mass loss from stars in all stages of evolution; (3) heavy mass loss from red supergiants only; and (4) sudden and very heavy mass loss from luminous yellow supergiants. The assumption of mass loss during the main-sequence phase of evolution is found to lead to a lowering of the luminosity and, unless the mass loss is extremely heavy, of the effective temperature as well. A comparison of the adopted mass-loss rates with observed rates suggests that stellar winds are probably not an important factor in the evolution of main-sequence stars and supergiants unless the initial masses are greater than about 30 solar masses.

Stothers, R.↗
NASA NTRS1986

The Cepheid mass problem and Cepheid binaries

Existing mass determinations for Cepheids with different periods are examined. Wesselink masses are independent of the adopted distance scale. For short periods (less than 6 days) they follow the sequence of evolutionary masses. For periods longer than 10 days they are lower by up to a factor of 2. The lower mass branch joins up with the bump masses. The new pulsational masses agree with the Wesselink masses for periods longer than 6 days. Cepheid masses determined by means of their giant companions also agree with the Wesselink masses and the new pulsational masses. While the error bars are large, the derived dynamical masses determined for S Mus and V636 Sco also agree with the low Wesselink and giant companion masses.

Bohm-Vitense, E.↗
NASA NTRS1993

Luminosity functions for very low mass stars and brown dwarfs

A theoretical investigation of the luminosity function for low-mass objects to constrain the stellar initial mass function at the low-mass end is reported. The ways in which luminosity functions for low-mass stars are affected by star formation histories, brown dwarf and premain-sequence cooling rates and main-sequence mass luminosity relations, and the IMF are examined. Cooling rates and the mass-luminosity relation are determined through a new series of evolutionary calculations for very low mass stars and brown dwarfs in the range 0.05-0.50 solar mass. Model luminosity functions are constructed for specific comparison with the results of four recent observational surveys. The likelihood that the stellar mass function in the solar neighborhood is increasing at masses near the bottom of the main sequence and perhaps at lower masses is confirmed. In the most optimistic case, brown dwarfs contribute half of the local missing disk mass. The actual contribution is likely to be considerably less.

Laughlin, Gregory↗
NASA NTRS1993

The evolution of massive stars including mass loss - Presupernova models and explosion

The evolution of massive stars of 35, 40, 60, and 85 solar masses is followed through all stages of nuclear burning to the point of Fe core collapse. Critical nuclear reaction and mass-loss rates are varied. Efficient mass loss during the Wolf-Rayet (WR) stage is likely to lead to final masses as small as 4 solar masses. For a reasonable parameterization of the mass loss, there may be convergence of all WR stars, both single and in binaries, to a narrow band of small final masses. Our representative model, a 4.25 solar-mass WR presupernova derived from a 60 solar mass star, is followed through a simulated explosion, and its explosive nucleosynthesis and light curve are determined. Its properties are similar to those observed in Type Ib supernovae. The effects of the initial mass and mass loss on the presupernova structure of small mass WR models is also explored. Important properties of the presupernova star and its explosion can only be obtained by following the complete evolution starting on the main sequence.

Woosley, S. E.↗
NASA NTRS2009

Evidence for an Intermediate Mass Black Hole in NGC 5408 X-1

We report the discovery with XMM-Newton of correlated spectral and timing behavior in the ultraluminous X-ray source (ULX) NGC 5408 X-1. An approx. 100 ksec pointing with XMM/Newton obtained in January, 2008 reveals a strong 10 mHz QPO in the > 1 keV flux, as well as flat-topped, band limited noise breaking to a power law. The energy spectrum is again dominated by two components, a 0.16 keV thermal disk and a power-law with an index of approx. 2.5. These new measurements, combined with results from our previous January 2006 pointing in which we first detected QPOs, show for the first time in a ULX a pattern of spectral and temporal correlations strongly analogous to that seen in Galactic black hole sources, but at much higher X-ray luminosity and longer characteristic time-scales. We find that the QPO frequency is proportional to the inferred disk flux, while the QPO and broad-band noise amplitude (root mean squared, rms) are inversely proportional to the disk flux. Assuming that QPO frequency scales inversely with black hole mass at a given power-law spectral index we derive mass estimates using the observed QPO frequency - spectral index relations from five stellar-mass black hole systems with dynamical mass constraints. The results from all sources are consistent with a mass range for NGC 5408 X-1 from 1000 - 9000 Stellar mass. We argue that these are conservative limits, and a more likely range is from 2000 - 5000 Stellar mass. Moreover, the recent relation from Gierlinski et al. that relates black hole mass to the strength of variability at high frequencies (above the break in the power spectrum), and the variability plane results of McHardy et al. and Koerding et al., are also suggestive of such a. high mass for NGC 5408 X-1. Importantly, none of the above estimates appears consistent with a black hole mass less than approx. 1000 Stellar mass for NGC 5408 X-1. We argue that these new findings strongly support the conclusion that NGC 5408 X-1 harbors an intermediate mass black hole.

Strohmayer, Tod E.↗
NASA NTRS2009

Modeling and Simulation of Variable Mass, Flexible Structures

The advent of the new Ares I launch vehicle has highlighted the need for advanced dynamic analysis tools for variable mass, flexible structures. This system is composed of interconnected flexible stages or components undergoing rapid mass depletion through the consumption of solid or liquid propellant. In addition to large rigid body configuration changes, the system simultaneously experiences elastic deformations. In most applications, the elastic deformations are compatible with linear strain-displacement relationships and are typically modeled using the assumed modes technique. The deformation of the system is approximated through the linear combination of the products of spatial shape functions and generalized time coordinates. Spatial shape functions are traditionally composed of normal mode shapes of the system or even constraint modes and static deformations derived from finite element models of the system. Equations of motion for systems undergoing coupled large rigid body motion and elastic deformation have previously been derived through a number of techniques [1]. However, in these derivations, the mode shapes or spatial shape functions of the system components were considered constant. But with the Ares I vehicle, the structural characteristics of the system are changing with the mass of the system. Previous approaches to solving this problem involve periodic updates to the spatial shape functions or interpolation between shape functions based on system mass or elapsed mission time. These solutions often introduce misleading or even unstable numerical transients into the system. Plus, interpolation on a shape function is not intuitive. This paper presents an approach in which the shape functions are held constant and operate on the changing mass and stiffness matrices of the vehicle components. Each vehicle stage or component finite element model is broken into dry structure and propellant models. A library of propellant models is used to describe the distribution of mass in the fuel tank or Solid Rocket Booster (SRB) case for various propellant levels. Based on the mass consumed by the liquid engine or SRB, the appropriate propellant model is coupled with the dry structure model for the stage. Then using vehicle configuration data, the integrated vehicle model is assembled and operated on by the constant system shape functions. The system mode shapes and frequencies can then be computed from the resulting generalized mass and stiffness matrices for that mass configuration. The rigid body mass properties of the vehicle are derived from the integrated vehicle model. The coupling terms between the vehicle rigid body motion and elastic deformation are also updated from the constant system shape functions and the integrated vehicle model. This approach was first used to analyze variable mass spinning beams and then prototyped into a generic dynamics simulation engine. The resulting code was tested against Crew Launch Vehicle (CLV-)class problems worked in the TREETOPS simulation package and by Wilson [2]. The Ares I System Integration Laboratory (SIL) is currently being developed at the Marshall Space Flight Center (MSFC) to test vehicle avionics hardware and software in a hardware-in-the-loop (HWIL) environment and certify that the integrated system is prepared for flight. The Ares I SIL utilizes the Ares Real-Time Environment for Modeling, Integration, and Simulation (ARTEMIS) tool to simulate the launch vehicle and stimulate avionics hardware. Due to the presence of vehicle control system filters and the thrust oscillation suppression system, which are tuned to the structural characteristics of the vehicle, ARTEMIS must incorporate accurate structural models of the Ares I launch vehicle. The ARTEMIS core dynamics simulation models the highly coupled nature of the vehicle flexible body dynamics, propellant slosh, and vehicle nozzle inertia effects combined with mass and flexible body properties that vary significant with time during the flight. All forces that act on the vehicle during flight must be simulated, including deflected engine thrust force, spatially distributed aerodynamic forces, gravity, and reaction control jet thrust forces. These forces are used to excite an integrated flexible vehicle, slosh, and nozzle dynamics model for the vehicle stack that simulates large rigid body translations and rotations along with small elastic deformations. Highly effective matrix math operations on a distributed, threaded high-performance simulation node allow ARTEMIS to retain up to 30 modes of flex for real-time simulation. Stage elements that separate from the stack during flight are propagated as independent rigid six degrees of freedom (6DOF) bodies. This paper will present the formulation of the resulting equations of motion, solutions to example problems, and describe the resulting dynamics simulation engine within ARTEMIS.

Tobbe, Patrick A.↗
NASA NTRS2011

Operational Parameters, Considerations, and Design Decisions for Resource-Constrained Ion Trap Mass Spectrometers

Mass spectrometers are increasingly finding applications in new and unique areas, often in situations where key operational resources (i.e. power, weight and size) are limited. One such example is the Mars Organic Molecule Analyzer (MOMA). This instrument is a joint venture between NASA and the European Space Agency (ESA) to develop an ion trap mass spectrometer for chemical analysis on Mars. The constraints on such an instrument are significant as are the performance requirements. While the ideal operating parameters for an ion trap are generally well characterized, methods to maintain analytical performance with limited power and system weight need to be investigated and tested. Methods Experiments have been performed on two custom ion trap mass spectrometers developed as prototypes for the MOMA instrument. This hardware consists of quadrupole ion trap electrodes that are 70% the size of common commercial instrumentation. The trapping RF voltage is created with a custom tank circuit that can be tuned over a range of RF frequencies and is driven using laboratory supplies and amplifiers. The entire instrument is controlled with custom Lab VIEW software that allows a high degree of flexibility in the definition of the scan function defining the ion trap experiment. Ions are typically generated via an internal electron ionization source, however, a laser desorption source is also in development for analysis of larger intact molecules. Preliminary Data The main goals in this work have been to reduce the power required to generate the radio frequency trapping field used in an ion trap mass spectrometer. Generally minimizing the power will also reduce the volume and mass of the electronics to support the instrument. In order to achieve optimum performance, commercial instruments typically utilize RF frequencies in the 1 MHz range. Without much concern for power usage, they simply generate the voltage required to access the mass range of interest. In order to reduce the required RF voltage (and power), operation of the ion trap at lower RF frequencies has been investigated. Surprisingly, the performance of the instrument has only been slightly degraded at RF frequencies all the way down to 500 kHz. Mass resolution is relatively stable to this point and depending on the resonant ejection point used, the peak intensity is also quite stable. To date only masses up to m/z 200 have been fully investigated, however, additional studies are planned to verify the performance with higher mass ions. The lower frequency and voltage should reduce the pseudo potential well depth, eventually affecting the trapping efficiency of the instrument -- effect that could manifest itself in significantly limiting the mass range of trapped ions. Other methods to reduce the RF power while maintaining analytical performance are also under investigation. This includes ion ejection at lower q(sub z) values to access a given mass with a lower RF voltage. The loss of mass resolution at lower q(sub eject) points has been measured and current work is underway to leverage scan speed and the use of non-linear resonances in order to counter this trend. The overall trap performance under this range of operating conditions will be presented with a goal of identifying what trade-offs are acceptable.

Danell, Ryan M.↗
NASA NTRS2016

Optimization of NTP System Truss to Reduce Radiation Shield Mass

The benefits of nuclear thermal propulsion are numerous and relevant to the current NASA mission goals involving but not limited to the crewed missions to mars and the moon. They do however also present new and unique challenges to the design and logistics of launching/operating spacecraft. One of these challenges, relevant to this discussion, is the significant mass of the shielding which is required to ensure an acceptable radiation environment for the spacecraft and crew. Efforts to reduce shielding mass are difficult to accomplish from material and geometric design points of the shield itself, however by increasing the distance between the nuclear engines and the main body of the spacecraft the required mass of the shielding is lessened considerably. The mass can be reduced significantly per unit length, though any additional mass added by the structure to create this distance serves to offset those savings, thus the design of a lightweight structure is ideal. The challenges of designing the truss are bounded by several limiting factors including; the loading conditions, the capabilities of the launch vehicle, and achieving the ideal truss length when factoring for the overall mass reduced. Determining the overall set of mass values for a truss of varying length is difficult since to maintain an optimally designed truss the geometry of the truss or its members must change. Thus the relation between truss mass and length for these loading scenarios is not linear, and instead has relation determined by the truss design. In order to establish a mass versus length trend for various truss designs to compare with the mass saved from the shield versus length, optimization software was used to find optimal geometric properties that still met the design requirements at established lengths. By solving for optimal designs at various lengths, mass trends could be determined. The initial design findings show a clear benefit to extending the engines as far from the main structure of the spacecraft as the launch vehicle's payload volume would allow when comparing mass savings verse the additional structure.

Scharber, Luke L.↗
NASA NTRS2022

Separating Propulsive Mass and Energy for Space Applications

Initially and traditionally, space access and in-space propulsion utilized combustion and expulsion of chemical “fuels” carried on board, with performance governed by the rocket equation. These chemicals produced energy, and after combustion, constituted the propulsive mass/momentum. The best such chemicals in terms of Isp that are deemed safe engineering/mission wise are hydrogen and oxygen, producing some 450 seconds of Isp. There are more reactive chemicals, such as fluorine, which have higher Isp, but also have serious safety issues. Propulsion in atmospheres can ingest atmospheric constituents (aka “airbreathing propulsion”) which provide additional propulsive mass when heated and a component of the combustion/energy generation process. This utilization of non-stored/carried propulsive mass provides partial separation of propulsive mass and energy and produces higher Isp. The other approach to separating propulsive mass and energy is to, either on board or added from offboard, supply additional energy, such as from nuclear or solar sources. This is the approach for fission nuclear thermal and electric propulsion, which can provide an Isp in excess of 800 seconds of Isp. Other processes or in addition to thermal expansion, such as electro-magnetics, can be employed to increase exit velocity and Isp. Separating propulsive mass and energy for space faring is commonly referred to as a means to circumvent the rocket equation and is capable of producing major benefits for the development of commercial deep space and space faring in general including affordable fast transits to mitigate the human health impacts of galactic cosmic radiation (GCR) and microG. The key to higher than chemical Isp, beyond H2-O2, and beyond the radiated energy and systems limitations of solar for space faring, is a light weight, high energy density source, either on board or via utilization of energy beaming to the vehicle. Traditional energy sources include chemical, heat, electrics, mechanical, photons, and nuclear. Propulsive mass can be carried on board, sourced beyond the surface of Earth via in-situ resource utilization (ISRU), and includes harvesting from atmospheres. High Isp via electromagnetic related propulsive mass acceleration requires sufficient ionization and conductivity. The major metrics for space propulsion are costs, safety, Isp, weight, and thrust level, the latter dependent upon mission requirements. High thrust for human missions is needed to reduce time exposed to radiation and microG and high thrust is required for space access. Costs of space access are reducing via reusability, printing manufacture, and robotization of manufacturing and operation. Chemical rockets provide high thrust from the expansion of the heated mass constituents at high mass flow. The other high thrust propulsion approach is magnetohydrodynamics (MHD), which, in addition to high thrust, has a high Isp of over 2,000 seconds of Isp. The VASIMIR engine offers some 5,000 seconds of Isp at high thrust [ref. 3]. Electric propulsion cycles are capable of Isp much higher than that but at low thrust levels using available energy sources. With cost as a major metric, reusable rockets and improved manufacturing and operations are rapidly greatly lowering the costs of space access, which could provide/supply in space fuel depots and affordable chemical propulsion for the desired human fast transits (e.g., some 200-day round trips to Mars). The other option for fast transits is VASIMIR, given a nuclear on-board energy source that has the requisite many megawatts of power and an alpha, kgs of weight/kW of energy produced, on the order of one (i.e., a light weight, high energy density energy source). The purpose of this report is to examine the options beyond traditional rocket engines, where propulsive mass and energy are combined, specifically the separation of propulsive mass and energy. This report considers the spectrum of advanced energetics, sources of propulsive mass, conductivity enhancement approaches, energy beaming possibilities, and candidate propulsion cycles. Suggestions are made for various combinatorial, system level, beyond traditional combustion rocket, space propulsion approaches for human deep space missions given the changing conditions of increased knowledge of deep space resources for ISRU, revolutionary energetics, and technology advancements writ large.

Dennis M. Bushnell↗
NASA NTRS2022

Four Sub-Jovian-Mass Planets Detected by High-Cadence Microlensing Surveys

Aims. With the aim of finding short-term planetary signals, we investigated the data collected from current high-cadence microlensing surveys. Methods. From this investigation, we found four planetary systems with low planet-to-host mass ratios, including OGLE-2017-BLG- 1691L, KMT-2021-BLG-0320L, KMT-2021-BLG-1303L, and KMT-2021-BLG-1554L. Despite the short durations, ranging from a few hours to a couple of days, the planetary signals were clearly detected by the combined data of the lensing surveys. We found that three of the planetary systems have mass ratios on the order of 10−4 and the other has a mass ratio that is slightly greater than 10−3. Results. The estimated masses indicate that all discovered planets have sub-Jovian masses. The planet masses of KMT-2021-BLG- 0320Lb, KMT-2021-BLG-1303Lb, and KMT-2021-BLG-1554Lb correspond to ∼0.10, ∼0.38, and ∼0.12 times the mass of the Jupiter, and the mass of OGLE-2017-BLG-1691Lb corresponds to that of the Uranus. The estimated mass of the planet host KMT-2021-BLG- 1554L, M host ∼ 0.08 M ⊙ , corresponds to the boundary between a star and a brown dwarf. Besides this system, the host stars of the other planetary systems are low-mass stars with masses in the range of ∼[0.3–0.6] M ⊙ . The discoveries of the planets fully demonstrate the capability of the current high-cadence microlensing surveys in detecting low-mass planets.

planets and satellites: detection↗
NASA NTRS1980

Echocardiographic left ventricular masses in distance runners and weight lifters

The relationships of different forms of exercise training to left ventricular mass and body mass are investigated by echocardiographic studies of weight lifters, long-distance runners, and comparatively sized untrained control subjects. Left ventricular mass determinations by the Penn convention reveal increased absolute left ventricular masses in long-distance runners and competitive weight lifters with respect to controls matched for age, body weight, and body surface area, and a significant correlation between ventricular mass and lean body mass. When normalized to lean body mass, the ventricular masses of distance runners are found to be significantly higher than those of the other groups, suggesting that dynamic training elevates left ventricular mass compared to static training and no training, while static training increases ventricular mass only to the extent that lean body mass is increased.

Longhurst, J. C.↗
NASA NTRS1990

The mass of the classical Cepheid SU Cygni

Velocities of the companion of the classical Cepheid SU Cyg have been measured from 13 IUE high-dispersion spectra. Because the companion is itself a member of a short-period binary system, its short-period orbit and the orbital amplitude of the center of mass in the long-period system are determined from a triple system solution. The velocity amplitude of the short-period center of mass in the long-period orbit is 32.2 + or - 1.6 km/s. This leads to a mass ratio between the mass of the Cepheid and the sum of the masses of the companions of 1.07 + or - 0.05. Combining this mass ratio with the mass function from the Cepheid orbit produces a purely dynamical lower limit to the mass of the Cepheid of 5.9 + or - 0.4 solar masses, in good agreement with the evolutionary mass for the Cepheid without mass loss or convective overshoot near the main sequence.

Evans, Nancy Remage↗
NASA NTRS1995

Radial velocities of very low mass stars and candidate brown dwarf members of the Hyades and Pleiades, 2

We have determined H alpha equivalent widths and radial velocities with 1 sigma accuracies of approximately 5 km/s for approximately 20 candidate very low mass members of the Pleiades cluster and for a few proposed very low mass members of the Hyades. Most of the Pleiades targets were selected from the recent Hambly, Hawkins, and Jameson proper motion survey, where they were identified as probable Pleiades brown dwarfs with an age spread from 3 to 70 Myr. Our spectroscopic data and a reinterpretation of the photometric data confirm that these objects are indeed likely Pleiades members; however, we believe that they more likely have masses slightly above the hydrogen burning mass limit and that there is no firm evidence for an age spread amongst these stars. All of the very low mass Pleiades and Hyades members show H alpha in emission. However, the ratio of H alpha flux to biometric flux in the Pleiades shows a maximum near M(sub Bol) approximately equal to 9.5 (M approximately equal to 0.3 solar mass) and a sharp decrease to lower masses. This break occurs at the approximate mass where low mass stars are expected to become fully convective, and it is tempting to assume that the decrease in H alpha flux is caused by some change in the behavior of stellar dynamos at this mass. We do not see a similar break in activity at this mass in the Hyades. We discuss possible evolutionary explanations for this difference in the H alpha activity between the two clusters.

Stauffer, John R.↗
NASA NTRS2009

Food Mass Reduction Trade Study

Future long duration manned space flights beyond low earth orbit will require the food system to remain safe, acceptable, and nutritious while efficiently balancing appropriate vehicle resources such as mass, volume, power, water, and crewtime. Often, this presents a challenge since maintaining the quality of the food system can result in a higher mass and volume. The Orion vehicle is significantly smaller than the Shuttle vehicle and the International Space Station and the mass and volume available for food is limited. Therefore, the food team has been challenged to reduce the mass of the packaged food from 1.82 kg per person per day to 1.14 kg per person per day. Past work has concentrated on how to reduce the mass of the packaging which contributes to about 15% of the total mass of the packaged food system. Designers have also focused on integrating and optimizing the Orion galley equipment as a system to reduce mass. To date, there has not been a significant effort to determine how to reduce the food itself. The objective of this project is to determine how the mass and volume of the packaged food can be reduced while maintaining caloric and hydration requirements. The following tasks are the key elements to this project: (1) Conduct further analysis of the ISS Standard Menu to determine moisture, protein, carbohydrate, and fat levels. (2) Conduct trade studies to determine how to bring the mass of the food system down. Trade studies may include removing the water of the total food system and/or increasing the fat content. (3) Determine the preferred method for delivery of the new food (e.g. bars, or beverages) and the degree of replacement. (4) Determine whether there are commercially available products that meet the requirements. By the end of this study, an estimate of the mass and volume savings will be provided to the Constellation Program. In addition, if new technologies need to be developed to achieve the mass savings, the technologies, timeline, and budget will be identified at the end of the project.

Perchonok, Michele H.↗