Voyager 2 Neptune targeting strategy

NASA Technical Reports Server (NTRS)

Potts, C. L.; Francis, K.; Matousek, S. E.; Cesarone, R. J.; Gray, D. L.

1989-01-01

The success of the Voyager 2 flybys of Neptune and Triton depends upon the ability to correct the spacecraft's trajectory. Accurate spacecraft delivery to the desired encounter conditions will promote the maximum science return. However, Neptune's great distance causes large a priori uncertainties in Neptune and Triton ephemerides and planetary system parameters. Consequently, the 'ideal' trajectory is unknown beforehand. The targeting challenge is to utilize the gradually improving knowledge as the spacecraft approaches Neptune to meet the science objectives, but with an overriding concern for spacecraft safety and a desire to limit propellant expenditure. A unique targeting strategy has been developed in response to this challenge. Through the use of a Monte Carlo simulation, candidate strategies are evaluated by the degree to which they meet these objectives and are compared against each other in determining the targeting strategy to be adopted.

Outgassing Data for Selecting Spacecraft Materials

NASA Technical Reports Server (NTRS)

Campbell, William A., Jr.; Scialdone, John J.

1993-01-01

This tenth compilation of outgassing data of materials intended for spacecraft use supersedes Reference Publication 1124, Revision 2, November 1990. The data were obtained at the Goddard Space Flight Center (GSFC), utilizing equipment developed at Stanford Research Institute (SRI) under contract to the Jet Propulsion Laboratory (JPL). SRI personnel developed an apparatus for determining the mass loss in vacuum and for collecting the outgassed products. The outgassing data have been presented in three different ways in order to facilitate material selection. In Section A, the materials are divided by category into the 18 probable uses, such as adhesives, greases, paints, potting compounds, and so forth. In Section B, all the materials contained in Section A are listed in alphabetical order by the manufacturer's identification. In Section C, the only materials listed are those having 'Total Mass Loss' (TML) and Collected Volatile Condensable Materials (CVCM) equal to or lower than a maximum 1.0 percent TML and a maximum 0.10 percent CVCM. These are grouped by use, as in Section A.

Pursuit/evasion in orbit

NASA Technical Reports Server (NTRS)

Kelley, H. J.; Cliff, E. M.; Lutze, F. H.

1981-01-01

Maneuvers available to a spacecraft having sufficient propellant to escape an antisatellite satellite (ASAT) attack are examined. The ASAT and the evading spacecraft are regarded as being in circular orbits, and equations of motion are developed for the ASAT to commence a two-impulse maneuver sequence. The ASAT employs thrust impulses which yield a minimum-time-to-rendezvous, considering available fuel. Optimal evasion is shown to involve only in-plane maneuvers, and begins as soon as the ASAT launch information is gathered and thrust activation can be initiated. A closest approach, along with a maximum evasion by the target spacecraft, is calculated to be 14,400 ft. Further research to account for ASATs in parking orbit and for generalization of a continuous control-modeled differential game is indicated.

Mesh Network Architecture for Enabling Inter-Spacecraft Communication

NASA Technical Reports Server (NTRS)

Becker, Christopher; Merrill, Garrick

2017-01-01

To enable communication between spacecraft operating in a formation or small constellation, a mesh network architecture was developed and tested using a time division multiple access (TDMA) communication scheme. The network is designed to allow for the exchange of telemetry and other data between spacecraft to enable collaboration between small spacecraft. The system uses a peer-to-peer topology with no central router, so that it does not have a single point of failure. The mesh network is dynamically configurable to allow for addition and subtraction of new spacecraft into the communication network. Flight testing was performed using an unmanned aerial system (UAS) formation acting as a spacecraft analogue and providing a stressing environment to prove mesh network performance. The mesh network was primarily devised to provide low latency, high frequency communication but is flexible and can also be configured to provide higher bandwidth for applications desiring high data throughput. The network includes a relay functionality that extends the maximum range between spacecraft in the network by relaying data from node to node. The mesh network control is implemented completely in software making it hardware agnostic, thereby allowing it to function with a wide variety of existing radios and computing platforms..

Electric propulsion options for 10 kW class earth space missions

NASA Technical Reports Server (NTRS)

Patterson, M. J.; Curran, Francis M.

1989-01-01

Five and 10 kW ion and arcjet propulsion system options for a near-term space demonstration experiment have been evaluated. Analyses were conducted to determine first-order propulsion system performance and system component mass estimates. Overall mission performance of the electric propulsion systems was quantified in terms of the maximum thrusting time, total impulse, and velocity increment capability available when integrated onto a generic spacecraft under fixed mission model assumptions. Maximum available thrusting times for the ion-propelled spacecraft options, launched on a DELTA II 6920 vehicle, range from approximately 8,600 hours for a 4-engine 10 kW system to more than 29,600 hours for a single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 1.2x10(7) to 2.1x10(7) N-s, and 3550 to 6200 m/s, respectively. Maximum available thrusting times for the arcjet propelled spacecraft launched on the DELTA II 6920 vehicle range from approximately 528 hours for the 6-engine 10 kW hydrazine system to 2328 hours for the single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 2.2x10(6) to 3.6x10(6) N-s, and approximately 662 to 1072 m/s, respectively.

Methods for utilizing maximum power from a solar array

NASA Technical Reports Server (NTRS)

Decker, D. K.

1972-01-01

A preliminary study of maximum power utilization methods was performed for an outer planet spacecraft using an ion thruster propulsion system and a solar array as the primary energy source. The problems which arise from operating the array at or near the maximum power point of its 1-V characteristic are discussed. Two closed loop system configurations which use extremum regulators to track the array's maximum power point are presented. Three open loop systems are presented that either: (1) measure the maximum power of each array section and compute the total array power, (2) utilize a reference array to predict the characteristics of the solar array, or (3) utilize impedance measurements to predict the maximum power utilization. The advantages and disadvantages of each system are discussed and recommendations for further development are made.

Deep space telecommunications and the solar cycle: A reappraisal

NASA Technical Reports Server (NTRS)

Berman, A. L.

1978-01-01

Observations of density enhancement in the near corona at solar cycle (sunspot) maximum have rather uncritically been interpreted to apply equally well to the extended corona, thus generating concern about the quality of outer planet navigational data at solar cycle maximum. Spacecraft have been deployed almost continuously during the recently completed solar cycle 20, providing two powerful new coronal investigatory data sources: (1) in-situ spacecraft plasma measurements at approximately 1 AU, and (2) plasma effects on monochromatic spacecraft signals at all signal closest approach points. A comprehensive review of these (solar cycle 20) data lead to the somewhat surprising conclusions that for the region of interest of navigational data, the highest levels of charged particle corruption of navigational data can be expected to occur at solar cycle minimum, rather than solar cycle maximum, as previously believed.

Model for Cumulative Solar Heavy Ion Energy and LET Spectra

NASA Technical Reports Server (NTRS)

Xapsos, Mike; Barth, Janet; Stauffer, Craig; Jordan, Tom; Mewaldt, Richard

2007-01-01

A probabilistic model of cumulative solar heavy ion energy and lineary energy transfer (LET) spectra is developed for spacecraft design applications. Spectra are given as a function of confidence level, mission time period during solar maximum and shielding thickness. It is shown that long-term solar heavy ion fluxes exceed galactic cosmic ray fluxes during solar maximum for shielding levels of interest. Cumulative solar heavy ion fluences should therefore be accounted for in single event effects rate calculations and in the planning of space missions.

A Power Conditioning Stage Based on Analog-Circuit MPPT Control and a Superbuck Converter for Thermoelectric Generators in Spacecraft Power Systems

NASA Astrophysics Data System (ADS)

Sun, Kai; Wu, Hongfei; Cai, Yan; Xing, Yan

2014-06-01

A thermoelectric generator (TEG) is a very important kind of power supply for spacecraft, especially for deep-space missions, due to its long lifetime and high reliability. To develop a practical TEG power supply for spacecraft, a power conditioning stage is indispensable, being employed to convert the varying output voltage of the TEG modules to a definite voltage for feeding batteries or loads. To enhance the system reliability, a power conditioning stage based on analog-circuit maximum-power-point tracking (MPPT) control and a superbuck converter is proposed in this paper. The input of this power conditioning stage is connected to the output of the TEG modules, and the output of this stage is connected to the battery and loads. The superbuck converter is employed as the main circuit, featuring low input current ripples and high conversion efficiency. Since for spacecraft power systems reliable operation is the key target for control circuits, a reset-set flip-flop-based analog circuit is used as the basic control circuit to implement MPPT, being much simpler than digital control circuits and offering higher reliability. Experiments have verified the feasibility and effectiveness of the proposed power conditioning stage. The results show the advantages of the proposed stage, such as maximum utilization of TEG power, small input ripples, and good stability.

COBE ground segment gyro calibration

NASA Technical Reports Server (NTRS)

Freedman, I.; Kumar, V. K.; Rae, A.; Venkataraman, R.; Patt, F. S.; Wright, E. L.

1991-01-01

Discussed here is the calibration of the scale factors and rate biases for the Cosmic Background Explorer (COBE) spacecraft gyroscopes, with the emphasis on the adaptation for COBE of an algorithm previously developed for the Solar Maximum Mission. Detailed choice of parameters, convergence, verification, and use of the algorithm in an environment where the reference attitudes are determined form the Sun, Earth, and star observations (via the Diffuse Infrared Background Experiment (DIRBE) are considered. Results of some recent experiments are given. These include tests where the gyro rate data are corrected for the effect of the gyro baseplate temperature on the spacecraft electronics.

Electric Propulsion Options for 10 kW Class Earth-Space Missions

NASA Technical Reports Server (NTRS)

Patterson, M. J.; Curran, Francis M.

1989-01-01

Five and 10 kW ion and arcjet propulsion system options for a near-term space demonstration experiment were evaluated. Analyses were conducted to determine first-order propulsion system performance and system component mass estimates. Overall mission performance of the electric propulsion systems was quantified in terms of the maximum thrusting time, total impulse, and velocity increment capability available when integrated onto a generic spacecraft under fixed mission model assumptions. Maximum available thrusting times for the ion-propelled spacecraft options, launched on a DELTA 2 6920 vehicle, range from approximately 8,600 hours for a 4-engine 10 kW system to more than 29,600 hours for a single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 1.2x10 (exp 7) to 2.1x10 (exp 7) N-s, and 3550 to 6200 m/s, respectively. Maximum available thrusting times for the arcjet propelled spacecraft launched on the DELTA 2 6920 vehicle range from approximately 528 hours for the 6-engine 10 kW hydrazine system to 2328 hours for the single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 2.2x10 (exp 6) to 3.6x10 (exp 6) N-s, and approximately 662 to 1072 m/s, respectively.

Performance of Nickel-Cadmium Batteries on the GOES I-K Series of Weather Satellites

NASA Technical Reports Server (NTRS)

Singhal, Sat P.; Rao, Gopalakrishna M.; Alsbach, Walter G.

1997-01-01

The US National Oceanic and Atmospheric Administration (NOAA) operates the Geostationary Operational Environmental Satellite (GOES) spacecraft (among others) to support weather forecasting, severe storm tracking, and meteorological research by the National Weather Service (NWS). The latest in the GOES series consists of 5 spacecraft (originally named GOES I-M), three of which are in orbit and two more in development. Each of five spacecraft carry two Nickel-Cadmium batteries, with batteries designed and manufactured by Space Systems Loral (SS/L) and cells manufactured by Gates Aerospace Batteries (sold to SAFT in 1993). The battery, which consists of 28 cells with a 12 Ah capacity, provides the spacecraft power needs during the ascent phase and during the semi-annual eclipse seasons lasting for approximately 45 days each. The maximum duration eclipses are 72 minutes long which result in a 60 percent depth of discharge (DOD) of the batteries. This paper provides a description of the batteries, reconditioning setup, DOD profile during a typical eclipse season, and flight performance from the 3 launched spacecraft (now GOES 8, 9, and 10) in orbit.

Performance of Nickel-Cadmium Batteries on the GOES 1-K Series of Weather Satellites

NASA Technical Reports Server (NTRS)

Singhal, Sat P.; Alsbach, Walter G.; Rao, Gopalakrishna M.

1998-01-01

The US National Oceanic and Atmospheric Administration (NOAA) operates the Geostationary Operational Environmental Satellite (GOES) spacecraft (among others) to support weather forecasting, severe storm tracking, and meteorological research by the National Weather Service (NWS). The latest in the GOES series consists of five spacecraft (originally named GOES 1-M), three of which are in orbit and and two more in development. Each of the five spacecraft carries two Nickel-Cadmium battery, with batteries designed by Space Systems Loral (SS/L) and cells manufactured by Gates Aerospace Batteries (sold to SAFT in 1993). The battery, which consists of 28 cells with a 12 Ah capacity, provides the spacecraft power needs during the ascent phase and during the semi-annual eclipse seasons lasting for approximately 45 days each. The maximum duration eclipses are 72 minutes long which result in a 60 percent depth of discharge (DOD) of the batteries. This paper provides a description of the batteries, reconditioning setup, DOD profile during a typical eclipse season, and flight performance from the three launched spacecraft (now GOES 8, 9, and 10) in orbit.

Lifetime predictions for the Solar Maximum Mission (SMM) and San Marco spacecraft

NASA Technical Reports Server (NTRS)

Smith, E. A.; Ward, D. T.; Schmitt, M. W.; Phenneger, M. C.; Vaughn, F. J.; Lupisella, M. L.

1989-01-01

Lifetime prediction techniques developed by the Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD) are described. These techniques were developed to predict the Solar Maximum Mission (SMM) spacecraft orbit, which is decaying due to atmospheric drag, with reentry predicted to occur before the end of 1989. Lifetime predictions were also performed for the Long Duration Exposure Facility (LDEF), which was deployed on the 1984 SMM repair mission and is scheduled for retrieval on another Space Transportation System (STS) mission later this year. Concepts used in the lifetime predictions were tested on the San Marco spacecraft, which reentered the Earth's atmosphere on December 6, 1988. Ephemerides predicting the orbit evolution of the San Marco spacecraft until reentry were generated over the final 90 days of the mission when the altitude was less than 380 kilometers. The errors in the predicted ephemerides are due to errors in the prediction of atmospheric density variations over the lifetime of the satellite. To model the time dependence of the atmospheric densities, predictions of the solar flux at the 10.7-centimeter wavelength were used in conjunction with Harris-Priester (HP) atmospheric density tables. Orbital state vectors, together with the spacecraft mass and area, are used as input to the Goddard Trajectory Determination System (GTDS). Propagations proceed in monthly segments, with the nominal atmospheric drag model scaled for each month according to the predicted monthly average value of F10.7. Calibration propagations are performed over a period of known orbital decay to obtain the effective ballistic coefficient. Progagations using plus or minus 2 sigma solar flux predictions are also generated to estimate the despersion in expected reentry dates. Definitive orbits are compared with these predictions as time expases. As updated vectors are received, these are also propagated to reentryto continually update the lifetime predictions.

Special features of the CLUSTER antenna and radial booms design, development and verification

NASA Technical Reports Server (NTRS)

Gianfiglio, G.; Yorck, M.; Luhmann, H. J.

1995-01-01

CLUSTER is a scientific space mission to in-situ investigate the Earth's plasma environment by means of four identical spin-stabilized spacecraft. Each spacecraft is provided with a set of four rigid booms: two Antenna Booms and two Radial Booms. This paper presents a summary of the boom development and verification phases addressing the key aspects of the Radial Boom design. In particular, it concentrates on the difficulties encountered in fulfilling simultaneously the requirements of minimum torque ratio and maximum allowed shock loads at boom latching for this two degree of freedom boom. The paper also provides an overview of the analysis campaign and testing program performed to achieve sufficient confidence in the boom performance and operation.

Motion-Based System Identification and Fault Detection and Isolation Technologies for Thruster Controlled Spacecraft

NASA Technical Reports Server (NTRS)

Wilson, Edward; Sutter, David W.; Berkovitz, Dustin; Betts, Bradley J.; Kong, Edmund; delMundo, Rommel; Lages, Christopher R.; Mah, Robert W.; Papasin, Richard

2003-01-01

By analyzing the motions of a thruster-controlled spacecraft, it is possible to provide on-line (1) thruster fault detection and isolation (FDI), and (2) vehicle mass- and thruster-property identification (ID). Technologies developed recently at NASA Ames have significantly improved the speed and accuracy of these ID and FDI capabilities, making them feasible for application to a broad class of spacecraft. Since these technologies use existing sensors, the improved system robustness and performance that comes with the thruster fault tolerance and system ID can be achieved through a software-only implementation. This contrasts with the added cost, mass, and hardware complexity commonly required by FDI. Originally developed in partnership with NASA - Johnson Space Center to provide thruster FDI capability for the X-38 during re-entry, these technologies are most recently being applied to the MIT SPHERES experimental spacecraft to fly on the International Space Station in 2004. The model-based FDI uses a maximum-likelihood calculation at its core, while the ID is based upon recursive least squares estimation. Flight test results from the SPHERES implementation, as flown aboard the NASA KC-1 35A 0-g simulator aircraft in November 2003 are presented.

Spacecraft attitude determination using the earth's magnetic field

NASA Technical Reports Server (NTRS)

Simpson, David G.

1989-01-01

A method is presented by which the attitude of a low-Earth orbiting spacecraft may be determined using a vector magnetometer, a digital Sun sensor, and a mathematical model of the Earth's magnetic field. The method is currently being implemented for the Solar Maximum Mission spacecraft (as a backup for the failing star trackers) as a way to determine roll gyro drift.

Ascent performance feasibility for next-generation spacecraft

NASA Astrophysics Data System (ADS)

Mancuso, Salvatore Massimo

This thesis deals with the optimization of the ascent trajectories for single-stage suborbital (SSSO), single-stage-to-orbit (SSTO), and two-stage-to-orbit (TSTO) rocket-powered spacecraft. The maximum payload weight problem has been solved using the sequential gradient-restoration algorithm. For the TSTO case, some modifications to the original version of the algorithm have been necessary in order to deal with discontinuities due to staging and the fact that the functional being minimized depends on interface conditions. The optimization problem is studied for different values of the initial thrust-to-weight ratio in the range 1.3 to 1.6, engine specific impulse in the range 400 to 500 sec, and spacecraft structural factor in the range 0.08 to 0.12. For the TSTO configuration, two subproblems are studied: uniform structural factor between stages and nonuniform structural factor between stages. Due to the regular behavior of the results obtained, engineering approximations have been developed which connect the maximum payload weight to the engine specific impulse and spacecraft structural factor; in turn, this leads to useful design considerations. Also, performance sensitivity to the scale of the aerodynamic drag is studied, and it is shown that its effect on payload weight is relatively small, even for drag changes approaching ± 50%. The main conclusions are that: the design of a SSSO configuration appears to be feasible; the design of a SSTO configuration might be comfortably feasible, marginally feasible, or unfeasible, depending on the parameter values assumed; the design of a TSTO configuration is not only feasible, but its payload appears to be considerably larger than that of a SSTO configuration. Improvements in engine specific impulse and spacecraft structural factor are desirable and crucial for SSTO feasibility; indeed, it appears that aerodynamic improvements do not yield significant improvements in payload weight.

Modeling of the Terminal Velocities of the Dust Ejected Material by the Impact

NASA Astrophysics Data System (ADS)

Rengel, M.; Küppers, M.; Keller, H. U.; Gutiérrez, P.

We compute the distribution of velocities of the particles ejected by the impact of the projectile released from NASA Deep Impact spacecraft on the nucleus of comet 9P/Tempel 1 on the successive 20 h following the collision. This is performed by the development and use of an ill-conditioned inverse problem approach, whose main ingredients are a set of observations taken by the Narrow Angle Camera (NAC) of OSIRIS onboard the Rosetta spacecraft, and a set of simple models of the expansion of the dust ejecta plume for different velocities. Terminal velocities are derived using a maximum likelihood estimator.

Proceedings of the Spacecraft Charging Technology Conference Held in Monterey, California on 31 October - 3 November 1989. Volume 2

DTIC Science & Technology

1989-11-01

STATEMENT 12b. DISTRIBUTION CODE Approved for public release; distribution unlimited 13. ABSTRACT (Maximum 200 words) The Spacecraft Charging... Distribution I D Availability Codes ’ Avail and/orDist Special VIvt PREFACE The Spacecraft Charging Technology conference was held at the Naval... distribution , the spacecraft will charge negatively during this time according to dV/dt = 47ta 2 Jth ev/° / C whose solution is V/0= - ln(l + t/t) "t = C 0

Spacecraft Charging and the Microwave Anisotropy Probe Spacecraft

NASA Technical Reports Server (NTRS)

Timothy, VanSant J.; Neergaard, Linda F.

1998-01-01

The Microwave Anisotropy Probe (MAP), a MIDEX mission built in partnership between Princeton University and the NASA Goddard Space Flight Center (GSFC), will study the cosmic microwave background. It will be inserted into a highly elliptical earth orbit for several weeks and then use a lunar gravity assist to orbit around the second Lagrangian point (L2), 1.5 million kilometers, anti-sunward from the earth. The charging environment for the phasing loops and at L2 was evaluated. There is a limited set of data for L2; the GEOTAIL spacecraft measured relatively low spacecraft potentials (approx. 50 V maximum) near L2. The main area of concern for charging on the MAP spacecraft is the well-established threat posed by the "geosynchronous region" between 6-10 Re. The launch in the autumn of 2000 will coincide with the falling of the solar maximum, a period when the likelihood of a substorm is higher than usual. The likelihood of a substorm at that time has been roughly estimated to be on the order of 20% for a typical MAP mission profile. Because of the possibility of spacecraft charging, a requirement for conductive spacecraft surfaces was established early in the program. Subsequent NASCAP/GEO analyses for the MAP spacecraft demonstrated that a significant portion of the sunlit surface (solar cell cover glass and sunshade) could have nonconductive surfaces without significantly raising differential charging. The need for conductive materials on surfaces continually in eclipse has also been reinforced by NASCAP analyses.

Microprocessor control of photovoltaic systems

NASA Technical Reports Server (NTRS)

Millner, A. R.; Kaufman, D. L.

1984-01-01

The present low power CMOS microprocessor controller for photovoltaic power systems possesses three programs, which are respectively intended for (1) conventional battery-charging systems with state-of-charge estimation and sequential shedding of subarrays and loads, (2) maximum power-controlled battery-charging systems, and (3) variable speed dc motor drives. Attention is presently given to the development of this terrestrial equipment for spacecraft use.

Computer-Aided Modeling and Analysis of Power Processing Systems (CAMAPPS), phase 1

NASA Technical Reports Server (NTRS)

Kim, S.; Lee, J.; Cho, B. H.; Lee, F. C.

1986-01-01

The large-signal behaviors of a regulator depend largely on the type of power circuit topology and control. Thus, for maximum flexibility, it is best to develop models for each functional block a independent modules. A regulator can then be configured by collecting appropriate pre-defined modules for each functional block. In order to complete the component model generation for a comprehensive spacecraft power system, the following modules were developed: solar array switching unit and control; shunt regulators; and battery discharger. The capability of each module is demonstrated using a simplified Direct Energy Transfer (DET) system. Large-signal behaviors of solar array power systems were analyzed. Stability of the solar array system operating points with a nonlinear load is analyzed. The state-plane analysis illustrates trajectories of the system operating point under various conditions. Stability and transient responses of the system operating near the solar array's maximum power point are also analyzed. The solar array system mode of operation is described using the DET spacecraft power system. The DET system is simulated for various operating conditions. Transfer of the software program CAMAPPS (Computer Aided Modeling and Analysis of Power Processing Systems) to NASA/GSFC (Goddard Space Flight Center) was accomplished.

Method and apparatus for calibrating the ionosphere and application to surveillance of geophysical events

NASA Technical Reports Server (NTRS)

Macdoran, P. F. (Inventor)

1984-01-01

The columnar electron content of the ionosphere between a spacecraft and a receiver is measured in realtime by cross correlating two coherently modulated signals transmitted at different frequencies (L1,L2) from the spacecraft to the receiver using a cross correlator. The time difference of arrival of the modulated signals is proportional to electron content of the ionosphere. A variable delay is adjusted relative to a fixed delay in the respective channels (L1,L2) to produce a maximum at the cross correlator output. The difference in delay required to produce this maximum is a measure of the columnar electron content of the ionosphere. A plurality of monitoring stations and spacecraft (Global Positioning System satellites) are employed to locate any terrestrial event that produces an ionospheric disturbance.

In-flight calibration and performance evaluation of the fixed head star trackers for the solar maximum mission

NASA Technical Reports Server (NTRS)

Thompson, R. H.; Gambardella, P. J.

1980-01-01

The Solar Maximum Mission (SMM) spacecraft provides an excellent opportunity for evaluating attitude determination accuracies achievable with tracking instruments such as fixed head star trackers (FHSTs). As a part of its payload, SMM carries a highly accurate fine pointing Sun sensor (FPSS). The EPSS provides an independent check of the pitch and yaw parameters computed from observations of stars in the FHST field of view. A method to determine the alignment of the FHSTs relative to the FPSS using spacecraft data is applied. Two methods that were used to determine distortions in the 8 degree by 8 degree field of view of the FHSTs using spacecraft data are also presented. The attitude determination accuracy performance of the in flight calibrated FHSTs is evaluated.

Temperature dependence of attitude sensor coalignments on the Solar Maximum Mission (SMM)

NASA Technical Reports Server (NTRS)

Pitone, D. S.; Eudell, A. H.; Patt, F. S.

1990-01-01

The temperature correlation of the relative coalignment between the fine-pointing sun sensor and fixed-head star trackers measured on the Solar Maximum Mission (SMM) is analyzed. An overview of the SMM, including mission history and configuration, is given. Possible causes of the misalignment variation are discussed, with focus placed on spacecraft bending due to solar-radiation pressure, electronic or mechanical changes in the sensors, uncertainty in the attitude solutions, and mounting-plate expansion and contraction due to thermal effects. Yaw misalignment variation from the temperature profile is assessed, and suggestions for spacecraft operations are presented, involving methods to incorporate flight measurements of the temperature-versus-alignment function and its variance in operational procedures and the spacecraft structure temperatures in the attitude telemetry record.

Nano-Satellite Avionics

NASA Technical Reports Server (NTRS)

Culver, Harry

1999-01-01

Abstract NASA's Goddard Space Flight Center (GSFC) is currently developing a new class of satellites called the nano-satellite (nano-sat). A major objective of this development effort is to provide the technology required to enable a constellation of tens to hundreds of nano-satellites to make both remote and in-situ measurements from space. The Nano-sat will be a spacecraft weighing a maximum of 10 kg, including the propellant mass, and producing at least 5 Watts of power to operate the spacecraft. The electronics are required to survive a total radiation dose rate of 100 krads for a mission lifetime of two years. There are many unique challenges that must be met in order to develop the avionics for such a spacecraft. The first challenge is to develop an architecture that will operate on the allotted 5 Watts and meet the diverging requirements of multiple missions. This architecture will need to incorporate a multitude of new advanced microelectronic technologies. The microelectronics developed must be a modular and scalable packaging of technology to solve the problem of developing a solution to both reduce cost and meet the requirements of various missions. This development will utilize the most cost effective approach, whether infusing commercially driven semiconductor devices into spacecraft applications or partnering with industry to design and develop low cost, low power, low mass, and high capacity data processing devices. This paper will discuss the nano-sat architecture and the major technologies that will be developed. The major technologies that will be covered include: (1) Light weight Low Power Electronics Packaging, (2) Radiation Hard/Tolerant, Low Power Processing Platforms, (3) High capacity Low Power Memory Systems (4) Radiation Hard reconfiguragble field programmable gate array (rFPGA)

Pulsed Thrust Method for Hover Formation Flying

NASA Technical Reports Server (NTRS)

Hope, Alan; Trask, Aaron

2003-01-01

A non-continuous thrust method for hover type formation flying has been developed. This method differs from a true hover which requires constant range and bearing from a reference vehicle. The new method uses a pulsed loop, or pogo, maneuver sequence that keeps the follower spacecraft within a defined box in a near hover situation. Equations are developed for the hover maintenance maneuvers. The constraints on the hover location, pulse interval, and maximum/minimum ranges are discussed.

30-kW SEP Spacecraft as Secondary Payloads for Low-Cost Deep Space Science Missions

NASA Technical Reports Server (NTRS)

Brophy, John R.; Larson, Tim

2013-01-01

The Solar Array System contracts awarded by NASA's Space Technology Mission Directorate are developing solar arrays in the 30 kW to 50 kW power range (beginning of life at 1 AU) that have significantly higher specific powers (W/kg) and much smaller stowed volumes than conventional rigid-panel arrays. The successful development of these solar array technologies has the potential to enable new types of solar electric propulsion (SEP) vehicles and missions. This paper describes a 30-kW electric propulsion vehicle built into an EELV Secondary Payload Adapter (ESPA) ring. The system uses an ESPA ring as the primary structure and packages two 15-kW Megaflex solar array wings, two 14-kW Hall thrusters, a hydrazine Reaction Control Subsystem (RCS), 220 kg of xenon, 26 kg of hydrazine, and an avionics module that contains all of the rest of the spacecraft bus functions and the instrument suite. Direct-drive is used to maximize the propulsion subsystem efficiency and minimize the resulting waste heat and required radiator area. This is critical for packaging a high-power spacecraft into a very small volume. The fully-margined system dry mass would be approximately 1120 kg. This is not a small dry mass for a Discovery-class spacecraft, for example, the Dawn spacecraft dry mass was only about 750 kg. But the Dawn electric propulsion subsystem could process a maximum input power of 2.5 kW, and this spacecraft would process 28 kW, an increase of more than a factor of ten. With direct-drive the specific impulse would be limited to about 2,000 s assuming a nominal solar array output voltage of 300 V. The resulting spacecraft would have a beginning of life acceleration that is more than an order of magnitude greater than the Dawn spacecraft. Since the spacecraft would be built into an ESPA ring it could be launched as a secondary payload to a geosynchronous transfer orbit significantly reducing the launch costs for a planetary spacecraft. The SEP system would perform the escape from Earth and then the heliocentric transfer to the science target.

Extreme Spacecraft Charging in Polar Low Earth Orbit

NASA Technical Reports Server (NTRS)

Colson, Andrew D.; Minow, Joseph I.; NeergaardParker, Linda

2012-01-01

Spacecraft in low altitude, high inclination (including sun-synchronous) orbits are widely used for remote sensing of the Earth's land surface and oceans, monitoring weather and climate, communications, scientific studies of the upper atmosphere and ionosphere, and a variety of other scientific, commercial, and military applications. These systems episodically charge to frame potentials in the kilovolt range when exposed to space weather environments characterized by a high flux of energetic (10 s kilovolt) electrons in regions of low background plasma density which is similar in some ways to the space weather conditions in geostationary orbit responsible for spacecraft charging to kilovolt levels. We first review the physics of space environment interactions with spacecraft materials that control auroral charging rates and the anticipated maximum potentials that should be observed on spacecraft surfaces during disturbed space weather conditions. We then describe how the theoretical values compare to the observational history of extreme charging in auroral environments. Finally, a set of extreme DMSP charging events are described varying in maximum negative frame potential from 0.6 kV to 2 kV, focusing on the characteristics of the charging events that are of importance both to the space system designer and to spacecraft operators. The goal of the presentation is to bridge the gap between scientific studies of auroral charging and the need for engineering teams to understand how space weather impacts both spacecraft design and operations for vehicles on orbital trajectories that traverse auroral charging environments.

Extreme Spacecraft Charging in Polar Low Earth Orbit

NASA Technical Reports Server (NTRS)

Colson, Andrew D.; Minow, Joseph I.; Parker, L. Neergaard

2012-01-01

Spacecraft in low altitude, high inclination (including sun -synchronous) orbits are widely used for remote sensing of the Earth fs land surface and oceans, monitoring weather and climate, communications, scientific studies of the upper atmosphere and ionosphere, and a variety of other scientific, commercial, and military applications. These systems episodically charge to frame potentials in the kilovolt range when exposed to space weather environments characterized by a high flux of energetic (approx.10 fs kilovolt) electrons in regions of low background plasma density. Auroral charging conditions are similar in some ways to the space weather conditions in geostationary orbit responsible for spacecraft charging to kilovolt levels. We first review the physics of space environment interactions with spacecraft materials that control auroral charging rates and the anticipated maximum potentials that should be observed on spacecraft surfaces during disturbed space weather conditions. We then describe how the theoretical values compare to the observational history of extreme charging in auroral environments. Finally, a set of extreme DMSP charging events are described varying in maximum negative frame potential from approx.0.6 kV to approx.2 kV, focusing on the characteristics of the charging events that are of importance both to the space system designer and to spacecraft operators. The goal of the presentation is to bridge the gap between scientific studies of auroral charging and the need for engineering teams to understand how space weather impacts both spacecraft design and operations for vehicles on orbital trajectories that traverse auroral charging environments.

Test of Shi et al. Method to Infer the Magnetic Reconnection Geometry from Spacecraft Data: MHD Simulation with Guide Field and Antiparallel Kinetic Simulation

NASA Technical Reports Server (NTRS)

Denton, R.; Sonnerup, B. U. O.; Swisdak, M.; Birn, J.; Drake, J. F.; Heese, M.

2012-01-01

When analyzing data from an array of spacecraft (such as Cluster or MMS) crossing a site of magnetic reconnection, it is desirable to be able to accurately determine the orientation of the reconnection site. If the reconnection is quasi-two dimensional, there are three key directions, the direction of maximum inhomogeneity (the direction across the reconnection site), the direction of the reconnecting component of the magnetic field, and the direction of rough invariance (the "out of plane" direction). Using simulated spacecraft observations of magnetic reconnection in the geomagnetic tail, we extend our previous tests of the direction-finding method developed by Shi et al. (2005) and the method to determine the structure velocity relative to the spacecraft Vstr. These methods require data from four proximate spacecraft. We add artificial noise and calibration errors to the simulation fields, and then use the perturbed gradient of the magnetic field B and perturbed time derivative dB/dt, as described by Denton et al. (2010). Three new simulations are examined: a weakly three-dimensional, i.e., quasi-two-dimensional, MHD simulation without a guide field, a quasi-two-dimensional MHD simulation with a guide field, and a two-dimensional full dynamics kinetic simulation with inherent noise so that the apparent minimum gradient was not exactly zero, even without added artificial errors. We also examined variations of the spacecraft trajectory for the kinetic simulation. The accuracy of the directions found varied depending on the simulation and spacecraft trajectory, but all the directions could be found within about 10 for all cases. Various aspects of the method were examined, including how to choose averaging intervals and the best intervals for determining the directions and velocity. For the kinetic simulation, we also investigated in detail how the errors in the inferred gradient directions from the unmodified Shi et al. method (using the unperturbed gradient) depended on the amplitude of the calibration errors. For an accuracy of 3 for the maximum gradient direction, the calibration errors could be as large as 3% of reconnection magnetic field, while for the same accuracy for the minimum gradient direction, the calibration errors could only be as large as 0.03% of the reconnection magnetic field. These results suggest that the maximum gradient direction can normally be determined by the unmodified Shi et al. method, while the modified method or some other method must be used to accurately determine the minimum gradient direction. The structure velocity was found with magnitude accurate to 2% and direction accurate to within 5%.

Scientific Analysis of Data for the ISTP/SOLARMAX Programs

NASA Technical Reports Server (NTRS)

Lazarus, Alan J.

2001-01-01

This Grant supplemented our work on data analysis from the Wind spacecraft which was one of the ISTRIA fleet of spacecraft. It was targeted at observations related to the time of solar maximum in 2000. The work we proposed to do under this grant included comparison of solar wind parameters obtained from different spacecraft in order to establish correlation lengths appropriate to the solar wind and also to compare parameters to explore solar cycle effects.

Toward a Dynamically Reconfigurable Computing and Communication System for Small Spacecraft

NASA Technical Reports Server (NTRS)

Kifle, Muli; Andro, Monty; Tran, Quang K.; Fujikawa, Gene; Chu, Pong P.

2003-01-01

Future science missions will require the use of multiple spacecraft with multiple sensor nodes autonomously responding and adapting to a dynamically changing space environment. The acquisition of random scientific events will require rapidly changing network topologies, distributed processing power, and a dynamic resource management strategy. Optimum utilization and configuration of spacecraft communications and navigation resources will be critical in meeting the demand of these stringent mission requirements. There are two important trends to follow with respect to NASA's (National Aeronautics and Space Administration) future scientific missions: the use of multiple satellite systems and the development of an integrated space communications network. Reconfigurable computing and communication systems may enable versatile adaptation of a spacecraft system's resources by dynamic allocation of the processor hardware to perform new operations or to maintain functionality due to malfunctions or hardware faults. Advancements in FPGA (Field Programmable Gate Array) technology make it possible to incorporate major communication and network functionalities in FPGA chips and provide the basis for a dynamically reconfigurable communication system. Advantages of higher computation speeds and accuracy are envisioned with tremendous hardware flexibility to ensure maximum survivability of future science mission spacecraft. This paper discusses the requirements, enabling technologies, and challenges associated with dynamically reconfigurable space communications systems.

Process and methodology of developing Cassini G and C Telemetry Dictionary

NASA Technical Reports Server (NTRS)

Kan, Edwin P.

1994-01-01

While the Cassini spacecraft telemetry design had taken on the new approach of 'packetized telemetry', the AACS (Attitude and Articulation Subsystem) had further extended into the design of 'mini-packets' in its telemetry system. Such telemetry packet and mini-packet design produced the AACS Telemetry Dictionary; iterations of the latter in turn provided changes to the former. The ultimate goals were to achieve maximum telemetry packing density, optimize the 'freshness' of more time-critical data, and to effect flexibility, i.e., multiple AACS data collection schemes, without needing to change the overall spacecraft telemetry mode. This paper describes such a systematic process and methodology, evidenced by various design products related to, or as part of, the AACS Telemetry Dictionary.

Design consideration for a nuclear electric propulsion system

NASA Technical Reports Server (NTRS)

Phillips, W. M.; Pawlik, E. V.

1978-01-01

A study is currently underway to design a nuclear electric propulsion vehicle capable of performing detailed exploration of the outer-planets. Primary emphasis is on the power subsystem. Secondary emphasis includes integration into a spacecraft, and integration with the thrust subsystem and science package or payload. The results of several design iterations indicate an all-heat-pipe system offers greater reliability, elimination of many technology development areas and a specific weight of under 20 kg/kWe at the 400 kWe power level. The system is compatible with a single Shuttle launch and provides greater safety than could be obtained with designs using pumped liquid metal cooling. Two configurations, one with the reactor and power conversion forward on the spacecraft with the ion engines aft and the other with reactor, power conversion and ion engines aft were selected as dual baseline designs based on minimum weight, minimum required technology development and maximum growth potential and flexibility.

Construction of optimum controls and trajectories of motion of the center of masses of a spacecraft equipped with the solar sail and low-thrust engine, using quaternions and Kustaanheimo-Stiefel variables

NASA Astrophysics Data System (ADS)

Sapunkov, Ya. G.; Chelnokov, Yu. N.

2014-11-01

The problem of optimum rendezvous of a controllable spacecraft (SC) with an uncontrollable spacecraft, moving over a Keplerian elliptic orbit in the gravitational field of the Sun, is considered. Control of the SC is performed using a solar sail and low-thrust engine. For solving the problem, the regular quaternion equations of the two-body problem with the Kustaanheimo-Stiefel variables and the Pontryagin maximum principle are used. The combined integral quality functional, which characterizes energy consumption for controllable SC transition from an initial to final state and the time spent for this transition, is used as a minimized functional. The differential boundary-value optimization problems are formulated, and their first integrals are found. Examples of numerical solution of problems are presented. The paper develops the application [1-6] of quaternion regular equations with the Kustaanheimo-Stiefel variables in the space flight mechanics.

High-latitude spacecraft charging in low-Earth polar orbit

NASA Astrophysics Data System (ADS)

Frooninckx, Thomas B.

Spacecraft charging within the upper ionosphere is commonly thought to be insignificant and thus has received little attention. Recent experimental evidence has shown that electric potential differences as severe as 680 volts can develop between Defense Meteorological Satellite Program (DMSP) polar-orbiting (840 kilometers) spacecraft and their high-latitude environment. To explore space vehicle charging in this region more fully, an analysis was performed using DMSP F6, F7, F8, and F9 satellite precipitating particle and ambient plasma measurements taken during the winters of 1986-87 (solar minimum) and 1989-90 (solar maximum). An extreme solar cycle dependence was discovered as charging occurred more frequently and with greater severity during the period of solar minimum. One hundred seventy charging events ranging from -46 to 1,430 volts were identified, and satellite measurements and Time Dependent Ionospheric Model (TDIM) output were used to characterize the environments which generated and inhibited these potentials. All current sources were considered to determine the cause of the solar cycle dependence.

Apollo experience report: Manned thermal-vacuum testing of spacecraft

NASA Technical Reports Server (NTRS)

Mclane, J. C., Jr.

1974-01-01

Manned thermal-vacuum tests of the Apollo spacecraft presented many first-time problems in the areas of test philosophy, operational concepts, and program implementation. The rationale used to resolve these problems is explained and examined critically in view of actual experience. The series of 12 tests involving 1517 hours of chamber operating time resulted in the disclosure of numerous equipment and procedural deficiencies of significance to the flight mission. Test experience and results in view of subsequent flight experience confirmed that thermal-vacuum testing of integrated manned spacecraft provides a feasible, cost-effective, and safe technique with which to obtain maximum confidence in spacecraft flight worthiness early in the program.

Energetic Phenomena on the Sun: The Solar Maximum Mission Flare Workshop. Proceedings

NASA Technical Reports Server (NTRS)

Kundu, Mukul (Editor); Woodgate, Bruce (Editor)

1986-01-01

The general objectives of the conference were as follows: (1) Synthesize flare studies after three years of Solar Maximum Mission (SSM) data analysis. Encourage a broader participation in the SMM data analysis and combine this more fully with theory and other data sources-data obtained with other spacecraft such as the HINOTORI, p78-1, and ISEE-3 spacecrafts, and with the Very Large Array (VLA) and many other ground-based instruments. Many coordinated data sets, unprecedented in their breadth of coverage and multiplicity of sources, had been obtained within the structure of the Solar Maximum Year (SMY). (2) Stimulate joint studies, and publication in the general scientific literature. The intended primary benefit was for informal collaborations to be started or broadened at the Workshops with subsequent publications. (3) Provide a special publication resulting from the Workshop.

Ku-Band rendezvous radar performance computer simulation model

NASA Technical Reports Server (NTRS)

Magnusson, H. G.; Goff, M. F.

1984-01-01

All work performed on the Ku-band rendezvous radar performance computer simulation model program since the release of the preliminary final report is summarized. Developments on the program fall into three distinct categories: (1) modifications to the existing Ku-band radar tracking performance computer model; (2) the addition of a highly accurate, nonrealtime search and acquisition performance computer model to the total software package developed on this program; and (3) development of radar cross section (RCS) computation models for three additional satellites. All changes in the tracking model involved improvements in the automatic gain control (AGC) and the radar signal strength (RSS) computer models. Although the search and acquisition computer models were developed under the auspices of the Hughes Aircraft Company Ku-Band Integrated Radar and Communications Subsystem program office, they have been supplied to NASA as part of the Ku-band radar performance comuter model package. Their purpose is to predict Ku-band acquisition performance for specific satellite targets on specific missions. The RCS models were developed for three satellites: the Long Duration Exposure Facility (LDEF) spacecraft, the Solar Maximum Mission (SMM) spacecraft, and the Space Telescopes.

Ku-Band rendezvous radar performance computer simulation model

NASA Astrophysics Data System (ADS)

Magnusson, H. G.; Goff, M. F.

1984-06-01

All work performed on the Ku-band rendezvous radar performance computer simulation model program since the release of the preliminary final report is summarized. Developments on the program fall into three distinct categories: (1) modifications to the existing Ku-band radar tracking performance computer model; (2) the addition of a highly accurate, nonrealtime search and acquisition performance computer model to the total software package developed on this program; and (3) development of radar cross section (RCS) computation models for three additional satellites. All changes in the tracking model involved improvements in the automatic gain control (AGC) and the radar signal strength (RSS) computer models. Although the search and acquisition computer models were developed under the auspices of the Hughes Aircraft Company Ku-Band Integrated Radar and Communications Subsystem program office, they have been supplied to NASA as part of the Ku-band radar performance comuter model package. Their purpose is to predict Ku-band acquisition performance for specific satellite targets on specific missions. The RCS models were developed for three satellites: the Long Duration Exposure Facility (LDEF) spacecraft, the Solar Maximum Mission (SMM) spacecraft, and the Space Telescopes.

Ceramic insulation/multifoil composite for thermal protection of reentry spacecraft

NASA Technical Reports Server (NTRS)

Pitts, W. C.; Kourtides, D. A.

1989-01-01

A new type of insulation blanket called Composite Flexible Blanket Insulation is proposed for thermal protection of advanced spacecraft in regions where the maximum temperature is not excessive. The blanket is a composite of two proven insulation materials: ceramic insulation blankets from Space Shuttle technology and multilayer insulation blankets from spacecraft thermal control technology. A potential heatshield weight saving of up to 500 g/sq m is predicted. The concept is described; proof of concept experimental data are presented; and a spaceflight experiment to demonstrate its actual performance is discussed.

Materials Advance Chemical Propulsion Technology

NASA Technical Reports Server (NTRS)

2012-01-01

In the future, the Planetary Science Division of NASA's Science Mission Directorate hopes to use better-performing and lower-cost propulsion systems to send rovers, probes, and observers to places like Mars, Jupiter, and Saturn. For such purposes, a new propulsion technology called the Advanced Materials Bipropellant Rocket (AMBR) was developed under NASA's In-Space Propulsion Technology (ISPT) project, located at Glenn Research Center. As an advanced chemical propulsion system, AMBR uses nitrogen tetroxide oxidizer and hydrazine fuel to propel a spacecraft. Based on current research and development efforts, the technology shows great promise for increasing engine operation and engine lifespan, as well as lowering manufacturing costs. In developing AMBR, ISPT has several goals: to decrease the time it takes for a spacecraft to travel to its destination, reduce the cost of making the propulsion system, and lessen the weight of the propulsion system. If goals like these are met, it could result in greater capabilities for in-space science investigations. For example, if the amount (and weight) of propellant required on a spacecraft is reduced, more scientific instruments (and weight) could be added to the spacecraft. To achieve AMBR s maximum potential performance, the engine needed to be capable of operating at extremely high temperatures and pressure. To this end, ISPT required engine chambers made of iridium-coated rhenium (strong, high-temperature metallic elements) that allowed operation at temperatures close to 4,000 F. In addition, ISPT needed an advanced manufacturing technique for better coating methods to increase the strength of the engine chamber without increasing the costs of fabricating the chamber.

Relativistic Spacecraft Propelled by Directed Energy

NASA Astrophysics Data System (ADS)

Kulkarni, Neeraj; Lubin, Philip; Zhang, Qicheng

2018-04-01

Achieving relativistic flight to enable extrasolar exploration is one of the dreams of humanity and the long-term goal of our NASA Starlight program. We derive a relativistic solution for the motion of a spacecraft propelled by radiation pressure from a directed energy (DE) system. Depending on the system parameters, low-mass spacecraft can achieve relativistic speeds, thus enabling interstellar exploration. The diffraction of the DE system plays an important role and limits the maximum speed of the spacecraft. We consider “photon recycling” as a possible method to achieving higher speeds. We also discuss recent claims that our previous work on this topic is incorrect and show that these claims arise from an improper treatment of causality.

Mars Observer trajectory and orbit design

NASA Technical Reports Server (NTRS)

Beerer, Joseph G.; Roncoli, Ralph B.

1991-01-01

The Mars Observer launch, interplanetary, Mars orbit insertion, and mapping orbit designs are described. The design objective is to enable a near-maximum spacecraft mass to be placed in orbit about Mars. This is accomplished by keeping spacecraft propellant requirements to a minimum, selecting a minimum acceptable launch period, equalizing the spacecraft velocity change requirement at the beginning and end of the launch period, and constraining the orbit insertion maneuvers to be coplanar. The mapping orbit design objective is to provide the opportunity for global observation of the planet by the science instruments while facilitating the spacecraft design. This is realized with a sun-synchronous near-polar orbit whose ground-track pattern covers the planet at progressively finer resolution.

Fire safety arrangement of inhabited pressurized compartments of manned spacecraft

NASA Astrophysics Data System (ADS)

Bolodian, Ivan; Melikhov, Anatoliy; Tanklevskiy, Leonid

2017-06-01

The article deals with innovative technical solutions that provide fire safety in inhabited pressurized compartments of manned spacecraft by means of a fireproof device of inhabited pressurized compartments via application of engineering means of fire prevention and fire spreading prevention by lowering fire load in an inhabited pressurized module up to the point when the maximum possible levels of fire factors in an inhabited pressurized compartment of a manned spacecraft are prevented. Represented technical solutions are used at the present time according to stated recommendations during provision of fire safety of equipment created by a number of Russian organizations for equipage of inhabited pressurized compartments of spacecraft of the Russian segment of International space station.

Optical effects produced by running onboard engines of low-earth-orbit spacecraft

NASA Astrophysics Data System (ADS)

Beletskiy, A. B.; Mihalev, A. V.; Hahinov, V. V.; Lebedev, V. P.

2016-12-01

This paper presents results of optical observations made during Radar-Progress Experiment performed on April 17, 2013 and July 30, 2014 after approach-correction engines (ACE) of Progress M-17M and Progress M-23M cargo spacecraft in the thermosphere had been started. A region of enhanced emission intensity was recorded during engine operation. This may have been related to the scatter of twilight solar emission along the cargo spacecraft exhaust and to the emergence of additional atomic oxygen [OI] emission at 630 nm. The maximum dimension of the observed emission region was ~330-350 km and ~250-270 km along and across the orbit respectively. For the first time after ACE had been started, an expansion rate of emission region was ~ 7 and ~ 3.5 km/s along and across the orbit respectively. The maximum intensity of the disturbance area for Progress M-17M is estimated as ~40-60 R at 2 nm. Progress M-23M Space Experiment recorded a minor disturbance of atmospheric [OI] 630.0 nm emissions, both in near and in far cargo spacecraft flight paths, which might have been associated with the ACE exhaust gas injection.

Deployment/retraction mechanism for solar maximum mission high gain antenna system

NASA Technical Reports Server (NTRS)

Bennett, N.; Preiswerk, P.

1979-01-01

A mechanism called a deployment/retraction assembly (DRA) which provides not only a stable, but a deployable platform for the high gain antenna system (HGAS) aboard the Solar Maximum Mission (SMM) spacecraft is described. The DRA also has the capability to retract the system upon command.

Distributioin, orientation and scales of the field-aligned currents measured by Swarm

NASA Astrophysics Data System (ADS)

Yang, J.; Dunlop, M. W.

2016-12-01

We have statistically studied the R1, R2 and net field aligned currents using the FAC data of the Swarm satellites. We also have investigated the statistical, dual-spacecraft correlations of field-aligned current signatures between two Swarm spacecraft (A and C). For the first time we have inferred the orientations of the current sheets of FACs directly, using the maximum correlations, obtained from sliding data segments, which show clear trends in magnetic local time (MLT). To compare with this we also check the MVAB method. To explore the scale and variability of the current sheet supposition, we investigate the MLT dependence of the maximum correlations in different time shift or longitude shift bins.

MEMS Louvers for Thermal Control

NASA Technical Reports Server (NTRS)

Champion, J. L.; Osiander, R.; Darrin, M. A. Garrison; Swanson, T. D.

1998-01-01

Mechanical louvers have frequently been used for spacecraft and instrument thermal control purposes. These devices typically consist of parallel or radial vanes, which can be opened or closed to vary the effective emissivity of the underlying surface. This project demonstrates the feasibility of using Micro-Electromechanical Systems (MEMS) technology to miniaturize louvers for such purposes. This concept offers the possibility of substituting the smaller, lighter weight, more rugged, and less costly MEMS devices for such mechanical louvers. In effect, a smart skin that self adjusts in response to environmental influences could be developed composed of arrays of thousands of miniaturized louvers. Several orders of magnitude size, weight, and volume decreases are potentially achieved using micro-electromechanical techniques. The use of this technology offers substantial benefits in spacecraft/instrument design, integration and testing, and flight operations. It will be particularly beneficial for the emerging smaller spacecraft and instruments of the future. In addition, this MEMS thermal louver technology can form the basis for related spacecraft instrument applications. The specific goal of this effort was to develop a preliminary MEMS device capable of modulating the effective emissivity of radiators on spacecraft. The concept pursued uses hinged panels, or louvers, in a manner such that heat emitted from the radiators is a function of louver angle. An electrostatic comb drive or other such actuator can control the louver position. The initial design calls for the louvers to be gold coated while the underlying surface is of high emissivity. Since, the base MEMS material, silicon, is transparent in the InfraRed (IR) spectrum, the device has a minimum emissivity when closed and a maximum emissivity when open. An initial set of polysilicon louver devices was designed at the Johns Hopkins Applied Physics Laboratory in conjunction with the Thermal Engineering Branch at NASA's Goddard Space Flight Center.

Heat Pipe Thermal Conditioning Panel

NASA Technical Reports Server (NTRS)

Saaski, E. W.

1973-01-01

The technology involved in designing and fabricating a heat pipe thermal conditioning panel to satisfy a broad range of thermal control system requirements on NASA spacecraft is discussed. The design specifications were developed for a 30 by 30 inch heat pipe panel. The fundamental constraint was a maximum of 15 gradient from source to sink at 300 watts input and a flux density of 2 watts per square inch. The results of the performance tests conducted on the panel are analyzed.

Spacecraft Station-Keeping Trajectory and Mission Design Tools

NASA Technical Reports Server (NTRS)

Chung, Min-Kun J.

2009-01-01

Two tools were developed for designing station-keeping trajectories and estimating delta-v requirements for designing missions to a small body such as a comet or asteroid. This innovation uses NPOPT, a non-sparse, general-purpose sequential quadratic programming (SQP) optimizer and the Two-Level Differential Corrector (T-LDC) in LTool (Libration point mission design Tool) to design three kinds of station-keeping scripts: vertical hovering, horizontal hovering, and orbiting. The T-LDC is used to differentially correct several trajectory legs that join hovering points. In a vertical hovering, the maximum and minimum range points must be connected smoothly while maintaining the spacecrafts range from a small body, all within the law of gravity and the solar radiation pressure. The same is true for a horizontal hover. A PatchPoint is an LTool class that denotes a space-time event with some extra information for differential correction, including a set of constraints to be satisfied by T-LDC. Given a set of PatchPoints, each with its own constraint, the T-LDC differentially corrects the entire trajectory by connecting each trajectory leg joined by PatchPoints while satisfying all specified constraints at the same time. Vertical and horizontal hover both are needed to minimize delta-v spent for station keeping. A Python I/F to NPOPT has been written to be used from an LTool script. In vertical hovering, the spacecraft stays along the line joining the Sun and a small body. An instantaneous delta-v toward the anti- Sun direction is applied at the closest approach to the small body for station keeping. For example, the spacecraft hovers between the minimum range (2 km) point and the maximum range (2.5 km) point from the asteroid 1989ML. Horizontal hovering buys more time for a spacecraft to recover if, for any reason, a planned thrust fails, by returning almost to the initial position after some time later via a near elliptical orbit around the small body. The mapping or staging orbit may be similarly generated using T-LDC with a set of constraints. Some delta-v tables are generated for several different asteroid masses.

Risk Analysis of On-Orbit Spacecraft Refueling Concepts

NASA Technical Reports Server (NTRS)

Cirillo, William M.; Stromgren, Chel; Cates, Grant R.

2010-01-01

On-orbit refueling of spacecraft has been proposed as an alternative to the exclusive use of Heavy-lift Launch Vehicles to enable human exploration beyond Low Earth Orbit (LEO). In these scenarios, beyond LEO spacecraft are launched dry (without propellant) or partially dry into orbit, using smaller or fewer element launch vehicles. Propellant is then launched into LEO on separate launch vehicles and transferred to the spacecraft. Refueling concepts are potentially attractive because they reduce the maximum individual payload that must be placed in Earth orbit. However, these types of approaches add significant complexity to mission operations and introduce more uncertainty and opportunities for failure to the mission. In order to evaluate these complex scenarios, the authors developed a Monte Carlo based discrete-event model that simulates the operational risks involved with such strategies, including launch processing delays, transportation system failures, and onorbit element lifetimes. This paper describes the methodology used to simulate the mission risks for refueling concepts, the strategies that were evaluated, and the results of the investigation. The results of the investigation show that scenarios that employ refueling concepts will likely have to include long launch and assembly timelines, as well as the use of spare tanker launch vehicles, in order to achieve high levels of mission success through Trans Lunar Injection.

Effort to recover SOHO spacecraft continue as investigation board focuses on most likely causes

NASA Astrophysics Data System (ADS)

1998-07-01

Meanwhile, the ESA/NASA investigation board concentrates its inquiry on three errors that appear to have led to the interruption of communications with SOHO on June 25. Officials remain hopeful that, based on ESA's successful recovery of the Olympus spacecraft after four weeks under similar conditions in 1991, recovery of SOHO may be possible. The SOHO Mission Interruption Joint ESA/NASA Investigation Board has determined that the first two errors were contained in preprogrammed command sequences executed on ground system computers, while the last error was a decision to send a command to the spacecraft in response to unexpected telemetry readings. The spacecraft is controlled by the Flight Operations Team, based at NASA's Goddard Space Flight Center, Greenbelt, MD. The first error was in a preprogrammed command sequence that lacked a command to enable an on-board software function designed to activate a gyro needed for control in Emergency Sun Reacquisition (ESR) mode. ESR mode is entered by the spacecraft in the event of anomalies. The second error, which was in a different preprogrammed command sequence, resulted in incorrect readings from one of the spacecraft's three gyroscopes, which in turn triggered an ESR. At the current stage of the investigation, the board believes that the two anomalous command sequences, in combination with a decision to send a command to SOHO to turn off a gyro in response to unexpected telemetry values, caused the spacecraft to enter a series of ESRs, and ultimately led to the loss of control. The efforts of the investigation board are now directed at identifying the circumstances that led to the errors, and at developing a recovery plan should efforts to regain contact with the spacecraft succeed. ESA and NASA engineers believe the spacecraft is currently spinning with its solar panels nearly edge-on towards the Sun, and thus not generating any power. Since the spacecraft is spinning around a fixed axis, as the spacecraft progresses in its orbit around the Sun, the orientation of the panels with respect to the Sun should gradually change. The orbit of the spacecraft and the seasonal change in the spacecraft-Sun alignment should result in the increased solar illumination of the spacecraft solar arrays over the next few months. The engineers predict that in late September 1998, illumination of the solar arrays and, consequently, power supplied to the spacecraft, should approach a maximum. The probability of successfully establishing contact reaches a maximum at this point. After this time, illumination of the solar arrays gradually diminishes as the spacecraft-Sun alignment continues to change. In an attempt to recover SOHO as soon as possible, the Flight Operations Team is uplinking commands to the spacecraft via NASA's Deep Space Network, managed by NASA's Jet Propulsion Laboratory, Pasadena, CA, approximately 12 hours per day with no success to date. A recovery plan is under development by ESA and NASA to provide for orderly restart of the spacecraft and to mitigate risks involved. The recovery of the Olympus spacecraft by ESA in 1991 under similar conditions leads to optimism that the SOHO spacecraft may be recoverable once contact is re-established. In May 1991, ESA's Olympus telecommunications satellite experienced a similar major anomaly which resulted in the loss of attitude, leading to intermittent power availability. As a consequence, there was inadequate communication, and the batteries and fuel froze. From analysis of the data available prior to the loss, there was confidence that the power situation would improve over the coming months. A recovery plan was prepared, supported by laboratory tests, to assess the characteristics of thawing batteries and propellants. Telecommand access of Olympus was regained four weeks later, and batteries and propellant tanks were thawed out progressively over the next four weeks. The attitude was then fully recovered and the payload switched back on three months after the incident. Equipment damage was sustained as a result of the low temperatures, but nothing significant enough to prevent the successful resumption of the mission. The experience of Olympus is being applied, where possible, to SOHO and increases the hope of also recovering this mission. Estimating the probability of recovery is made difficult by a number of unknown spacecraft conditions. Like Olympus, the hydrazine fuel and batteries may be frozen. Thermal stress may have damaged some of the scientific instruments as well. If the rate of spin is excessive, there may have been structural damage. SOHO engineers can reliably predict the spacecraft's orbit through November 1998. After that time, the long-term orbital behavior becomes dependent on the initial velocity conditions of the spacecraft at the time of the telemetry loss. These are not known precisely, due to spacecraft thruster activity that continued after loss of telemetry, so orbital prediction becomes very difficult.

Evaluation of simple deployment mechanism of multiple rovers by microgravity experiments using a drop tower

NASA Astrophysics Data System (ADS)

Yoshimitsu, Tetsuo; Yano, Hajime; Kubota, Takashi; Adachi, Tadashi; Ishigami, Genya

2012-07-01

Introduction, Japan has announced the official development of ``Hayabusa-2'', the second sample return mission to a Near-Earth asteroid. When the development is made smoothly, Hayabusa-2 will be launched in 2014. The predecessor spacecraft ``Hayabusa'' made a great success when it returned to the Earth in June 2010 with a capsule containing some particles obtained from S-type asteroid ``Itokawa.'' Rover system, The authors installed a tiny hopping rover called ``MINERVA'' into Hayabusa spacecraft. MINERVA weights only 591[g] but has an autonomous exploration capability on the microgravity environment on the small solar system bodies. MINERVA was successfully deployed from the mother spacecraft on 12 Nov 2005 at the vicinity of the target asteroid. But unfortunately it became a solar orbiting satellite since the relative position and the speed of the mother spacecraft around the target asteroid were worst. Nevertheless it worked well, demonstrating an autnomous capability and had survived until the comunication link was lost. The authors plan to install some rovers also into Hayabusa-2. The total concept is the same but this time multiple rovers are considered. Deployment mechanism, Two rovers are installed in one container and are developed at the same time. The maximum allowed weight for the container including two rovers is 2.5[kg] and we have to seek for a simple and a light-weighted deployment system. We developed a new deployment system drastically sophisticated from the one used for MINERVA in Hayabusa mission. Both the cover and the rovers are pushed by the springs after the tightly winded wire has been cut by the deployment trigger form the spacecraft. The new deployment system enables the following things. The cover and the rovers are deployed in different directions in one action. The uncertainty of the deployment speed is decreased. Microgravity experiment, Thanks to the courtesy of DLR (German Aerospace Center) based on the international cooperation between Germany and Japan on the development of Hayabusa-2 spacecraft, we had an opprotunity to evaluate the performance of the above mentioned deployment system by using the drop tower in Bremen. We made six drops in January 2012 to look at the speed and the direction of the cover and the rovers using various parameters, which were all performed successfully. This paper reports the result of the microgravity experiments.

Small Body GN and C Research Report: G-SAMPLE - An In-Flight Dynamical Method for Identifying Sample Mass [External Release Version

NASA Technical Reports Server (NTRS)

Carson, John M., III; Bayard, David S.

2006-01-01

G-SAMPLE is an in-flight dynamical method for use by sample collection missions to identify the presence and quantity of collected sample material. The G-SAMPLE method implements a maximum-likelihood estimator to identify the collected sample mass, based on onboard force sensor measurements, thruster firings, and a dynamics model of the spacecraft. With G-SAMPLE, sample mass identification becomes a computation rather than an extra hardware requirement; the added cost of cameras or other sensors for sample mass detection is avoided. Realistic simulation examples are provided for a spacecraft configuration with a sample collection device mounted on the end of an extended boom. In one representative example, a 1000 gram sample mass is estimated to within 110 grams (95% confidence) under realistic assumptions of thruster profile error, spacecraft parameter uncertainty, and sensor noise. For convenience to future mission design, an overall sample-mass estimation error budget is developed to approximate the effect of model uncertainty, sensor noise, data rate, and thrust profile error on the expected estimate of collected sample mass.

Automated maneuver planning using a fuzzy logic algorithm

NASA Technical Reports Server (NTRS)

Conway, D.; Sperling, R.; Folta, D.; Richon, K.; Defazio, R.

1994-01-01

Spacecraft orbital control requires intensive interaction between the analyst and the system used to model the spacecraft trajectory. For orbits with right mission constraints and a large number of maneuvers, this interaction is difficult or expensive to accomplish in a timely manner. Some automation of maneuver planning can reduce these difficulties for maneuver-intensive missions. One approach to this automation is to use fuzzy logic in the control mechanism. Such a prototype system currently under development is discussed. The Tropical Rainfall Measurement Mission (TRMM) is one of several missions that could benefit from automated maneuver planning. TRMM is scheduled for launch in August 1997. The spacecraft is to be maintained in a 350-km circular orbit throughout the 3-year lifetime of the mission, with very small variations in this orbit allowed. Since solar maximum will occur as early as 1999, the solar activity during the TRMM mission will be increasing. The increasing solar activity will result in orbital maneuvers being performed as often as every other day. The results of automated maneuver planning for the TRMM mission will be presented to demonstrate the prototype of the fuzzy logic tool.

Preliminary design of the thermal protection system for solar probe

NASA Technical Reports Server (NTRS)

Dirling, R. B., Jr.; Loomis, W. C.; Heightland, C. N.

1982-01-01

A preliminary design of the thermal protection system for the NASA Solar Probe spacecraft is presented. As presently conceived, the spacecraft will be launched by the Space Shuttle on a Jovian swing-by trajectory and at perihelion approach to three solar radii of the surface of the Earth's sun. The system design satisfies maximum envelope, structural integrity, equipotential, and mass loss/contamination requirements by employing lightweight carbon-carbon emissive shields. The primary shield is a thin shell, 15.5-deg half-angle cone which absorbs direct solar flux at up to 10-deg off-nadir spacecraft pointing angles. Secondary shields of sandwich construction and low thickness-direction thermal conductivity are used to reduce the primary shield infrared radiation to the spacecraft payload.

Design, Integration, Certification and Testing of the Orion Crew Module Propulsion System

NASA Technical Reports Server (NTRS)

McKay, Heather; Freeman, Rich; Cain, George; Albright, John D.; Schoenberg, Rich; Delventhal, Rex

2014-01-01

The Orion Multipurpose Crew Vehicle (MPCV) is NASA's next generation spacecraft for human exploration of deep space. Lockheed Martin is the prime contractor for the design, development, qualification and integration of the vehicle. A key component of the Orion Crew Module (CM) is the Propulsion Reaction Control System, a high-flow hydrazine system used during re-entry to orient the vehicle for landing. The system consists of a completely redundant helium (GHe) pressurization system and hydrazine fuel system with monopropellant thrusters. The propulsion system has been designed, integrated, and qualification tested in support of the Orion program's first orbital flight test, Exploration Flight Test One (EFT-1), scheduled for 2014. A subset of the development challenges and lessons learned from this first flight test campaign will be discussed in this paper for consideration when designing future spacecraft propulsion systems. The CONOPS and human rating requirements of the CM propulsion system are unique when compared with a typical satellite propulsion reaction control system. The system requires a high maximum fuel flow rate. It must operate at both vacuum and sea level atmospheric pressure conditions. In order to meet Orion's human rating requirements, multiple parts of the system must be redundant, and capable of functioning after spacecraft system fault events.

System design of the Pioneer Venus spacecraft. Volume 9: Attitude control/mechanisms subsystems studies

NASA Technical Reports Server (NTRS)

Neil, A. L.

1973-01-01

The Pioneer Venus mission study was conducted for a probe spacecraft and an orbiter spacecraft to be launched by either a Thor/Delta or an Atlas/Centaur launch vehicle. Both spacecraft are spin stabilized. The spin speed is controlled by ground commands to as low as 5 rpm for science instrument scanning on the orbiter and as high as 71 rpm for small probes released from the probe bus. A major objective in the design of the attitude control and mechanism subsystem (ACMS) was to provide, in the interest of costs, maximum commonality of the elements between the probe bus and orbiter spacecraft configurations. This design study was made considering the use of either launch vehicle. The basic functional requirements of the ACMS are derived from spin axis pointing and spin speed control requirements implicit in the acquisition, cruise, encounter and orbital phases of the Pioneer Venus missions.

Redundancy for electric motors in spacecraft applications

NASA Technical Reports Server (NTRS)

Smith, Robert J.; Flew, Alastair R.

1986-01-01

The parts of electric motors which should be duplicated in order to provide maximum reliability in spacecraft application are identified. Various common types of redundancy are described. The advantages and disadvantages of each are noted. The principal types are illustrated by reference to specific examples. For each example, constructional details, basic performance data and failure modes are described, together with a discussion of the suitability of particular redundancy techniques to motor types.

The quasi-inertial and wide-deadband modes as backup attitude options for the Skylab mission

NASA Technical Reports Server (NTRS)

Elrod, B. D.

1971-01-01

The quasi-inertial (QI) and wide deadband (WDB) modes were investigated as alternatives to the solar inertial (SI) mode in case two control moment gyros fail during the Skylab mission. Both modes provide a substantial reduction in propellant requirements from the solar interial hold requirement with either the orbital assembly/thruster attitude control system or service module reaction control system. Spacecraft motion in the QI mode is produced by a command rate and results in a small amplitude oscillation (17 deg, maximum) about the SI orientation. In the WDB mode a somewhat similar, but larger amplitude motion (35 deg maximum) about the SI orientation is developed by appropriate choice of controller deadbands and switch line slopes.

SMART-OLEV—An orbital life extension vehicle for servicing commercial spacecrafts in GEO

NASA Astrophysics Data System (ADS)

Kaiser, Clemens; Sjöberg, Fredrik; Delcura, Juan Manuel; Eilertsen, Baard

2008-07-01

Orbital Satellite Services Limited (OSSL) is a satellite servicing company that is developing an orbit life extension vehicle (OLEV) to extend the operational lifetime of geostationary satellites. The industrial consortium of SSC (Sweden), Kayser-Threde (Germany) and Sener (Spain) is in charge to develop and industrialize the space and ground segment. It is a fully commercial program with support of several space agencies during the development phase. The business plan is based on life extension for high value commercial satellites while also providing the satellite operators with various fleet management services such as graveyard burns, slot transfers and on orbit protection against replacement satellite or launch failures. The OLEV spacecraft will be able to dock with a geostationary satellite and uses an electrical propulsion system to extend its life by taking over the attitude control and station keeping functions. The OLEV system is building on the SMART-1 platform developed by Swedish Space Corporation. It was developed for ESA as a technology test-bed to demonstrate the use of electrical propulsion for interplanetary orbit transfer manoeuvres. The concept is called SMART-OLEV and takes advantage of the low cost, low mass SMART-1 platform by a maximum use of recurrent platform technology.

Power management and control for space systems

NASA Technical Reports Server (NTRS)

Finke, R. C.; Myers, I. T.; Terdan, F. F.; Stevens, N. J.

1978-01-01

Power management and control technology for the large, high-power spacecraft of the 1980's is discussed. Systems weight optimization that indicate a need for higher bus voltages are shown. Environmental interactions that are practical limits for the maximum potential on exposed surfaces are shown. A dual-voltage system is proposed that would provide the weight savings of a high-voltage distribution system and take into account the potential environmental interactions. The technology development of new components and circuits is also discussed.

A depolarization and attenuation experiment using the CTS satellite. Volume 1: Experiment description

NASA Technical Reports Server (NTRS)

Bostian, C. W.; Holt, S. B., Jr.; Kauffman, S. R.; Manus, E. A.; Marshall, R. E.; Stutzman, W. L.; Wiley, P. H.

1976-01-01

An experiment for measuring precipitation attenuation and depolarization on the Communications Technology Satellite (CTS) 11.7 GHz downlink is described. Attenuation and depolarization of the signal received from the spacecraft is monitored on a 24 hour basis. Data is correlated with ground weather conditions. Theoretical models for millimeter wave propagation through rain are refined for maximum agreement with observed data. Techniques are developed for predicting and mimimizing the effects of rain scatter and depolarization on future satellite communication systems.

Imaging Flash Lidar for Autonomous Safe Landing and Spacecraft Proximity Operation

NASA Technical Reports Server (NTRS)

Amzajerdian, Farzin; Roback, Vincent E.; Brewster, Paul F.; Hines, Glenn D.; Bulyshev, Alexander E.

2016-01-01

3-D Imaging flash lidar is recognized as a primary candidate sensor for safe precision landing on solar system bodies (Moon, Mars, Jupiter and Saturn moons, etc.), and autonomous rendezvous proximity operations and docking/capture necessary for asteroid sample return and redirect missions, spacecraft docking, satellite servicing, and space debris removal. During the final stages of landing, from about 1 km to 500 m above the ground, the flash lidar can generate 3-Dimensional images of the terrain to identify hazardous features such as craters, rocks, and steep slopes. The onboard fli1ght computer can then use the 3-D map of terrain to guide the vehicle to a safe location. As an automated rendezvous and docking sensor, the flash lidar can provide relative range, velocity, and bearing from an approaching spacecraft to another spacecraft or a space station from several kilometers distance. NASA Langley Research Center has developed and demonstrated a flash lidar sensor system capable of generating 16k pixels range images with 7 cm precision, at a 20 Hz frame rate, from a maximum slant range of 1800 m from the target area. This paper describes the lidar instrument design and capabilities as demonstrated by the closed-loop flight tests onboard a rocket-propelled free-flyer vehicle (Morpheus). Then a plan for continued advancement of the flash lidar technology will be explained. This proposed plan is aimed at the development of a common sensor that with a modest design adjustment can meet the needs of both landing and proximity operation and docking applications.

Metrics on the relative spacecraft motion invariant manifold.

PubMed

Gurfil, P; Kholshevnikov, Konstantin V

2005-12-01

This paper establishes a methodology for obtaining the general solution to the spacecraft relative motion problem by utilizing Cartesian configuration space in conjunction with classical orbital elements. The geometry of the relative motion configuration space is analyzed, and the relative motion invariant manifold is determined. Most importantly, the geometric structure of the relative motion problem is used to derive useful metrics for quantification of the minimum, maximum, and mean distance between spacecraft for commensurable and non-commensurable mean motions. A number of analytic solutions, as well as useful examples, are provided, illustrating the calculated bounds. A few particular cases are given that yield simple solutions.

Comparison of candidate solar array maximum power utilization approaches. [for spacecraft propulsion

NASA Technical Reports Server (NTRS)

Costogue, E. N.; Lindena, S.

1976-01-01

A study was made of five potential approaches that can be utilized to detect the maximum power point of a solar array while sustaining operations at or near maximum power and without endangering stability or causing array voltage collapse. The approaches studied included: (1) dynamic impedance comparator, (2) reference array measurement, (3) onset of solar array voltage collapse detection, (4) parallel tracker, and (5) direct measurement. The study analyzed the feasibility and adaptability of these approaches to a future solar electric propulsion (SEP) mission, and, specifically, to a comet rendezvous mission. Such missions presented the most challenging requirements to a spacecraft power subsystem in terms of power management over large solar intensity ranges of 1.0 to 3.5 AU. The dynamic impedance approach was found to have the highest figure of merit, and the reference array approach followed closely behind. The results are applicable to terrestrial solar power systems as well as to other than SEP space missions.

Wind Observations of Anomalous Cosmic Rays from Solar Minimum to Maximum

NASA Technical Reports Server (NTRS)

Reames, D. V.; McDonald, F. B.

2003-01-01

We report the first observation near Earth of the time behavior of anomalous cosmic-ray N, O, and Ne ions through the period surrounding the maximum of the solar cycle. These observations were made by the Wind spacecraft during the 1995-2002 period spanning times from solar minimum through solar maximum. Comparison of anomalous and galactic cosmic rays provides a powerful tool for the study of the physics of solar modulation throughout the solar cycle.

Adverse Space Weather at the Solar Cycle Minimum

NASA Astrophysics Data System (ADS)

Baker, D. N.; Kanekal, S. G.; McCollough, J. P.; Singer, H. J.; Chappell, S. P.; Allen, J. H.

2008-05-01

It is commonly understood that many types of adverse space weather (solar flares, coronal mass ejections, geomagnetic storms) occur most commonly around the maximum of the 11-year sunspot activity cycle. Other types of well-known space weather such as relativistic electron events in the Earth's outer magnetosphere (that produce deep dielectric charging in spacecraft systems) are usually associated with the period just after sunspot maximum. At the present time, we are in the very lowest activity phase of the sunspot cycle (solar minimum). As such we would not expect much in the way of adverse space weather events. However, in early to mid-February of 2008 quite prominent solar coronal holes produced two high-speed streams that in turn stimulated very large, long-duration relativistic electron enhancements in Earth's magnetosphere. These seem to have been associated with several spacecraft operational anomalies at various spacecraft orbital locations. We describe these recent space weather events and assess their operational significance in this presentation. These results show that substantial space weather events can and do occur even during the quietest parts of the solar cycle.

Toxicological approach to setting spacecraft maximum allowable concentrations for carbon monoxide

NASA Technical Reports Server (NTRS)

Wong, K. L.; Limero, T. F.; James, J. T.

1992-01-01

The Spacecraft Maximum Allowable Concentrations (SMACs) are exposure limits for airborne chemicals used by NASA in spacecraft. The aim of these SMACs is to protect the spacecrew against adverse health effects and performance decrements that would interfere with mission objectives. Because of the 1 and 24 hr SMACs are set for contingencies, minor reversible toxic effects that do not affect mission objectives are acceptable. The 7, 30, or 180 day SMACs are aimed at nominal operations, so they are established at levels that would not cause noncarcinogenic toxic effects and more than one case of tumor per 1000 exposed individuals over the background. The process used to set the SMACs for carbon monoxide (CO) is described to illustrate the approach used by NASA. After the toxicological literature on CO was reviewed, the data were summarized and separated into acute, subchronic, and chronic toxicity data. CO's toxicity depends on the formation of carboxyhemoglobin (COHb) in the blood, reducing the blood's oxygen carrying capacity. The initial task was to estimate the COHb levels that would not produce toxic effects in the brain and heart.

Development of a Solar Array Drive Assembly for CubeSat

NASA Technical Reports Server (NTRS)

Passaretti, Mike; Hayes, Ron

2010-01-01

Small satellites and in particular CubeSats, have increasingly become more viable as platforms for payloads typically requiring much larger bus structures. As advances in technology make payloads and instruments for space missions smaller, lighter and more power efficient, a niche market is emerging from the university community to perform rapidly developed, low-cost missions on very small spacecraft - micro, nano, and picosatellites. In just the last few years, imaging, biological and new technology demonstration missions have been either proposed or have flown using variations of the CubeSat structure as a basis. As these missions have become more complex, and the CubeSat standard has increased in both size (number of cubes) and mass, available power has become an issue. Body-mounted solar cells provide a minimal amount of power; deployable arrays improve on that baseline but are still limited. To truly achieve maximum power, deployed tracked arrays are necessary. To this end, Honeybee Robotics Spacecraft Mechanisms Corporation, along with MMA of Nederland Colorado, has developed a solar array drive assembly (SADA) and deployable solar arrays specifically for CubeSat missions. In this paper, we discuss the development of the SADA.

Comparison of polar cap potential drops estimated from solar wind and ground magnetometer data - CDAW 6

NASA Technical Reports Server (NTRS)

Reiff, P. H.; Spiro, R. W.; Wolf, R. A.; Kamide, Y.; King, J. H.

1985-01-01

It is pointed out that the maximum electrostatic potential difference across the polar cap, Phi, is a fundamental measure of the coupling between the solar wind and the earth's magnetosphere/ionosphere sytem. During the Coordinated Data Analysis Workshop (CDAW) 6 intervals, no suitably instrumented spacecraft was in an appropriate orbit to determine the polar-cap potential drop directly. However, two recently developed independent techniques make it possible to estimate the polar-cap potential drop for times when direct spacecraft data are not available. The present investigation is concerned with a comparison of cross-polar-cap potential drop estimates calculated for the two CDAW 6 intervals on the basis of these two techniques. In the case of one interval, the agreement between the potential drops and Joule heating rates is relatively good. In the second interval, however, the agreement is not very good. Explanations for this discrepancy are discussed.

Thermal Management Architecture for Future Responsive Spacecraft

NASA Astrophysics Data System (ADS)

Bugby, D.; Zimbeck, W.; Kroliczek, E.

2009-03-01

This paper describes a novel thermal design architecture that enables satellites to be conceived, configured, launched, and operationally deployed very quickly. The architecture has been given the acronym SMARTS for Satellite Modular and Reconfigurable Thermal System and it involves four basic design rules: modest radiator oversizing, maximum external insulation, internal isothermalization and radiator heat flow modulation. The SMARTS philosophy is being developed in support of the DoD Operationally Responsive Space (ORS) initiative which seeks to drastically improve small satellite adaptability, deployability, and design flexibility. To illustrate the benefits of the philosophy for a prototypical multi-paneled small satellite, the paper describes a SMARTS thermal control system implementation that uses: panel-to-panel heat conduction, intra-panel heat pipe isothermalization, radiator heat flow modulation via a thermoelectric cooler (TEC) cold-biased loop heat pipe (LHP) and maximum external multi-layer insulation (MLI). Analyses are presented that compare the traditional "cold-biasing plus heater power" passive thermal design approach to the SMARTS approach. Plans for a 3-panel SMARTS thermal test bed are described. Ultimately, the goal is to incorporate SMARTS into the design of future ORS satellites, but it is also possible that some aspects of SMARTS technology could be used to improve the responsiveness of future NASA spacecraft. [22 CFR 125.4(b)(13) applicable

Dynamic Method for Identifying Collected Sample Mass

NASA Technical Reports Server (NTRS)

Carson, John

2008-01-01

G-Sample is designed for sample collection missions to identify the presence and quantity of sample material gathered by spacecraft equipped with end effectors. The software method uses a maximum-likelihood estimator to identify the collected sample's mass based on onboard force-sensor measurements, thruster firings, and a dynamics model of the spacecraft. This makes sample mass identification a computation rather than a process requiring additional hardware. Simulation examples of G-Sample are provided for spacecraft model configurations with a sample collection device mounted on the end of an extended boom. In the absence of thrust knowledge errors, the results indicate that G-Sample can identify the amount of collected sample mass to within 10 grams (with 95-percent confidence) by using a force sensor with a noise and quantization floor of 50 micrometers. These results hold even in the presence of realistic parametric uncertainty in actual spacecraft inertia, center-of-mass offset, and first flexibility modes. Thrust profile knowledge is shown to be a dominant sensitivity for G-Sample, entering in a nearly one-to-one relationship with the final mass estimation error. This means thrust profiles should be well characterized with onboard accelerometers prior to sample collection. An overall sample-mass estimation error budget has been developed to approximate the effect of model uncertainty, sensor noise, data rate, and thrust profile error on the expected estimate of collected sample mass.

L^1 -optimality conditions for the circular restricted three-body problem

NASA Astrophysics Data System (ADS)

Chen, Zheng

2016-11-01

In this paper, the L^1 -minimization for the translational motion of a spacecraft in the circular restricted three-body problem (CRTBP) is considered. Necessary conditions are derived by using the Pontryagin Maximum Principle (PMP), revealing the existence of bang-bang and singular controls. Singular extremals are analyzed, recalling the existence of the Fuller phenomenon according to the theories developed in (Marchal in J Optim Theory Appl 11(5):441-486, 1973; Zelikin and Borisov in Theory of Chattering Control with Applications to Astronautics, Robotics, Economics, and Engineering. Birkhäuser, Basal 1994; in J Math Sci 114(3):1227-1344, 2003). The sufficient optimality conditions for the L^1 -minimization problem with fixed endpoints have been developed in (Chen et al. in SIAM J Control Optim 54(3):1245-1265, 2016). In the current paper, we establish second-order conditions for optimal control problems with more general final conditions defined by a smooth submanifold target. In addition, the numerical implementation to check these optimality conditions is given. Finally, approximating the Earth-Moon-Spacecraft system by the CRTBP, an L^1 -minimization trajectory for the translational motion of a spacecraft is computed by combining a shooting method with a continuation method in (Caillau et al. in Celest Mech Dyn Astron 114:137-150, 2012; Caillau and Daoud in SIAM J Control Optim 50(6):3178-3202, 2012). The local optimality of the computed trajectory is asserted thanks to the second-order optimality conditions developed.

The Hughes HS601HP spacecraft power subsystem

DOE Office of Scientific and Technical Information (OSTI.GOV)

Krummann, W.; Ayvazian, H.

1998-07-01

The introduction of the Hughes HS 601HP (high power) spacecraft product line continuous the highly successful HS601 three axis stabilized geosynchronus spacecraft with increased power capabilities for larger payload applications. The enhanced power capabilities of the HS 601HP are built upon the heritage of 29 HS601 spacecraft presently in operation. The HS 601HP accommodates payload power ranges of 3 to 7 kilowatts and provides a smooth transition from the lower power HS 601 spacecraft to the HS 702 spacecraft, which has a payload capability up to 13 kilowatts. The HS 601HP spacecraft is designed for a 15 year life withmore » minimal operator interaction. The HS 601HP power subsystem provides a regulated power bus with a voltage range of 52 to 53 volts during all operational phases. The power subsystem is tailored to the specific needs of the spacecraft by selecting standard products from the HS 601HP power catalog. The solar arrays, battery, power control electronics and power distribution electronics are all modular and configurable to the requirements of the spacecraft. The HS 601HP solar array is the primary power source for the spacecraft. The solar array is comprised of two sets of planar solar panels (solar wings) which track the sun in a single spacecraft axis. The solar cells are selected from three different types based upon the spacecraft power generation requirements; silicon, single junction gallium arsenide or dual junction gallium arsenide. The maximum power capability at end of life (15 years, summer solstice) ranges from 4 to 7.7 kilowatts for the three types of solar cells. The HS 601HP battery is the power source for the spacecraft during eclipse and peak sunlight power periods. The battery is comprised of four individual battery packs connected in series to produce a single battery. Each battery pack can accommodate a maximum of eight battery cells with a capacity of 350 ampere-hours. The battery pack also provides for mounting of all electronics utilized by the battery, such as cell bypassing. The power electronics for the HS 601HP spacecraft provide for a tightly regulated power bus whether in sunlight or eclipse (battery discharge) operation. The bus voltage during sunlight is maintained by two bus voltage limiters (BVL), located on the yoke of each solar wing. The BVL maintains the regulated power bus at 52.9 volts by shunting excess solar wing power when not required by the spacecraft. The bus voltage during eclipse is maintained by two battery power controllers (BPC) located on the spacecraft bus shelf. The BPC maintains the regulated power bus at 52.2 volts during battery discharge and also provides for battery charging when excess solar array power is available. The power from the solar array or battery is distributed to the spacecraft by bus and payload power distribution units (PDU). The HS 601HP spacecraft product line now has three spacecraft in orbit. The first was launched in early November of 1997 with the second and third launched in late November and early December of 1997, respectively. The power systems are performing as designed and correlate well with the predicted performance calculations. Several more HS 601HP are scheduled to launch during 1998.« less

Venus Gravity: 180th Degree and Order Model

NASA Technical Reports Server (NTRS)

Konopliv, A. S.; Banerdt, W. B.; Sjogren, W. L.

1998-01-01

The Megallan Doppler radiometric tracking data provides unprecedented precision for spacecraft based gravity measurements with the maximum resolution approaching spherical harmonic degree and order 180 in selected equatorial regions.

Fire safety applications for spacecraft

NASA Technical Reports Server (NTRS)

Friedman, Robert; Olson, Sandra L.

1989-01-01

Fire safety for spacecraft is reviewed by first describing current practices, many of which are adapted directly from aircraft. Then, current analyses and experimental knowledge in low-gravity combustion, with implications for fire safety are discussed. In orbiting spacecraft, the detection and suppression of flames are strongly affected by the large reduction in buoyant flows under low gravity. Generally, combustion intensity is reduced in low gravity. There are some notable exceptions, however, one example being the strong enhancement of flames by low-velocity ventilation flows in space. Finally, the future requirements in fire safety, particularly the needs of long-duration space stations in fire prevention, detection, extinguishment, and atmospheric control are examined. The goal of spacecraft fire-safety investigations is the establishment of trade-offs that promote maximum safety without hampering the useful human and scientific activities in space.

Motor/generator and electronic control considerations for energy storage flywheels

NASA Technical Reports Server (NTRS)

Nola, F. J.

1984-01-01

A spacecraft electric power supply system is described. Requirements of the system are to accelerate a momentum wheel to a fixed maximum speed when solar energy is available and to maintain a constant voltage on the spacecraft bus under varying loads when solar energy is not available. Candidate motor types, pulse width modulated current control systems, and efficiency considerations are discussed. In addition, the Lunar Roving Vehicle motors are described along with their respective efficiencies.

Optimization techniques applied to passive measures for in-orbit spacecraft survivability

NASA Technical Reports Server (NTRS)

Mog, Robert A.; Price, D. Marvin

1987-01-01

Optimization techniques applied to passive measures for in-orbit spacecraft survivability, is a six-month study, designed to evaluate the effectiveness of the geometric programming (GP) optimization technique in determining the optimal design of a meteoroid and space debris protection system for the Space Station Core Module configuration. Geometric Programming was found to be superior to other methods in that it provided maximum protection from impact problems at the lowest weight and cost.

Balancing reliability and cost to choose the best power subsystem

NASA Technical Reports Server (NTRS)

Suich, Ronald C.; Patterson, Richard L.

1991-01-01

A mathematical model is presented for computing total (spacecraft) subsystem cost including both the basic subsystem cost and the expected cost due to the failure of the subsystem. This model is then used to determine power subsystem cost as a function of reliability and redundancy. Minimum cost and maximum reliability and/or redundancy are not generally equivalent. Two example cases are presented. One is a small satellite, and the other is an interplanetary spacecraft.

Prospective Out-of-ecliptic White-light Imaging of Interplanetary Corotating Interaction Regions at Solar Maximum

NASA Astrophysics Data System (ADS)

Xiong, Ming; Davies, Jackie A.; Li, Bo; Yang, Liping; Liu, Ying D.; Xia, Lidong; Harrison, Richard A.; Keiji, Hayashi; Li, Huichao

2017-07-01

Interplanetary corotating interaction regions (CIRs) can be remotely imaged in white light (WL), as demonstrated by the Solar Mass Ejection Imager (SMEI) on board the Coriolis spacecraft and Heliospheric Imagers (HIs) on board the twin Solar TErrestrial RElations Observatory (STEREO) spacecraft. The interplanetary WL intensity, due to Thomson scattering of incident sunlight by free electrons, is jointly determined by the 3D distribution of electron number density and line-of-sight (LOS) weighting factors of the Thomson-scattering geometry. The 2D radiance patterns of CIRs in WL sky maps look very different from different 3D viewpoints. Because of the in-ecliptic locations of both the STEREO and Coriolis spacecraft, the longitudinal dimension of interplanetary CIRs has, up to now, always been integrated in WL imagery. To synthesize the WL radiance patterns of CIRs from an out-of-ecliptic (OOE) vantage point, we perform forward magnetohydrodynamic modeling of the 3D inner heliosphere during Carrington Rotation CR1967 at solar maximum. The mixing effects associated with viewing 3D CIRs are significantly minimized from an OOE viewpoint. Our forward modeling results demonstrate that OOE WL imaging from a latitude greater than 60° can (1) enable the garden-hose spiral morphology of CIRs to be readily resolved, (2) enable multiple coexisting CIRs to be differentiated, and (3) enable the continuous tracing of any interplanetary CIR back toward its coronal source. In particular, an OOE view in WL can reveal where nascent CIRs are formed in the extended corona and how these CIRs develop in interplanetary space. Therefore, a panoramic view from a suite of wide-field WL imagers in a solar polar orbit would be invaluable in unambiguously resolving the large-scale longitudinal structure of CIRs in the 3D inner heliosphere.

Solar maximum mission/ultraviolet spectrometer and polarimeter studies

NASA Technical Reports Server (NTRS)

Henze, William, Jr.

1993-01-01

This final report for NASA Contract No. NAS8-35921 describes various studies performed for the Ultraviolet Spectrometer and Polarimeter (UVSP) experiment, one of several instruments on the Solar Maximum Mission (SMM) satellite which was launched on 14 February 1980. The UVSP consisted primarily of a Gregorian telescope and an Ebert-Fastie spectrometer with a polarimeter that could be inserted into the light path. The spacecraft and most of the instruments, including the UVSP, operated successfully until 23 November 1980, when part of the SMM attitude control system (fine pointing control) failed. The UVSP was then unable to observe the Sun until 18 April 1984, when the SMM was visited by the space shuttle and the attitude control module was replaced by astronauts. The SMM mission ended when the spacecraft reentered the atmosphere of the Earth and was thereby destroyed on 2 December 1989. The topics covered in this report include the following: (1) ultraviolet stellar polarimetry (probably the first such attempted measurement); no polarization as detected and the upper limits, based on the sensitivity as determined by the observed count rate, are rather high; (2) an investigation into the possible position of the UVSP wavelength drive after it became stuck on 26 April 1985; (3) fast timing tests for sit-and-state observations involving one or two detectors; (4) development of computer subroutines to allow the calculation of the component of the SMM spacecraft orbital velocity along the line of sight to the Sun at any desired time during the 1984/1985 period when the UVSP wavelength drive was operating properly; (5) listing of published research papers; (6) description of the UVSP catalog of observations; and (7) description of UVSP calibration report and data users guide.

Carbon-Carbon Composite Radiator Development for the EO-1 Spacecraft

NASA Technical Reports Server (NTRS)

Vaughn, Wallace; Shinn, Elizabeth; Rawal, Suraj; Wright, Joe

2004-01-01

The Carbon-Carbon Space Radiator Partnership (CSRP), an informal partnership of Government and industrial personnel, was formed to promote the use of Carbon-carbon composites (C-C) as engineering materials for spacecraft thermal management applications . As a part of this effort the partnership has built a structural radiator for the Earth Orbiter - 1 (EO-1) spacecraft. This radiator, using C-C face-sheets with an aluminum honeycomb core, will demonstrate both the thermal and structural properties of C-C under actual service conditions as well as provide performance data from space flight. This paper will present results from the design of the radiator, the thermal/mechanical tests of the facesheet materials, and sub-component test results on the C-C/Al honeycomb sandwich material. The 29- by 28-inch radiator was designed to support two electronics boxes with a combined heat output of 60 watts maximum and a weight of 58 lbs. The analysis of the radiator design shows that the radiator constructed with 20-mil-thick facesheets of a P30-fiber-reinforced C-C from BFGoodrich is able to meet or exceed all the required thermal and mechanical requirements.

Applications Explorer Missions (AEM): Mission planners handbook

NASA Technical Reports Server (NTRS)

Smith, S. R. (Editor)

1974-01-01

The Applications Explorer Missions (AEM) Program is a planned series of space applications missions whose purpose is to perform various tasks that require a low cost, quick reaction, small spacecraft in a dedicated orbit. The Heat Capacity Mapping Mission (HCMM) is the first mission of this series. The spacecraft described in this document was conceived to support a variety of applications instruments and the HCMM instrument in particular. The maximum use of commonality has been achieved. That is, all of the subsystems employed are taken directly or modified from other programs such as IUE, IMP, RAE, and Nimbus. The result is a small versatile spacecraft. The purpose of this document, the AEM Mission Planners Handbook (AEM/MPH) is to describe the spacecraft and its capabilities in general and the HCMM in particular. This document will also serve as a guide for potential users as to the capabilities of the AEM spacecraft and its achievable orbits. It should enable each potential user to determine the suitability of the AEM concept to his mission.

Quasi-Sun-Pointing of Spacecraft Using Radiation Pressure

NASA Technical Reports Server (NTRS)

Spilker, Thomas

2003-01-01

A report proposes a method of utilizing solar-radiation pressure to keep the axis of rotation of a small spin-stabilized spacecraft pointed approximately (typically, within an angle of 10 deg to 20 deg) toward the Sun. Axisymmetry is not required. Simple tilted planar vanes would be attached to the outer surface of the body, so that the resulting spacecraft would vaguely resemble a rotary fan, windmill, or propeller. The vanes would be painted black for absorption of Solar radiation. A theoretical analysis based on principles of geometric optics and mechanics has shown that torques produced by Solar-radiation pressure would cause the axis of rotation to precess toward Sun-pointing. The required vane size would be a function of the angular momentum of the spacecraft and the maximum acceptable angular deviation from Sun-pointing. The analysis also shows that the torques produced by the vanes would slowly despin the spacecraft -- an effect that could be counteracted by adding specularly reflecting "spin-up" vanes.

Trace Contaminant Control for the International Space Station's Node 1- Analysis, Design, and Verification

NASA Technical Reports Server (NTRS)

Perry, J. L.

2017-01-01

Trace chemical contaminant generation inside crewed spacecraft cabins is a technical and medical problem that must be continuously evaluated. Although passive control through materials selection and active control by adsorption and catalytic oxidation devices is employed during normal operations of a spacecraft, contaminant buildup can still become a problem. Buildup is particularly troublesome during the stages between the final closure of a spacecraft during ground processing and the time that a crewmember enters for the first time during the mission. Typically, the elapsed time between preflight closure and first entry on orbit for spacecraft such as Spacelab modules was 30 days. During that time, the active contamination control systems are not activated and contaminants can potentially build up to levels which exceed the spacecraft maximum allowable concentrations (SMACs) specified by NASA toxicology experts. To prevent excessively high contamination levels at crew entry, the Spacelab active contamination control system was operated for 53 hours just before launch.

Evaluation of a Conductive Elastomer Seal for Spacecraft

NASA Technical Reports Server (NTRS)

Daniels, C. C.; Mather, J. L.; Oravec, H. A.; Dunlap, P. H., Jr.

2016-01-01

An electrically conductive elastomer was evaluated as a material candidate for a spacecraft seal. The elastomer used electrically conductive constituents as a means to reduce the resistance between mating interfaces of a sealed joint to meet spacecraft electrical bonding requirements. The compound's outgassing levels were compared against published NASA requirements. The compound was formed into a hollow O-ring seal and its compression set was measured. The O-ring seal was placed into an interface and the electrical resistance and leak rate were quantified. The amount of force required to fully compress the test article in the sealing interface and the force needed to separate the joint were also measured. The outgassing and resistance measurements were below the maximum allowable levels. The room temperature compression set and leak rates were fairly high when compared against other typical spacecraft seal materials, but were not excessive. The compression and adhesion forces were desirably low. Overall, the performance of the elastomer compound was sufficient to be considered for future spacecraft seal applications.

Solar wind structure suggested by bimodal correlations of solar wind speed and density between the spacecraft SOHO and Wind

NASA Astrophysics Data System (ADS)

Ogilvie, K. W.; Coplan, M. A.; Roberts, D. A.; Ipavich, F.

2007-08-01

We calculate the cross-spacecraft maximum lagged-cross-correlation coefficients for 2-hour intervals of solar wind speed and density measurements made by the plasma instruments on the Solar and Heliospheric Observatory (SOHO) and Wind spacecraft over the period from 1996, the minimum of solar cycle 23, through the end of 2005. During this period, SOHO was located at L1, about 200 R E upstream from the Earth, while Wind spent most of the time in the interplanetary medium at distances of more than 100 R E from the Earth. Yearly histograms of the maximum, time-lagged correlation coefficients for both the speed and density are bimodal in shape, suggesting the existence of two distinct solar wind regimes. The larger correlation coefficients we suggest are due to structured solar wind, including discontinuities and shocks, while the smaller are likely due to Alfvénic turbulence. While further work will be required to firmly establish the physical nature of the two populations, the results of the analysis are consistent with a solar wind that consists of turbulence from quiet regions of the Sun interspersed with highly filamentary structures largely convected from regions in the inner solar corona. The bimodal appearance of the distributions is less evident in the solar wind speed than in the density correlations, consistent with the observation that the filamentary structures are convected with nearly constant speed by the time they reach 1 AU. We also find that at solar minimum the fits for the density correlations have smaller high-correlation components than at solar maximum. We interpret this as due to the presence of more relatively uniform Alfvénic regions at solar minimum than at solar maximum.

Aeroheating Thermal Model Correlation for Mars Global Surveyor (MGS) Solar Array

NASA Technical Reports Server (NTRS)

Amundsen, Ruth M.; Dec, John A.; George, Benjamin E.

2003-01-01

The Mars Global Surveyor (MGS) Spacecraft made use of aerobraking to gradually reduce its orbit period from a highly elliptical insertion orbit to its final science orbit. Aerobraking produces a high heat load on the solar arrays, which have a large surface area exposed to the airflow and relatively low mass. To accurately model the complex behavior during aerobraking, the thermal analysis needed to be tightly coupled to the spatially varying, time dependent aerodynamic heating. Also, the thermal model itself needed to accurately capture the behavior of the solar array and its response to changing heat load conditions. The correlation of the thermal model to flight data allowed a validation of the modeling process, as well as information on what processes dominate the thermal behavior. Correlation in this case primarily involved detailing the thermal sensor nodes, using as-built mass to modify material property estimates, refining solar cell assembly properties, and adding detail to radiation and heat flux boundary conditions. This paper describes the methods used to develop finite element thermal models of the MGS solar array and the correlation of the thermal model to flight data from the spacecraft drag passes. Correlation was made to data from four flight thermal sensors over three of the early drag passes. Good correlation of the model was achieved, with a maximum difference between the predicted model maximum and the observed flight maximum temperature of less than 5%. Lessons learned in the correlation of this model assisted in validating a similar model and method used for the Mars Odyssey solar array aeroheating analysis, which were used during onorbit operations.

SMART-1: key technologies and autonomy implementations

NASA Astrophysics Data System (ADS)

Elfving, A.; Stagnaro, L.; Winton, A.

2003-01-01

SMART-1 is the first of the Small Missions for Advanced Research in Technology of the ESA Horizons 2000 Science Plan. The main mission objective of SMART-1 is to demonstrate innovative and key technologies for scientific deep-space missions. One of the key technologies is the solar electric propulsion used as primary propulsion. The electric propulsion will be using 1400 W to transfer the 350 kg spacecraft from an Ariane 5 standard GTO to an elliptic Moon polar orbit, 10000×300 km. The total mission time is 24 months including a maximum of 18 months transfer time. The spacecraft development entered the detailed design and implementation phase in October 1999, under the responsibility of the Swedish Space Cooperation as prime contractor, and the flight acceptance is targeted for the 3rd quarter of 2002. The committed total life cost budget is 84 million Euro. Apart from the in-orbit demonstration of electric propulsion as primary propulsion, SMART-1 is implementing many other enabling technologies for deep-space missions such as deep-space transponder with communication in X-band for uplink and in X- and Ka-band for downlink, highly integrated and radiation tolerant microprocessors, FPGAs and memories, high on-board autonomy driven by ground communication only once per 4 days, maximum available power to electric propulsion by centralised and software-controlled power management, minimum propellant consumption through innovative angular momentum management. In addition, the spacecraft avionics design is tailored to the low-cost philosophy by enabling flexible integration of commercial off the shelf (COTS) equipment. The scientific instruments, five in total, support characterisation of the electric propulsion thrust environment during the long transfer phase and detailed imaging and spectroscopy of the lunar surface in visible, infrared and X-ray during the Moon orbiting phase. Several of the instruments implement new enabling technologies, e.g. swept charge devices for the X-ray spectrometer and quarts gratings for the near-infrared reflectance spectrometer. The paper summarises the baseline mission design, the projected scientific and technology output, and describe the spacecraft bus design. The main part of the paper highlights and elaborate on a number of spacecraft bus technologies; the on-board autonomy, the avionics design, and the X/Ka-band transponder.

Space Debris Mitigation Efforts through the Disposition of the Service Module of the Unmanned Space Experiment Recovery System (USERS)

NASA Astrophysics Data System (ADS)

Ijichi, Koichi; Ushikoshi, Atsuo; Nakamura, Shuji; Kanai, Hiroshi

The Unmanned Space Experiment Recovery System (USERS) Project has been completed with full success, and the Service Module (SEM) of the USERS Spacecraft, which supported the recovery portion of the spacecraft which was left on the orbit, was properly disposed to the maximum extent as possible according to the IADC debris mitigation guideline and re-entered the atmosphere on June 15, 2007 (JST). USERS spacecraft disposition by possible means available at the mission completion showed good example of realizing debris mitigation purpose in spite of originally different design baseline, and obtained actual data and experiences to be reflected for future space programs.

Practical Applications of Cosmic Ray Science: Spacecraft, Aircraft, Ground Based Computation and Control Systems and Human Health and Safety

NASA Technical Reports Server (NTRS)

Atwell, William; Koontz, Steve; Normand, Eugene

2012-01-01

In this paper we review the discovery of cosmic ray effects on the performance and reliability of microelectronic systems as well as on human health and safety, as well as the development of the engineering and health science tools used to evaluate and mitigate cosmic ray effects in earth surface, atmospheric flight, and space flight environments. Three twentieth century technological developments, 1) high altitude commercial and military aircraft; 2) manned and unmanned spacecraft; and 3) increasingly complex and sensitive solid state micro-electronics systems, have driven an ongoing evolution of basic cosmic ray science into a set of practical engineering tools (e.g. ground based test methods as well as high energy particle transport and reaction codes) needed to design, test, and verify the safety and reliability of modern complex electronic systems as well as effects on human health and safety. The effects of primary cosmic ray particles, and secondary particle showers produced by nuclear reactions with spacecraft materials, can determine the design and verification processes (as well as the total dollar cost) for manned and unmanned spacecraft avionics systems. Similar considerations apply to commercial and military aircraft operating at high latitudes and altitudes near the atmospheric Pfotzer maximum. Even ground based computational and controls systems can be negatively affected by secondary particle showers at the Earth's surface, especially if the net target area of the sensitive electronic system components is large. Accumulation of both primary cosmic ray and secondary cosmic ray induced particle shower radiation dose is an important health and safety consideration for commercial or military air crews operating at high altitude/latitude and is also one of the most important factors presently limiting manned space flight operations beyond low-Earth orbit (LEO).

Development of a hardware-in-the-loop testbed to demonstrate multiple spacecraft operations in proximity

NASA Astrophysics Data System (ADS)

Eun, Youngho; Park, Sang-Young; Kim, Geuk-Nam

2018-06-01

This paper presents a new state-of-the-art ground-based hardware-in-the-loop test facility, which was developed to verify and demonstrate autonomous guidance, navigation, and control algorithms for space proximity operations and formation flying maneuvers. The test facility consists of two complete spaceflight simulators, an aluminum-based operational arena, and a set of infrared motion tracking cameras; thus, the testbed is capable of representing space activities under circumstances prevailing on the ground. The spaceflight simulators have a maximum of five-degree-of-freedom in a quasi-momentum-free environment, which is produced by a set of linear/hemispherical air-bearings and a horizontally leveled operational arena. The tracking system measures the real-time three-dimensional position and attitude to provide state variables to the agents. The design of the testbed is illustrated in detail for every element throughout the paper. The practical hardware characteristics of the active/passive measurement units and internal actuators are identified in detail from various perspectives. These experimental results support the successful development of the entire facility and enable us to implement and verify the spacecraft proximity operation strategy in the near future.

Modular, Reconfigurable, and Rapid Response Space Systems: The Remote Sensing Advanced Technology Microsatellite

NASA Technical Reports Server (NTRS)

Esper, Jaime; Andary, Jim; Oberright, John; So, Maria; Wegner, Peter; Hauser, Joe

2004-01-01

Modular, Reconfigurable, and Rapid-response (MR(sup 2)) space systems represent a paradigm shift in the way space assets of all sizes are designed, manufactured, integrated, tested, and flown. This paper will describe the MR(sup 2) paradigm in detail, and will include guidelines for its implementation. The Remote Sensing Advanced Technology microsatellite (RSAT) is a proposed flight system test-bed used for developing and implementing principles and best practices for MR(sup 2) spacecraft, and their supporting infrastructure. The initial goal of this test-bed application is to produce a lightweight (approx. 100 kg), production-minded, cost-effective, and scalable remote sensing micro-satellite capable of high performance and broad applicability. Such applications range from future distributed space systems, to sensor-webs, and rapid-response satellite systems. Architectures will be explored that strike a balance between modularity and integration while preserving the MR(sup 2) paradigm. Modularity versus integration has always been a point of contention when approaching a design: whereas one-of-a-kind missions may require close integration resulting in performance optimization, multiple and flexible application spacecraft benefit &om modularity, resulting in maximum flexibility. The process of building spacecraft rapidly (< 7 days), requires a concerted and methodical look at system integration and test processes and pitfalls. Although the concept of modularity is not new and was first developed in the 1970s by NASA's Goddard Space Flight Center (Multi-Mission Modular Spacecraft), it was never modernized and was eventually abandoned. Such concepts as the Rapid Spacecraft Development Office (RSDO) became the preferred method for acquiring satellites. Notwithstanding, over the past 30 years technology has advanced considerably, and the time is ripe to reconsider modularity in its own right, as enabler of R(sup 2), and as a key element of transformational systems. The MR2 architecture provides a competitive advantage over the old modular approach in its rapid response to market needs that are difficult to predict both from the perspectives of evolving technology, as well as mission and application requirements.

NanoSail-D: A Solar Sail Demonstration Mission

NASA Technical Reports Server (NTRS)

Johnson, Les; Whorton, Mark; Heaton, Andy; Pinson, robin; Laue, Greg; Adams, Charles

2009-01-01

During the past decade, within the United States, NASA Marshall Space Flight Center (MSFC) was heavily engaged in the development of revolutionary new technologies for in-space propulsion. One of the major in-space propulsion technologies developed was a solar sail propulsion system. Solar sail propulsion uses the solar radiation pressure exerted by the momentum transfer of reflected photons to generate a net force on a spacecraft. To date, solar sail propulsion systems have been designed for large spacecraft in the tens to hundreds of kilograms mass range. Recently, however, MSFC has been investigating the application of solar sails for small satellite propulsion. Likewise, NASA Ames Research Center (ARC) has been developing small spacecraft missions that have a need for amass-efficient means of satisfying deorbit requirements. Hence, a synergistic collaboration was established between these two NASA field Centers with the objective of conducting a flight demonstration of solar sail technologies for small satellites. The NanoSail-D mission flew onboard the ill-fated Falcon Rocket launched August 2, 2008, and, due to the failure of that rocket, never achieved orbit. The NanoSail-D flight spare is ready for flight and a suitable launch arrangement is being actively pursued. Both the original sailcraft and the flight spare are hereafter referred to as NanoSail-D. The sailcraft consists of a sail subsystem stowed in a three-element CubeSat. Shortly after deployment of the NanoSail-D, the solar sail will deploy and mission operations will commence. This demonstration flight has two primary technical objectives: (1) to successfully stow and deploy the sail and (2) to demonstrate deorbit functionality. Given a near-term opportunity for launch on Falcon, the project was given the challenge of delivering the flight hardware in 6 mo, which required a significant constraint on flight system functionality. As a consequence, passive attitude stabilization of the spacecraft will be achieved using permanent magnets to detumble and orient the body with the magnetic field lines and then rely on atmospheric drag to passively stabilize the sailcraft in an essentially maximum drag attitude. This paper will present an introduction to solar sail propulsion systems and an overview of the NanoSail-D spacecraft.

Standardization and program effect analysis (Study 2.4). Volume 2: Equipment commonality analysis. [cost savings of using flight-proven components in designing spacecraft

NASA Technical Reports Server (NTRS)

Shiokari, T.

1975-01-01

The feasibility and cost savings of using flight-proven components in designing spacecraft were investigated. The components analyzed were (1) large space telescope, (2) stratospheric aerosol and gas equipment, (3) mapping mission, (4) solar maximum mission, and (5) Tiros-N. It is concluded that flight-proven hardware can be used with not-too-extensive modification, and significant savings can be realized. The cost savings for each component are presented.

Flight performance of the Pioneer Venus Orbiter solar array

NASA Technical Reports Server (NTRS)

Goldhammer, L. J.; Powe, J. S.; Smith, Marcie

1987-01-01

The Pioneer Venus Orbiter (PVO) solar panel power output capability has degraded much more severely than has the power output capability of solar panels that have operated in earth-orbiting spacecraft for comparable periods of time. The incidence of solar proton events recorded by the spacecraft's scientific instruments accounts for this phenomenon only in part. It cannot explain two specific forms of anomalous behavior observed: 1) a variation of output per spin with roll angle, and 2) a gradual degradation of the maximum output. Analysis indicates that the most probable cause of the first anomaly is that the solar cells underneath the spacecraft's magnetometer boom have been damaged by a reverse biasing of the cells that occurs during pulsed shadowing of the cells by the boom as the spacecraft rotates. The second anomaly might be caused by the effects on the solar array of substances from the upper atmosphere of Venus.

SNAP 19 Viking Program. Bimonthly technical progress report, October 1979-November 1979

DOE Office of Scientific and Technical Information (OSTI.GOV)

Not Available

1979-12-01

Monitoring and evaluation of Viking 1 Lander power system data continued. The RTG series power range as measured at the PCDA was 65 to 68 watts at fin root temperatures between 280/sup 0/F and 310/sup 0/F. The Mars landed performance history of Viking 1 include both the minimum and maximum data for each of the SOL days. Monitoring and evaluation of Viking 2 Lander power system data continued. The RTG series power range as measured at the PCDA was 71 to 72 watts at fin root temperatures between 230/sup 0/F and 260/sup 0/F. The Mars landed performance history of Vikingmore » 2 include both the minimum and maximum data for each of the SOL days. The performance of both power systems continues to be very satisfactory. Power system performance data for Pioneer 10 and Pioneer 11 spacecraft were monitored through the reporting period. The estimated RTG system net power was 116 watts for Pioneer 10 and 118 watts for Pioneer Saturn. The September 1 encounter with Saturn appears to have had no deleterious effect on the RTG's of the spacecraft power system. The telemetry signals from both spacecrafts remain satisfactory.« less

Drag De-Orbit Device: A New Standard Re-Entry Actuator for CubeSats

NASA Technical Reports Server (NTRS)

Guglielmo, David; Omar, Sanny; Bevilacqua, Riccardo

2017-01-01

With the advent of CubeSats, research in Low Earth Orbit (LEO) becomes possible for universities and small research groups. Only a handful of launch sites can be used, due to geographical and political restrictions. As a result, common orbits in LEO are becoming crowded due to the additional launches made possible by low-cost access to space. CubeSat design principles require a maximum of a 25-year orbital lifetime in an effort to reduce the total number of spacecraft in orbit at any time. Additionally, since debris may survive re-entry, it is ideal to de-orbit spacecraft over unpopulated areas to prevent casualties. The Drag Deorbit Device (D3) is a self-contained targeted re-entry subsystem intended for CubeSats. By varying the cross-wind area, the atmospheric drag can be varied in such a way as to produce desired maneuvers. The D3 is intended to be used to remove spacecraft from orbit to reach a desired target interface point. Additionally, attitude stabilization is performed by the D3 prior to deployment and can replace a traditional ADACS on many missions.This paper presents the hardware used in the D3 and operation details. Four stepper-driven, repeatedly retractable booms are used to modify the cross-wind area of the D3 and attached spacecraft. Five magnetorquers (solenoids) over three axes are used to damp rotational velocity. This system is expected to be used to improve mission flexibility and allow additional launches by reducing the orbital lifetime of spacecraft.The D3 can be used to effect a re-entry to any target interface point, with the orbital inclination limiting the maximum latitude. In the chance that the main spacecraft fails, a timer will automatically deploy the booms fully, ensuring the spacecraft will at the minimum reenter the atmosphere in the minimum possible time, although not necessarily at the desired target interface point. Although this does not reduce the risk of casualties, the 25-year lifetime limit is still respected, allowing a reduction of the risk associated with a hardware failure.

Research to Support the Determination of Spacecraft Maximum Acceptable Concentrations of Potential Atmospheric Contaminants

NASA Technical Reports Server (NTRS)

Orr, John L.

1997-01-01

In many ways, the typical approach to the handling of bibliographic material for generating review articles and similar manuscripts has changed little since the use of xerographic reproduction has become widespread. The basic approach is to collect reprints of the relevant material and place it in folders or stacks based on its dominant content. As the amount of information available increases with the passage of time, the viability of this mechanical approach to bibliographic management decreases. The personal computer revolution has changed the way we deal with many familiar tasks. For example, word processing on personal computers has supplanted the typewriter for many applications. Similarly, spreadsheets have not only replaced many routine uses of calculators but have also made possible new applications because the cost of calculation is extremely low. Objective The objective of this research was to use personal computer bibliographic software technology to support the determination of spacecraft maximum acceptable concentration (SMAC) values. Specific Aims The specific aims were to produce draft SMAC documents for hydrogen sulfide and tetrachloroethylene taking maximum advantage of the bibliographic software.

Space Weather Impacts on Spacecraft Design and Operations in Auroral Charging Environments

NASA Technical Reports Server (NTRS)

Minow, Joseph I.; Parker, Linda N.

2012-01-01

Spacecraft in low altitude, high inclination (including sun-synchronous) orbits are widely used for remote sensing of the Earth s land surface and oceans, monitoring weather and climate, communications, scientific studies of the upper atmosphere and ionosphere, and a variety of other scientific, commercial, and military applications. These systems are episodically exposed to environments characterized by a high flux of energetic (approx.1 to 10 s kilovolt) electrons in regions of very low background plasma density which is similar in some ways to the space weather conditions in geostationary orbit responsible for spacecraft charging to kilovolt levels. While it is well established that charging conditions in geostationary orbit are responsible for many anomalies and even spacecraft failures, to date there have been relatively few such reports due to charging in auroral environments. This presentation first reviews the physics of the space environment and its interactions with spacecraft materials that control auroral charging rates and the anticipated maximum potentials that should be observed on spacecraft surfaces during disturbed space weather conditions. We then describe how the theoretical values compare to the observational history of extreme charging in auroral environments and discuss how space weather impacts both spacecraft design and operations for vehicles on orbital trajectories that traverse auroral charging environments.

Spacecraft Mission Design for the Mitigation of the 2017 PDC Hypothetical Asteroid Threat

NASA Technical Reports Server (NTRS)

Barbee, Brent W.; Sarli, Bruno V.; Lyzhoft, Josh; Chodas, Paul W.; Englander, Jacob A.

2017-01-01

This paper presents a detailed mission design analysis results for the 2017 Planetary Defense Conference (PDC) Hypothetical Asteroid Impact Scenario, documented at https:cneos.jpl.nasa.govpdcspdc17. The mission design includes campaigns for both reconnaissance (flyby or rendezvous) of the asteroid (to characterize it and the nature of the threat it poses to Earth) and mitigation of the asteroid, via kinetic impactor deflection, nuclear explosive device (NED) deflection, or NED disruption. Relevant scenario parameters are varied to assess the sensitivity of the design outcome, such as asteroid bulk density, asteroid diameter, momentum enhancement factor, spacecraft launch vehicle, and mitigation system type. Different trajectory types are evaluated in the mission design process from purely ballistic to those involving optimal midcourse maneuvers, planetary gravity assists, and/or low-thrust solar electric propulsion. The trajectory optimization is targeted around peak deflection points that were found through a novel linear numerical technique method. The optimization process includes constrain parameters, such as Earth departure date, launch declination, spacecraft, asteroid relative velocity and solar phase angle, spacecraft dry mass, minimum/maximum spacecraft distances from Sun and Earth, and Earth-spacecraft communications line of sight. Results show that one of the best options for the 2017 PDC deflection is solar electric propelled rendezvous mission with a single spacecraft using NED for the deflection.

Spacecraft Maximum Allowable Concentrations for Airborne Contaminants

NASA Technical Reports Server (NTRS)

James, John T.

2008-01-01

The enclosed table lists official spacecraft maximum allowable concentrations (SMACs), which are guideline values set by the NASA/JSC Toxicology Group in cooperation with the National Research Council Committee on Toxicology (NRCCOT). These values should not be used for situations other than human space flight without careful consideration of the criteria used to set each value. The SMACs take into account a number of unique factors such as the effect of space-flight stress on human physiology, the uniform good health of the astronauts, and the absence of pregnant or very young individuals. Documentation of the values is given in a 5 volume series of books entitled "Spacecraft Maximum Allowable Concentrations for Selected Airborne Contaminants" published by the National Academy Press, Washington, D.C. These books can be viewed electronically at http://books.nap.edu/openbook.php?record_id=9786&page=3. Short-term (1 and 24 hour) SMACs are set to manage accidental releases aboard a spacecraft and permit risk of minor, reversible effects such as mild mucosal irritation. In contrast, the long-term SMACs are set to fully protect healthy crewmembers from adverse effects resulting from continuous exposure to specific air pollutants for up to 1000 days. Crewmembers with allergies or unusual sensitivity to trace pollutants may not be afforded complete protection, even when long-term SMACs are not exceeded. Crewmember exposures involve a mixture of contaminants, each at a specific concentration (C(sub n)). These contaminants could interact to elicit symptoms of toxicity even though individual contaminants do not exceed their respective SMACs. The air quality is considered acceptable when the toxicity index (T(sub grp)) for each toxicological group of compounds is less than 1, where T(sub grp), is calculated as follows: T(sub grp) = C(sub 1)/SMAC(sub 1) + C(sub 2/SMAC(sub 2) + ...+C(sub n)/SMAC(sub n).

Obtaining the Greatest Scientific Benefit from Observational Platforms by Consideration of the Relative Benefit of Observations

NASA Technical Reports Server (NTRS)

Chelberg, David; Drews, Frank; Fleeman, David; Welch, Lonnie; Marquart, Jane; Pfarr, Barbara

2003-01-01

One of the current trends in spacecraft software design is to increase the autonomy of onboard flight and science software. This is especially true when real-time observations may affect the observation schedule of a mission. For many science missions, such as those conducted by the Swift Burst Alert Telescope, the ability of the spacecraft to autonomously respond in real-time to unpredicted science events is crucial for mission success. We apply utility theory within resource management middleware to optimize the real-time performance of application software and achieve maximum system level benefit. We then explore how this methodology can be extended to manage both software and observational resources onboard a spacecraft to achieve the best possible observations.

Validation of Modified Wine-Rack Thermal Design for Nickel-Hydrogen Batteries in Landsat-7 Spacecraft Thermal Vacuum Test and in Flight

NASA Technical Reports Server (NTRS)

Choi, Michael K.

1999-01-01

A heritage wine-rack thermal/mechanical design for the nickel-hydrogen batteries was the baseline at the Landsat-7 Preliminary Design Review. An integrated thermal and power analysis of the batteries performed by the author in 1994 revealed that the maximum cell-to-cell gradient was 6.6 C. The author proposed modifying the heritage wine-rack design by enhancing heat conduction from cells to cells, and from cells to battery frame. At the 1995 Intersociety Energy Conversion Engineering Conference (IECEC), the author presented a paper on methods of modifying the wine-rack design. It showed that the modified wine-rack option, which uses a metallic filler, could reduce the maximum cell-to-cell temperature gradient to 1.30 C, and could also reduce the maximum cell temperature by as much as 80 C. That design concept was adopted by the Landsat7 Project Office, and a design change was made at the Critical Design Review. Results of the spacecraft thermal vacuum and thermal balance tests, and temperature data in flight show that the temperatures of the battery cells are very uniform. The maximum cell-to-cell gradient is 1.50 C. They validate the modified wine-rack thermal design.

Differential Drag Demonstration: A Post-Mission Experiment with the EO-1 Spacecraft

NASA Technical Reports Server (NTRS)

Hull, Scott; Shelton, Amanda; Richardson, David

2017-01-01

Differential drag is a technique for altering the semi-major axis, velocity, and along-track position of a spacecraft in low Earth orbit. It involves varying the spacecrafts cross-sectional area relative to its velocity direction by temporarily changing attitude and solar array angles, thus varying the amount of atmospheric drag on the spacecraft. The technique has recently been proposed and used by at least three satellite systems for initial separation of constellation spacecraft after launch, stationkeeping during the mission, and potentially for conjunction avoidance. Similarly, differential drag has been proposed as a control strategy for rendezvous, removing the need for active propulsion. In theory, some operational missions that lack propulsion capability could use this approach for conjunction avoidance, though options are typically constrained for spacecraft that are already in orbit. Shortly before the spacecraft was decommissioned, an experiment was performed using NASAs EO-1 spacecraft in order to demonstrate differential drag on an operational spacecraft in orbit, and discover some of the effects differential drag might manifest. EO-1 was not designed to maintain off-nominal orientations for long periods, and as a result the team experienced unanticipated challenges during the experiment. This paper will discuss operations limitations identified before the experiment, as well as those discovered during the experiment. The effective displacement that resulted from increasing the drag area for 39 hours will be compared to predictions as well as the expected position if the spacecraft maintained nominal operations. A hypothetical scenario will also be examined, studying the relative risks of maintaining an operational spacecraft bus in order to maintain the near-maximum drag area orientation and hasten reentry.

Differential Drag Demonstration: A Post-Mission Experiment with the EO-1 Spacecraft

NASA Technical Reports Server (NTRS)

Hull, Scott; Shelton, Amanda; Richardson, David

2017-01-01

Differential drag is a technique for altering the semimajor axis, velocity, and along-track position of a spacecraft in low Earth orbit. It involves varying the spacecraft's cross-sectional area relative to its velocity direction by temporarily changing attitude and solar array angles, thus varying the amount of atmospheric drag on the spacecraft. The technique has recently been proposed and used by at least three satellite systems for initial separation of constellation spacecraft after launch, stationkeeping during the mission, and potentially for conjunction avoidance. Similarly, differential drag has been proposed as a control strategy for rendezvous, removing the need for active propulsion. In theory, some operational missions that lack propulsion capability could use this approach for conjunction avoidance, though options are typically constrained for spacecraft that are already in orbit. Shortly before the spacecraft was decommissioned, an experiment was performed using NASA's EO-1 spacecraft in order to demonstrate differential drag on an operational spacecraft in orbit, and discover some of the effects differential drag might manifest. EO-1 was not designed to maintain off-nominal orientations for long periods, and as a result the team experienced unanticipated challenges during the experiment. This paper will discuss operations limitations identified before the experiment, as well as those discovered during the experiment. The effective displacement that resulted from increasing the drag area for 39 hours will be compared to predictions as well as the expected position if the spacecraft maintained nominal operations. A hypothetical scenario will also be examined, studying the relative risks of maintaining an operational spacecraft bus in order to maintain the near-maximum drag area orientation and hasten reentry.

Preliminary studies on the planetary entry to Jupiter by aerocapture technique

NASA Astrophysics Data System (ADS)

Aso, Shigeru; Yasaka, Tetsuo; Hirayama, Hiroshi; Poetro, Ridanto Eko; Hatta, Shinji

2006-10-01

Preliminary studies on the planetary entry to Jupiter by aerocapture technique are studied in order to complete technological challenges to deliver scientific probe with low cost and smaller mass of the spacecraft to Jupiter. Jupiter aerocapture corridor determination based on maximum deceleration limit of 5g (lower corridor) and aerocapture capability (upper corridor) at Jupiter are carefully considered and calculated. The results show about 1700 m/s of saving velocity due to aerocapture could be possible in some cases for the spacecraft to be captured by Jovian gravitational field. However, the results also show that Jovian aerocapture is not available in some cases. Hence, careful selection is needed to realize Jovian aerocapture. Also the numerical simulation of aerodynamic heating to the spacecraft has been conducted. DSMC method is used for the simulation of flow fields around the spacecraft. The transient changes of drag due to Jovian atmosphere and total heat loads to the spacecraft are obtained. The results show that the estimated heat loads could be within allowable amount heat load when some ablation heat shield technique is applied.

Preliminary studies on the planetary entry to Jupiter by aerocapture technique

NASA Astrophysics Data System (ADS)

Aso, Shigeru; Yasaka, Tetsuo; Hirayama, Hiroshi; Eko Poetro, Ridanto; Hatta, Shinji

2003-11-01

Preliminary studies on the planetary entry to Jupiter by aerocapture technique are studied in order to complete technological challenges to deliver scientific probe with low cost and smaller mass of the spacecraft to Jupiter. Jupiter aerocapture corridor determination based on maximum deceleration limit of 5g (lower corridor) and aerocapture capability (upper corridor) at Jupiter are carefully considered and calculated. The results show about 1700 m/s of saving velocity due to aerocapture could be possible in some cases for the spacecraft to be captured by Jovian gravitational field. However, the results also show that Jovian aerocapture is not available in some cases. Hence, careful selection is needed to realise Jovian aerocapture. Also the numerical simulation of aerodynamic heating to the spacecraft has been conducted. DSMC method is used for the simulation of flow fields around the spacecraft. The transient changes of drag due to Jovian atmosphere and total heat loads to the spacecraft are obtained. The results show the estimated heat loads could be within allowable amount heat load when some ablation heat shield technique is applied.

Evaluation of aperture cover tank vent nozzles for the IRAS spacecraft

NASA Technical Reports Server (NTRS)

Richter, R.

1983-01-01

The influence of coefficients for the three axes of the Infrared Astronomical Satellite (IRAS) were established to determine the maximum allowable thrust difference between the two vent nozzles of the aperture cover tank low thrust vent system and their maximum misalignment. Test data generated by flow and torque measurements permitted the selection of two nozzles whose thrust differential was within the limit of the attitude control capability. Based on thrust stand data, a thrust vector misalignment was indicated that was slightly higher than permissible for the worst case, i.e., considerable degradation of the torque capacity of the attitude control system combined with venting of helium at its upper limit. The probability of destabilizing the IRAS spacecraft by activating the venting system appeared to be very low. The selection and mounting of the nozzles have satisfied all the requirements for the safe venting of helium.

SCL: An off-the-shelf system for spacecraft control

NASA Astrophysics Data System (ADS)

Buckley, Brian; Vangaasbeck, James

1994-11-01

In this age of shrinking military, civil, and commercial space budgets, an off-the-shelf solution is needed to provide a multimission approach to spacecraft control. A standard operational interface which can be applied to multiple spacecraft allows a common approach to ground and space operations. A trend for many space programs has been to reduce operational staff by applying autonomy to the spacecraft and to the ground stations. The Spacecraft Command Language (SCL) system developed by Interface and Control Systems, Inc. (ICS) provides an off-the-shelf solution for spacecraft operations. The SCL system is designed to provide a hyper-scripting interface which remains standard from program to program. The spacecraft and ground station hardware specifics are isolated to provide the maximum amount of portability from system to system. Uplink and downlink interfaces are also isolated to allow the system to perform independent of the communications protocols chosen. The SCL system can be used for both the ground stations and the spacecraft, or as a value added package for existing ground station environments. The SCL system provides an expanded stored commanding capability as well as a rule-based expert system on-board. The expert system allows reactive control on-board the spacecraft for functions such as electrical power systems (EPS), thermal control, etc. which have traditionally been performed on the ground. The SCL rule and scripting capability share a common syntax allowing control of scripts from rules and rules from scripts. Rather than telemeter over sampled data to the ground, the SCL system maintains a database on-board which is available for interrogation by the scripts and rules. The SCL knowledge base is constructed on the ground and uploaded to the spacecraft. The SCL system follows an open-systems approach allowing other tasks to communicate with SCL on the ground and in space. The SCL system was used on the Clementine program (launched January 25, 1994) and is required to have bidirectional communications with the guidance, navigation, and control (GNC) algorithms which were written as another task. Sequencing of the spacecraft maneuvers are handled by SCL, but the low-level thruster pulse commands are handled by the GNC software. Attitude information is reported back as telemetry, allowing the SCL expert system to inference on the changing data. The Clementine SCL flight software was largely reused from another Naval Center for Space Technology (NCST) satellite program.

SCL: An off-the-shelf system for spacecraft control

NASA Technical Reports Server (NTRS)

Buckley, Brian; Vangaasbeck, James

1994-01-01

In this age of shrinking military, civil, and commercial space budgets, an off-the-shelf solution is needed to provide a multimission approach to spacecraft control. A standard operational interface which can be applied to multiple spacecraft allows a common approach to ground and space operations. A trend for many space programs has been to reduce operational staff by applying autonomy to the spacecraft and to the ground stations. The Spacecraft Command Language (SCL) system developed by Interface and Control Systems, Inc. (ICS) provides an off-the-shelf solution for spacecraft operations. The SCL system is designed to provide a hyper-scripting interface which remains standard from program to program. The spacecraft and ground station hardware specifics are isolated to provide the maximum amount of portability from system to system. Uplink and downlink interfaces are also isolated to allow the system to perform independent of the communications protocols chosen. The SCL system can be used for both the ground stations and the spacecraft, or as a value added package for existing ground station environments. The SCL system provides an expanded stored commanding capability as well as a rule-based expert system on-board. The expert system allows reactive control on-board the spacecraft for functions such as electrical power systems (EPS), thermal control, etc. which have traditionally been performed on the ground. The SCL rule and scripting capability share a common syntax allowing control of scripts from rules and rules from scripts. Rather than telemeter over sampled data to the ground, the SCL system maintains a database on-board which is available for interrogation by the scripts and rules. The SCL knowledge base is constructed on the ground and uploaded to the spacecraft. The SCL system follows an open-systems approach allowing other tasks to communicate with SCL on the ground and in space. The SCL system was used on the Clementine program (launched January 25, 1994) and is required to have bidirectional communications with the guidance, navigation, and control (GNC) algorithms which were written as another task. Sequencing of the spacecraft maneuvers are handled by SCL, but the low-level thruster pulse commands are handled by the GNC software. Attitude information is reported back as telemetry, allowing the SCL expert system to inference on the changing data. The Clementine SCL flight software was largely reused from another Naval Center for Space Technology (NCST) satellite program. This paper details the SCL architecture and how an off-the-shelf solution makes sense for multimission spacecraft programs. The Clementine mission will be used as a case study in the application of the SCL to a 'fast track' program. The benefits of such a system in a 'better, cheaper, faster' climate will be discussed.

Development of termination and utilization concepts for flat conductor cables. Volume 1: Development of low profile flat conductor cable connecting device and permanent splice

NASA Technical Reports Server (NTRS)

1972-01-01

The development of low-profile flat conductor cable (FCC) connecting device and FCC permanent splice methods are discussed. The design goal for the low-profile connecting device was to mate and unmate FCC harness to a typical spacecraft component with a maximum height of 3/8 in. The results indicate that the design, fabrication, and processing of the low-profile connecting device are feasible and practical. Some redesign will be required to achieve the goal of 3/8 in. Also, failures were experienced subsequent to salt spray and humidity exposure. Five different FCC permanent splice methods were considered. Subsequent to evaluation of these five methods, two design concepts were chosen for development tests.

Low thrust spacecraft transfers optimization method with the stepwise control structure in the Earth-Moon system in terms of the L1-L2 transfer

NASA Astrophysics Data System (ADS)

Fain, M. K.; Starinova, O. L.

2016-04-01

The paper outlines the method for determination of the locally optimal stepwise control structure in the problem of the low thrust spacecraft transfer optimization in the Earth-Moon system, including the L1-L2 transfer. The total flight time as an optimization criterion is considered. The optimal control programs were obtained by using the Pontryagin's maximum principle. As a result of optimization, optimal control programs, corresponding trajectories, and minimal total flight times were determined.

Spacecraft shielding for a Mars mission

NASA Astrophysics Data System (ADS)

O'Brien, K.

Calculations of the effective radiation dose due to cosmic rays in the interplanetary medium between Earth and Mars show that, as in the atmosphere above the Pfotzer Maximum, the dose rate increases with increasing wall thickness. An unshielded space crew member would receive almost 70 rem (0.70 Sv) a year. The effect of a typically proposed composite space-craft hull of aluminum and polyethylene would increase the dose rate by a few percent. However, 100 g/cm2 of almost any light material would more than double the cosmic radiation exposure of the crew.

Max '91: Flare research at the next solar maximum

NASA Technical Reports Server (NTRS)

Dennis, Brian; Canfield, Richard; Bruner, Marilyn; Emslie, Gordon; Hildner, Ernest; Hudson, Hugh; Hurford, Gordon; Lin, Robert; Novick, Robert; Tarbell, Ted

1988-01-01

To address the central scientific questions surrounding solar flares, coordinated observations of electromagnetic radiation and energetic particles must be made from spacecraft, balloons, rockets, and ground-based observatories. A program to enhance capabilities in these areas in preparation for the next solar maximum in 1991 is recommended. The major scientific issues are described, and required observations and coordination of observations and analyses are detailed. A program plan and conceptual budgets are provided.

Maximum Langmuir Fields in Planetary Foreshocks Determined from the Electrostatic Decay Threshold

NASA Technical Reports Server (NTRS)

Robinson, P. A.; Cairns, Iver H.

1995-01-01

Maximum electric fields of Langmuir waves at planetary foreshocks are estimated from the threshold for electrostatic decay, assuming it saturates beam driven growth, and incorporating heliospheric variation of plasma density and temperature. Comparisons with spacecraft observations yields good quantitative agreement. Observations in type 3 radio sources are also in accord with this interpretation. A single mechanism can thus account for the highest fields of beam driven waves in both contexts.

Lightweight Battery Charge Regulator Used to Track Solar Array Peak Power

NASA Technical Reports Server (NTRS)

Soeder, James F.; Button, Robert M.

1999-01-01

A battery charge regulator based on the series-connected boost regulator (SCBR) technology has been developed for high-voltage spacecraft applications. The SCBR regulates the solar array power during insolation to prevent battery overcharge or undercharge conditions. It can also be used to provide regulated battery output voltage to spacecraft loads if necessary. This technology uses industry-standard dc-dc converters and a unique interconnection to provide size, weight, efficiency, fault tolerance, and modularity benefits over existing systems. The high-voltage SCBR shown in the photograph has demonstrated power densities of over 1000 watts per kilogram (W/kg). Using four 150-W dc-dc converter modules, it can process 2500 W of power at 120 Vdc with a minimum input voltage of 90 Vdc. Efficiency of the SCBR was 94 to 98 percent over the entire operational range. Internally, the unit is made of two separate SCBR s, each with its own analog control circuitry, to demonstrate the modularity of the technology. The analog controllers regulate the output current and incorporate the output voltage limit with active current sharing between the two units. They also include voltage and current telemetry, on/off control, and baseplate temperature sensors. For peak power tracking, the SCBR was connected to a LabView-based data acquisition system for telemetry and control. A digital control algorithm for tracking the peak power point of a solar array was developed using the principle of matching the source impedance with the load impedance for maximum energy transfer. The algorithm was successfully demonstrated in a simulated spacecraft electrical system at the Boeing PhantomWorks High Voltage Test Facility in Seattle, Washington. The system consists of a 42-string, high-voltage solar array simulator, a 77-cell, 80-ampere-hour (A-hr) nickel-hydrogen battery, and a constant power-load module. The SCBR and the LabView control algorithm successfully tracked the solar array peak power point through various load transients, including sunlight discharge transients when the total load exceeded the maximum solar array output power.

Goddard Space Flight Center solar array missions, requirements and directions

NASA Technical Reports Server (NTRS)

Gaddy, Edward; Day, John

1994-01-01

The Goddard Space Flight Center (GSFC) develops and operates a wide variety of spacecraft for conducting NASA's communications, space science, and earth science missions. Some are 'in house' spacecraft for which the GSFC builds the spacecraft and performs all solar array design, analysis, integration, and test. Others are 'out of house' spacecraft for which an aerospace contractor builds the spacecraft and develops the solar array under direction from GSFC. The experience of developing flight solar arrays for numerous GSFC 'in house' and 'out of house' spacecraft has resulted in an understanding of solar array requirements for many different applications. This presentation will review those solar array requirements that are common to most GSFC spacecraft. Solar array technologies will be discussed that are currently under development and that could be useful to future GSFC spacecraft.

Environmental impact statement for the Kennedy Space Center, 1978 - 1979 revision

NASA Technical Reports Server (NTRS)

1979-01-01

The ongoing operation of KSC for expendable launch vehicles and automated spacecraft, continued development of facility capabilities, and the approved follow-on operations of the Space Transportation System and associated payloads are described. Emphasis is placed on the expendable launch vehicle and space shuttle traffic projected as of January, 1979. The maximum potential effect on the environment is addressed. Site specific environmental effects are summarized. It is indicated that all potential impacts will be localized, of short duration, controllable, and of minimum severity. The impact on land use, air and water quality, weather, and noise effects is covered.

Spacecraft Mission Design for the Mitigation of the 2017 PDC Hypothetical Asteroid Threat

NASA Technical Reports Server (NTRS)

Barbee, Brent W.; Sarli, Bruno V.; Lyzhoft, Joshua; Chodas, Paul W.; Englander, Jacob A.

2017-01-01

This paper presents a detailed mission design analysis results for the 2017 Planetary Defense Conference (PDC) Hypothetical Asteroid Impact Scenario, documented at https://cneos.jpl.nasa.gov/ pd/cs/pdc17/. The mission design includes campaigns for both reconnaissance (flyby or rendezvous) of the asteroid (to characterize it and the nature of the threat it poses to Earth) and mitigation of the asteroid, via kinetic impactor deflection, nuclear explosive device (NED) deflection, or NED disruption. Relevant scenario parameters are varied to assess the sensitivity of the design outcome, such as asteroid bulk density, asteroid diameter, momentum enhancement factor, spacecraft launch vehicle, and mitigation system type. Different trajectory types are evaluated in the mission design process from purely ballistic to those involving optimal midcourse maneuvers, planetary gravity assists, and/or lowthrust solar electric propulsion. The trajectory optimization is targeted around peak deflection points that were found through a novel linear numerical technique method. The optimization process includes constrain parameters, such as Earth departure date, launch declination, spacecraft/asteroid relative velocity and solar phase angle, spacecraft dry mass, minimum/maximum spacecraft distances from Sun and Earth, and Earth/spacecraft communications line of sight. Results show that one of the best options for the 2017 PDC deflection is solar electric propelled rendezvous mission with a single spacecraft using NED for the deflection

Distributed network scheduling

NASA Technical Reports Server (NTRS)

Clement, Bradley J.; Schaffer, Steven R.

2004-01-01

Distributed Network Scheduling is the scheduling of future communications of a network by nodes in the network. This report details software for doing this onboard spacecraft in a remote network. While prior work on distributed scheduling has been applied to remote spacecraft networks, the software reported here focuses on modeling communication activities in greater detail and including quality of service constraints. Our main results are based on a Mars network of spacecraft and include identifying a maximum opportunity of improving traverse exploration rate a factor of three; a simulation showing reduction in one-way delivery times from a rover to Earth from as much as 5 to 1.5 hours; simulated response to unexpected events averaging under an hour onboard; and ground schedule generation ranging from seconds to 50 minutes for 15 to 100 communication goals.

Optimization of the interplanetary trajectories of spacecraft with a solar electric propulsion power plant of minimal power

NASA Astrophysics Data System (ADS)

Ivanyukhin, A. V.; Petukhov, V. G.

2016-12-01

The problem of optimizing the interplanetary trajectories of a spacecraft (SC) with a solar electric propulsion system (SEPS) is examined. The problem of investigating the permissible power minimum of the solar electric propulsion power plant required for a successful flight is studied. Permissible ranges of thrust and exhaust velocity are analyzed for the given range of flight time and final mass of the spacecraft. The optimization is performed according to Portnyagin's maximum principle, and the continuation method is used for reducing the boundary problem of maximal principle to the Cauchy problem and to study the solution/ parameters dependence. Such a combination results in the robust algorithm that reduces the problem of trajectory optimization to the numerical integration of differential equations by the continuation method.

Development of Power Electronics for a 0.2kW-Class Ion Thruster

NASA Technical Reports Server (NTRS)

Pinero, Luis R.; Patterson, Michael J.; Bowers, Glen E.

1997-01-01

Applications that might benefit from low power ion propulsion systems include Earth-orbit magnetospheric mapping satellite constellations, low Earth-orbit satellites, geosynchronous Earth-orbit satellite north-south stationkeeping, and asteroid orbiters. These spacecraft are likely to have masses on the order of 50 to 500 kg with up to 0.5 kW of electrical power available. A power processing unit for a 0.2 kW-class ion thruster is currently under development for these applications. The first step in this effort is the development and testing of a 0.24 kW beam power supply. The design incorporates a 20 kHz full bridge topology with multiple secondaries connected in series to obtain outputs of up to 1200 V(sub DC). A current-mode control pulse width modulation circuit built using discrete components was selected for this application. An input voltage of 28 +/- 4 V(sub DC) was assumed, since the small spacecraft for which this system is targeted are anticipated to have unregulated low voltage busses. Efficiencies in excess of 91 percent were obtained at maximum output power. The total mass of the breadboard was less than 1.0 kg and the component mass was 0.53 kg. It is anticipated that a complete flight power processor could weigh about 2.0 kg.

X-Ray Detection and Processing Models for Spacecraft Navigation and Timing

NASA Technical Reports Server (NTRS)

Sheikh, Suneel; Hanson, John

2013-01-01

The current primary method of deepspace navigation is the NASA Deep Space Network (DSN). High-performance navigation is achieved using Delta Differential One-Way Range techniques that utilize simultaneous observations from multiple DSN sites, and incorporate observations of quasars near the line-of-sight to a spacecraft in order to improve the range and angle measurement accuracies. Over the past four decades, x-ray astronomers have identified a number of xray pulsars with pulsed emissions having stabilities comparable to atomic clocks. The x-ray pulsar-based navigation and time determination (XNAV) system uses phase measurements from these sources to establish autonomously the position of the detector, and thus the spacecraft, relative to a known reference frame, much as the Global Positioning System (GPS) uses phase measurements from radio signals from several satellites to establish the position of the user relative to an Earth-centered fixed frame of reference. While a GPS receiver uses an antenna to detect the radio signals, XNAV uses a detector array to capture the individual xray photons from the x-ray pulsars. The navigation solution relies on detailed xray source models, signal processing, navigation and timing algorithms, and analytical tools that form the basis of an autonomous XNAV system. Through previous XNAV development efforts, some techniques have been established to utilize a pulsar pulse time-of-arrival (TOA) measurement to correct a position estimate. One well-studied approach, based upon Kalman filter methods, optimally adjusts a dynamic orbit propagation solution based upon the offset in measured and predicted pulse TOA. In this delta position estimator scheme, previously estimated values of spacecraft position and velocity are utilized from an onboard orbit propagator. Using these estimated values, the detected arrival times at the spacecraft of pulses from a pulsar are compared to the predicted arrival times defined by the pulsar s pulse timing model. A discrepancy provides an estimate of the spacecraft position offset, since an error in position will relate to the measured time offset of a pulse along the line of sight to the pulsar. XNAV researchers have been developing additional enhanced approaches to process the photon TOAs to arrive at an estimate of spacecraft position, including those using maximum-likelihood estimation, digital phase locked loops, and "single photon processing" schemes that utilize all available time data associated with each photon. Using pulsars from separate, non-coplanar locations provides range and range-rate measurements in each pulsar s direction. Combining these different pulsar measurements solves for offsets in position and velocity in three dimensions, and provides accurate overall navigation for deep space vehicles.

Saturn Apollo Program

NASA Image and Video Library

1976-06-01

This illustration depicts the launch configuration of the Apollo spacecraft for the Apollo-Soyuz Test Project (ASTP). The ASTP was the first international docking of the U.S.'s Apollo spacecraft and the U.S.S.R.'s Soyuz spacecraft in space. A joint engineering team from the two countries met to develop a docking system that permitted the two spacecraft to link in space and allowed the two crews to travel from one spacecraft to the other. This system entailed developing a large habitable Docking Module (DM) to be carried on the Apollo spacecraft to facilitate the joining of two dissimilar spacecraft. The Marshall Space Flight Center was responsible for development and sustaining engineering of the Saturn IB launch vehicle during the mission.

Simple Statistical Model to Quantify Maximum Expected EMC in Spacecraft and Avionics Boxes

NASA Technical Reports Server (NTRS)

Trout, Dawn H.; Bremner, Paul

2014-01-01

This study shows cumulative distribution function (CDF) comparisons of composite a fairing electromagnetic field data obtained by computational electromagnetic 3D full wave modeling and laboratory testing. Test and model data correlation is shown. In addition, this presentation shows application of the power balance and extention of this method to predict the variance and maximum exptected mean of the E-field data. This is valuable for large scale evaluations of transmission inside cavities.

Prospective Out-of-ecliptic White-light Imaging of Interplanetary Corotating Interaction Regions at Solar Maximum

DOE Office of Scientific and Technical Information (OSTI.GOV)

Xiong, Ming; Yang, Liping; Liu, Ying D.

Interplanetary corotating interaction regions (CIRs) can be remotely imaged in white light (WL), as demonstrated by the Solar Mass Ejection Imager (SMEI) on board the Coriolis spacecraft and Heliospheric Imagers (HIs) on board the twin Solar TErrestrial RElations Observatory ( STEREO ) spacecraft. The interplanetary WL intensity, due to Thomson scattering of incident sunlight by free electrons, is jointly determined by the 3D distribution of electron number density and line-of-sight (LOS) weighting factors of the Thomson-scattering geometry. The 2D radiance patterns of CIRs in WL sky maps look very different from different 3D viewpoints. Because of the in-ecliptic locations ofmore » both the STEREO and Coriolis spacecraft, the longitudinal dimension of interplanetary CIRs has, up to now, always been integrated in WL imagery. To synthesize the WL radiance patterns of CIRs from an out-of-ecliptic (OOE) vantage point, we perform forward magnetohydrodynamic modeling of the 3D inner heliosphere during Carrington Rotation CR1967 at solar maximum. The mixing effects associated with viewing 3D CIRs are significantly minimized from an OOE viewpoint. Our forward modeling results demonstrate that OOE WL imaging from a latitude greater than 60° can (1) enable the garden-hose spiral morphology of CIRs to be readily resolved, (2) enable multiple coexisting CIRs to be differentiated, and (3) enable the continuous tracing of any interplanetary CIR back toward its coronal source. In particular, an OOE view in WL can reveal where nascent CIRs are formed in the extended corona and how these CIRs develop in interplanetary space. Therefore, a panoramic view from a suite of wide-field WL imagers in a solar polar orbit would be invaluable in unambiguously resolving the large-scale longitudinal structure of CIRs in the 3D inner heliosphere.« less

Design and Development of an Equipotential Voltage Reference (Grounding) System for a Low-Cost Rapid-Development Modular Spacecraft Architecture

NASA Technical Reports Server (NTRS)

Lukash, James A.; Daley, Earl

2011-01-01

This work describes the design and development effort to adapt rapid-development space hardware by creating a ground system using solutions of low complexity, mass, & cost. The Lunar Atmosphere and Dust Environment Explorer (LADEE) spacecraft is based on the modular common spacecraft bus architecture developed at NASA Ames Research Center. The challenge was building upon the existing modular common bus design and development work and improving the LADEE spacecraft design by adding an Equipotential Voltage Reference (EVeR) system, commonly referred to as a ground system. This would aid LADEE in meeting Electromagnetic Environmental Effects (E3) requirements, thereby making the spacecraft more compatible with itself and its space environment. The methods used to adapt existing hardware are presented, including provisions which may be used on future spacecraft.

Autonomous Aerobraking Using Thermal Response Surface Analysis

NASA Technical Reports Server (NTRS)

Prince, Jill L.; Dec, John A.; Tolson, Robert H.

2007-01-01

Aerobraking is a proven method of significantly increasing the science payload that can be placed into low Mars orbits when compared to an all propulsive capture. However, the aerobraking phase is long and has mission cost and risk implications. The main cost benefit is that aerobraking permits the use of a smaller and cheaper launch vehicle, but additional operational costs are incurred during the long aerobraking phase. Risk is increased due to the repeated thermal loading of spacecraft components and the multiple attitude and propulsive maneuvers required for successful aerobraking. Both the cost and risk burdens can be significantly reduced by automating the aerobraking operations phase. All of the previous Mars orbiter missions that have utilized aerobraking have increasingly relied on onboard calculations during aerobraking. Even though the temperature of spacecraft components has been the limiting factor, operational methods have relied on using a surrogate variable for mission control. This paper describes several methods, based directly on spacecraft component maximum temperature, for autonomously predicting the subsequent aerobraking orbits and prescribing apoapsis propulsive maneuvers to maintain the spacecraft within specified temperature limits. Specifically, this paper describes the use of thermal response surface analysis in predicting the temperature of the spacecraft components and the corresponding uncertainty in this temperature prediction.

Best-Fit Conic Approximation of Spacecraft Trajectory

NASA Technical Reports Server (NTRS)

Singh, Gurkipal

2005-01-01

A computer program calculates a best conic fit of a given spacecraft trajectory. Spacecraft trajectories are often propagated as conics onboard. The conic-section parameters as a result of the best-conic-fit are uplinked to computers aboard the spacecraft for use in updating predictions of the spacecraft trajectory for operational purposes. In the initial application for which this program was written, there is a requirement to fit a single conic section (necessitated by onboard memory constraints) accurate within 200 microradians to a sequence of positions measured over a 4.7-hour interval. The present program supplants a prior one that could not cover the interval with fewer than four successive conic sections. The present program is based on formulating the best-fit conic problem as a parameter-optimization problem and solving the problem numerically, on the ground, by use of a modified steepest-descent algorithm. For the purpose of this algorithm, optimization is defined as minimization of the maximum directional propagation error across the fit interval. In the specific initial application, the program generates a single 4.7-hour conic, the directional propagation of which is accurate to within 34 microradians easily exceeding the mission constraints by a wide margin.

Life test failure of harmonic gears in a Two-axis Gimbal for the Mars Reconnaissance Orbiter Spacecraft

NASA Technical Reports Server (NTRS)

Johnson, Michael R.; Gehling, Russ; Head, Ray

2006-01-01

This paper will present a process for increasing the stiffness of harmonic gear assemblies and recommend a maximum stiffness point that, if exceeded, compromises the reliability of the gear components for long life applications.

Lyman alpha SMM/UVSP absolute calibration and geocoronal correction

NASA Technical Reports Server (NTRS)

Fontenla, Juan M.; Reichmann, Edwin J.

1987-01-01

Lyman alpha observations from the Ultraviolet Spectrometer Polarimeter (UVSP) instrument of the Solar Maximum Mission (SMM) spacecraft were analyzed and provide instrumental calibration details. Specific values of the instrument quantum efficiency, Lyman alpha absolute intensity, and correction for geocoronal absorption are presented.

Spacecraft-spacecraft radio-metric tracking: Signal acquisition requirements and application to Mars approach navigation

NASA Technical Reports Server (NTRS)

Kahn, R. D.; Thurman, S.; Edwards, C.

1994-01-01

Doppler and ranging measurements between spacecraft can be obtained only when the ratio of the total received signal power to noise power density (P(sub t)/N(sub 0)) at the receiving spacecraft is sufficiently large that reliable signal detection can be achieved within a reasonable time period. In this article, the requirement on P(sub t)/N(sub 0) for reliable carrier signal detection is calculated as a function of various system parameters, including characteristics of the spacecraft computing hardware and a priori uncertainty in spacecraft-spacecraft relative velocity and acceleration. Also calculated is the P(sub t)/N(sub 0) requirements for reliable detection of a ranging signal, consisting of a carrier with pseudonoise (PN) phase modulation. Once the P(sub t)/N(sub 0) requirement is determined, then for a given set of assumed spacecraft telecommunication characteristics (transmitted signal power, antenna gains, and receiver noise temperatures) it is possible to calculate the maximum range at which a carrier signal or ranging signal may be acquired. For example, if a Mars lander and a spacecraft approaching Mars are each equipped with 1-m-diameter antennas, the transmitted power is 5 W, and the receiver noise temperatures are 350 K, then S-band carrier signal acquisition can be achieved at ranges exceeding 10 million km. An error covariance analysis illustrates the utility of in situ Doppler and ranging measurements for Mars approach navigation. Covariance analysis results indicate that navigation accuracies of a few km can be achieved with either data type. The analysis also illustrates dependency of the achievable accuracy on the approach trajectory velocity.

Advanced nickel-hydrogen spacecraft battery development

NASA Technical Reports Server (NTRS)

Coates, Dwaine K.; Fox, Chris L.; Standlee, D. J.; Grindstaff, B. K.

1994-01-01

Eagle-Picher currently has several advanced nickel-hydrogen (NiH2) cell component and battery designs under development including common pressure vessel (CPV), single pressure vessel (SPV), and dependent pressure vessel (DPV) designs. A CPV NiH2 battery, utilizing low-cost 64 mm (2.5 in.) cell diameter technology, has been designed and built for multiple smallsat programs, including the TUBSAT B spacecraft which is currently scheduled (24 Nov. 93) for launch aboard a Russian Proton rocket. An advanced 90 mm (3.5 in.) NiH2 cell design is currently being manufactured for the Space Station Freedom program. Prototype 254 mm (10 in.) diameter SPV batteries are currently under construction and initial boilerplate testing has shown excellent results. NiH2 cycle life testing is being continued at Eagle-Picher and IPV cells have currently completed more than 89,000 accelerated LEO cycles at 15% DOD, 49,000 real-time LEO cycles at 30 percent DOD, 37,800 cycles under a real-time LEO profile, 30 eclipse seasons in accelerated GEO, and 6 eclipse seasons in real-time GEO testing at 75 percent DOD maximum. Nickel-metal hydride battery development is continuing for both aerospace and electric vehicle applications. Eagle-Picher has also developed an extensive range of battery evaluation, test, and analysis (BETA) measurement and control equipment and software, based on Hewlett-Packard computerized data acquisition/control hardware.

Development of an expert system prototype for determining software functional requirements for command management activities at NASA Goddard

NASA Technical Reports Server (NTRS)

Liebowitz, J.

1986-01-01

The development of an expert system prototype for software functional requirement determination for NASA Goddard's Command Management System, as part of its process of transforming general requests into specific near-earth satellite commands, is described. The present knowledge base was formulated through interactions with domain experts, and was then linked to the existing Knowledge Engineering Systems (KES) expert system application generator. Steps in the knowledge-base development include problem-oriented attribute hierarchy development, knowledge management approach determination, and knowledge base encoding. The KES Parser and Inspector, in addition to backcasting and analogical mapping, were used to validate the expert system-derived requirements for one of the major functions of a spacecraft, the solar Maximum Mission. Knowledge refinement, evaluation, and implementation procedures of the expert system were then accomplished.

Long-term downward trend in total solar irradiance.

PubMed

Willson, R C; Hudson, H S; Frohlich, C; Brusa, R W

1986-11-28

The first 5 years (from 1980 to 1985) of total solar irradiance observations by the first Active Cavity Radiometer Irradiance Monitor (ACRIM I) experiment on board the Solar Maximum Mission spacecraft show a clearly defined downward trend of -0.019% per year. The existence of this trend has been confirmed by the internal self-calibrations of ACRIM I, by independent measurements from sounding rockets and balloons, and by observations from the Nimbus-7 spacecraft. The trend appears to be due to unpredicted variations of solar luminosity on time scales of years, and it may be related to solar cycle magnetic activity.

Development of concepts for satellite retrieval devices

NASA Technical Reports Server (NTRS)

Pruett, E. C.; Robertson, K. B., III; Loughead, T. E.

1979-01-01

The teleoperator being developed to augment the Space Transportation System (STS) for satellite placement, retrieval, or servicing at altitudes or orbital planes where it would be impractical to use the shuttle is primarily a general purpose propulsion stage that can be fitted with manipulator arms, automated servicers and satellite retrieval devices for particular missions. Design concepts for a general purpose retrieval device for docking with a satellite to which a grappling fixture has been attached, and for a retrieval device for docking with the Solar Maximum Mission (SMM) spacecraft were defined. The mechanical aspects of these two devices are discussed as well as the crew operations involved and problems created by the requirement for remote control. Drawings for the two retrieval device concepts are included.

Apollo LM guidance computer software for the final lunar descent.

NASA Technical Reports Server (NTRS)

Eyles, D.

1973-01-01

In all manned lunar landings to date, the lunar module Commander has taken partial manual control of the spacecraft during the final stage of the descent, below roughly 500 ft altitude. This report describes programs developed at the Charles Stark Draper Laboratory, MIT, for use in the LM's guidance computer during the final descent. At this time computational demands on the on-board computer are at a maximum, and particularly close interaction with the crew is necessary. The emphasis is on the design of the computer software rather than on justification of the particular guidance algorithms employed. After the computer and the mission have been introduced, the current configuration of the final landing programs and an advanced version developed experimentally by the author are described.

Developing Architectures and Technologies for an Evolvable NASA Space Communication Infrastructure

NASA Technical Reports Server (NTRS)

Bhasin, Kul; Hayden, Jeffrey

2004-01-01

Space communications architecture concepts play a key role in the development and deployment of NASA's future exploration and science missions. Once a mission is deployed, the communication link to the user needs to provide maximum information delivery and flexibility to handle the expected large and complex data sets and to enable direct interaction with the spacecraft and experiments. In human and robotic missions, communication systems need to offer maximum reliability with robust two-way links for software uploads and virtual interactions. Identifying the capabilities to cost effectively meet the demanding space communication needs of 21st century missions, proper formulation of the requirements for these missions, and identifying the early technology developments that will be needed can only be resolved with architecture design. This paper will describe the development of evolvable space communication architecture models and the technologies needed to support Earth sensor web and collaborative observation formation missions; robotic scientific missions for detailed investigation of planets, moons, and small bodies in the solar system; human missions for exploration of the Moon, Mars, Ganymede, Callisto, and asteroids; human settlements in space, on the Moon, and on Mars; and great in-space observatories for observing other star systems and the universe. The resulting architectures will enable the reliable, multipoint, high data rate capabilities needed on demand to provide continuous, maximum coverage of areas of concentrated activities, such as in the vicinity of outposts in-space, on the Moon or on Mars.

Approaching Solar Maximum 24 with Stereo-Multipoint Observations of Solar Energetic Particle Events

NASA Technical Reports Server (NTRS)

Dresing, N.; Cohen, C. M. S.; Gomez-Herrero, R.; Heber, B.; Klassen, A.; Leske, R. A.; Mason, G. M.; Mewaldt, R. A.; von Rosenvinge, T. T.

2014-01-01

Since the beginning of the Solar Terrestrial Relations Observatory (STEREO) mission at the end of 2006, the two spacecraft have now separated by more than 130? degrees from the Earth. A 360-degree view of the Sun has been possible since February 2011, providing multipoint in situ and remote sensing observations of unprecedented quality. Combining STEREO observations with near-Earth measurements allows the study of solar energetic particle (SEP) events over a wide longitudinal range with minimal radial gradient effects. This contribution provides an overview of recent results obtained by the STEREO/IMPACT team in combination with observations by the ACE and SOHO spacecraft. We focus especially on multi-spacecraft investigations of SEP events. The large longitudinal spread of electron and 3He-rich events as well as unusual anisotropies will be presented and discussed.

Re-scheduling as a tool for the power management on board a spacecraft

NASA Technical Reports Server (NTRS)

Albasheer, Omar; Momoh, James A.

1995-01-01

The scheduling of events on board a spacecraft is based on forecast energy levels. The real time values of energy may not coincide with the forecast values; consequently, a dynamic revising to the allocation of power is needed. The re-scheduling is also needed for other reasons on board a spacecraft like the addition of new event which must be scheduled, or a failure of an event due to many different contingencies. This need of rescheduling is very important to the survivability of the spacecraft. In this presentation, a re-scheduling tool will be presented as a part of an overall scheme for the power management on board a spacecraft from the allocation of energy point of view. The overall scheme is based on the optimal use of energy available on board a spacecraft using expert systems combined with linear optimization techniques. The system will be able to schedule maximum number of events utilizing most energy available. The outcome is more events scheduled to share the operation cost of that spacecraft. The system will also be able to re-schedule in case of a contingency with minimal time and minimal disturbance of the original schedule. The end product is a fully integrated planning system capable of producing the right decisions in short time with less human error. The overall system will be presented with the re-scheduling algorithm discussed in detail, then the tests and results will be presented for validations.

Technology for small spacecraft

NASA Technical Reports Server (NTRS)

1994-01-01

This report gives the results of a study by the National Research Council's Panel on Small Spacecraft Technology that reviewed NASA's technology development program for small spacecraft and assessed technology within the U.S. government and industry that is applicable to small spacecraft. The panel found that there is a considerable body of advanced technology currently available for application by NASA and the small spacecraft industry that could provide substantial improvement in capability and cost over those technologies used for current NASA small spacecraft. These technologies are the result of developments by commercial companies, Department of Defense agencies, and to a lesser degree NASA. The panel also found that additional technologies are being developed by these same entities that could provide additional substantial improvement if development is successfully completed. Recommendations for future technology development efforts by NASA across a broad technological spectrum are made.

Economic and technical aspects of repair, servicing, and retrieval of low earth orbit free flying spacecraft

NASA Technical Reports Server (NTRS)

Cepollina, F. J.

1982-01-01

The economic and technical aspects of the Solar Maximum Observatory Repair Mission at NASA are presented, in an effort to demonstrate the Space Shuttle capability to rendezvous with and repair on-orbit the Solar Maximum Observatory (SMM). A failure in the Attitude Control Subsystem (ACS) after 10 months of operation caused a loss in precision pointing capability. The Multimission Modular Spacecraft (MMS) used for the mission, was designed with on-orbit repairability, and to correct various instrument anomalies, repiar kits such as an electronics box, a thermal aperture closure, and a high energy particle reflection baffle will be used. In addition, a flight support system will be used to berth, electrically safe, and support all the repair activities. A two year effort is foreseen, and the economic return on SMM will be $176 M, in addition to two to three years of solar observation. The mission will eventually conduct studies on flare as a function of solar cycle.

Proceedings of the Spacecraft Charging Technology Conference

NASA Technical Reports Server (NTRS)

Pike, C. P. (Editor); Lovell, R. R. (Editor)

1977-01-01

Over 50 papers from the spacecraft charging conference are included on subjects such as: (1) geosynchronous plasma environment, (2) spacecraft modeling, (3) spacecraft materials characterization, (4) spacecraft materials development, and (5) satellite design and test.

Pioneer Venus data analysis for the retarding potential analyzer

NASA Technical Reports Server (NTRS)

Knudsen, William C.

1993-01-01

This report describes the data analysis and archiving activities, analysis results, and instrument performance of the orbiter retarding potential analyzer (ORPA) flown on the Pioneer Venus Orbiter spacecraft during the period, Aug. 1, 1988 to Sept. 30, 1993. During this period, the periapsis altitude of the Orbiter spacecraft descended slowly from 1900 km altitude, at which altitude the spacecraft was outside the Venus ionosphere, to approximately 130 km altitude in Oct. 1992 at which time communication with the spacecraft ceased as a result of entry of the spacecraft into the Venus atmosphere. The quantity of ORPA data returned during this reporting period was greatly reduced over that recovered in the previous years of the mission because of the reduced power capability of the spacecraft, loss of half of the onboard data storage, and partial failure of the ORPA. Despite the reduction in available data, especially ionospheric data, important scientific discoveries resulted from this extended period of the Pioneer Venus mission. The most significant discovery was that of a strong solar cycle change in the size of the dayside ionosphere and the resulting shutoff of flow of dayside ions into the nightside hemisphere. The large, topside O+ F2 ionospheric layer observed during the first three years of the Pioneer Venus mission, a period of solar cycle maximum activity, is absent during the solar cycle minimum activity period.

Spacecraft Charging Hazards In Low-earth Orbit

NASA Astrophysics Data System (ADS)

Anderson, P. C.

The space environment in low-Earth orbit (LEO) has until recently been considered quite benign to high levels of spacecraft charging. However, it has been found that the DMSP spacecraft at 840 km can charge to very large negative voltages (up to - 2000 V) when encountering intense precipitating electron events (auroral arcs) while traversing the auroral zone. The occurrence frequency of charging events, defined as when the spacecraft charged to levels exceeding 100 V negative, was highly correlated with the 11-year solar cycle with the largest number of events occurring during solar minimum. This was due to the requirement that the background thermal plasma den- sity be low, at most 104 cm-2. During solar maximum, the plasma density is typically well above that level due to the solar EUV ionizing radiation, and although the oc- currence frequency of auroral arcs is considerably greater than at solar minimum, the occurrence of high-level charging is minimal. Indeed, of the over 1200 events found during the most recent solar cycle, none occurred during the last solar maximum. This has implications to a number of LEO satellite programs, including the International Space Station (ISS). The plasma density in the ISS orbit, at a much lower altitude than DMSP, is well above that at 840 km and rarely below 104 cm-2. However, in the wake of the ISS, the plasma density can be 2 orders of magnitude or more lower than the background density and thus conditions are ripe for significant charging effects. With an inclination of 51.6 degrees, the ISS does enter the auroral zone, particularly during geomagnetic storms and substorms when the auroral boundary can penetrate to very low latitudes. This has significant implications for EVA operations in the ISS wake.

MIDEX Advanced Modular and Distributed Spacecraft Avionics Architecture

NASA Technical Reports Server (NTRS)

Ruffa, John A.; Castell, Karen; Flatley, Thomas; Lin, Michael

1998-01-01

MIDEX (Medium Class Explorer) is the newest line in NASA's Explorer spacecraft development program. As part of the MIDEX charter, the MIDEX spacecraft development team has developed a new modular, distributed, and scaleable spacecraft architecture that pioneers new spaceflight technologies and implementation approaches, all designed to reduce overall spacecraft cost while increasing overall functional capability. This resultant "plug and play" system dramatically decreases the complexity and duration of spacecraft integration and test, providing a basic framework that supports spacecraft modularity and scalability for missions of varying size and complexity. Together, these subsystems form a modular, flexible avionics suite that can be modified and expanded to support low-end and very high-end mission requirements with a minimum of redesign, as well as allowing a smooth, continuous infusion of new technologies as they are developed without redesigning the system. This overall approach has the net benefit of allowing a greater portion of the overall mission budget to be allocated to mission science instead of a spacecraft bus. The MIDEX scaleable architecture is currently being manufactured and tested for use on the Microwave Anisotropy Probe (MAP), an inhouse program at GSFC.

Temperature dependence of attitude sensor coalignments on the Solar Maximum Mission (SMM)

NASA Technical Reports Server (NTRS)

Pitone, D. S.; Eudell, A. H.; Patt, F. S.

1989-01-01

Results are presented on the temperature correlation of the relative coalignment between the fine pointing sun sensor (FPSS) and fixed head star trackers (FHSTs) on the Solar Maximum Mission (SMM). This correlation can be caused by spacecraft electronic and mechanical effects. Routine daily measurements reveal a time dependent sensor coalignment variation. The magnitude of the alignment variation is on the order of 120 arc seconds (arc sec), which greatly exceeds the prelaunch thermal structural analysis estimate of 15 acr sec. Differences between FPSS-only and FHST-only yaw solutions as a function of mission day are correlated with the relevant spacecraft temperature. If unaccounted for, the sensor misalignments due to thermal effects are a significant source of error in attitude determination accuracy. Prominent sources of temperature variation are identified and correlated with the temperature profile observed on the SMM. It was determined that even relatively small changes in spacecraft temperature can affect the coalignments between the attitude hardware on the SMM and the science instrument support plate and that frequent recalibration of sensor alignments is necessary to compensate for this effect. An alterntive to frequent recalibration is to model the variation of alignments as a function of temperature and use this to maintain accurate ground or onboard alignment estimates. These flight data analysis results may be important consierations for prelaunch analysis of future missions.

An Empirical Comparison between Two Recursive Filters for Attitude and Rate Estimation of Spinning Spacecraft

NASA Technical Reports Server (NTRS)

Harman, Richard R.

2006-01-01

The advantages of inducing a constant spin rate on a spacecraft are well known. A variety of science missions have used this technique as a relatively low cost method for conducting science. Starting in the late 1970s, NASA focused on building spacecraft using 3-axis control as opposed to the single-axis control mentioned above. Considerable effort was expended toward sensor and control system development, as well as the development of ground systems to independently process the data. As a result, spinning spacecraft development and their resulting ground system development stagnated. In the 1990s, shrinking budgets made spinning spacecraft an attractive option for science. The attitude requirements for recent spinning spacecraft are more stringent and the ground systems must be enhanced in order to provide the necessary attitude estimation accuracy. Since spinning spacecraft (SC) typically have no gyroscopes for measuring attitude rate, any new estimator would need to rely on the spacecraft dynamics equations. One estimation technique that utilized the SC dynamics and has been used successfully in 3-axis gyro-less spacecraft ground systems is the pseudo-linear Kalman filter algorithm. Consequently, a pseudo-linear Kalman filter has been developed which directly estimates the spacecraft attitude quaternion and rate for a spinning SC. Recently, a filter using Markley variables was developed specifically for spinning spacecraft. The pseudo-linear Kalman filter has the advantage of being easier to implement but estimates the quaternion which, due to the relatively high spinning rate, changes rapidly for a spinning spacecraft. The Markley variable filter is more complicated to implement but, being based on the SC angular momentum, estimates parameters which vary slowly. This paper presents a comparison of the performance of these two filters. Monte-Carlo simulation runs will be presented which demonstrate the advantages and disadvantages of both filters.

Quantification of encapsulated bioburden in spacecraft polymer materials by cultivation-dependent and molecular methods.

PubMed

Bauermeister, Anja; Mahnert, Alexander; Auerbach, Anna; Böker, Alexander; Flier, Niwin; Weber, Christina; Probst, Alexander J; Moissl-Eichinger, Christine; Haberer, Klaus

2014-01-01

Bioburden encapsulated in spacecraft polymers (such as adhesives and coatings) poses a potential risk to jeopardize scientific exploration of other celestial bodies. This is particularly critical for spacecraft components intended for hard landing. So far, it remained unclear if polymers are indeed a source of microbial contamination. In addition, data with respect to survival of microbes during the embedding/polymerization process are sparse. In this study we developed testing strategies to quantitatively examine encapsulated bioburden in five different polymers used frequently and in large quantities on spaceflight hardware. As quantitative extraction of the bioburden from polymerized (solid) materials did not prove feasible, contaminants were extracted from uncured precursors. Cultivation-based analyses revealed <0.1-2.5 colony forming units (cfu) per cm3 polymer, whereas quantitative PCR-based detection of contaminants indicated considerably higher values, despite low DNA extraction efficiency. Results obtained from this approach reflect the most conservative proxy for encapsulated bioburden, as they give the maximum bioburden of the polymers irrespective of any additional physical and chemical stress occurring during polymerization. To address the latter issue, we deployed an embedding model to elucidate and monitor the physiological status of embedded Bacillus safensis spores in a cured polymer. Staining approaches using AlexaFluor succinimidyl ester 488 (AF488), propidium monoazide (PMA), CTC (5-cyano-2,3-diotolyl tetrazolium chloride) demonstrated that embedded spores retained integrity, germination and cultivation ability even after polymerization of the adhesive Scotch-Weld 2216 B/A. Using the methods presented here, we were able to estimate the worst case contribution of encapsulated bioburden in different polymers to the bioburden of spacecraft. We demonstrated that spores were not affected by polymerization processes. Besides Planetary Protection considerations, our results could prove useful for the manufacturing of food packaging, pharmacy industry and implant technology.

Development of Rene 41 honeycomb structure as an integral cryogenic tankage/fuselage concept for future space transportation systems

NASA Technical Reports Server (NTRS)

Shideler, J. L.; Swegle, A. R.; Fields, R. A.

1982-01-01

The status of the structural development of an integral cryogenic-tankage/hot-fuselage concept for future space transportation systems is reviewed. The concept comprises a honeycomb sandwich structure that serves the combined functions of containing the cryogenic fuel, supporting the vehicle loads, and protecting the spacecraft from entry heating. The inner face sheet is exposed to cryogenic temperature of -423 F during boost; the outer face sheet, which is slotted to reduce thermal stress, is exposed to a maximum temperature of 1400 F during a high-altitude gliding entry. Attention is given to the development of a fabrication process for a Rene 41 honeycomb sandwich panel with a core density of less than 1 percent that is consistent with desirable heat treatment processes for high strength.

14 CFR 1216.305 - Criteria for actions requiring environmental assessments.

Code of Federal Regulations, 2012 CFR

2012-01-01

... assessment are: (1) Specific spacecraft development and flight projects in space science. (2) Specific spacecraft development and flight projects in space and terrestrial applications. (3) Specific experimental... spacecraft development and flight projects. (2) R&D activities in space and terrestrial applications (e.g...

Crew Exploration Vehicle (CEV) Water Landing Simulation

NASA Technical Reports Server (NTRS)

Littell, Justin D.; Lawrence, Charles; Carney, Kelly S.

2007-01-01

Crew Exploration Vehicle (CEV) water splashdowns were simulated in order to find maximum acceleration loads on the astronauts and spacecraft under various landing conditions. The acceleration loads were used in a Dynamic Risk Index (DRI) program to find the potential risk for injury posed on the astronauts for a range of landing conditions. The DRI results showed that greater risks for injury occurred for two landing conditions; when the vertical velocity was large and the contact angle between the spacecraft and the water impact surface was zero, and when the spacecraft was in a toe down configuration and both the vertical and horizontal landing velocities were large. Rollover was also predicted to occur for cases where there is high horizontal velocity and low contact angles in a toe up configuration, and cases where there was a high horizontal velocity with high contact angles in a toe down configuration.

Mars atmospheric losses induced by the solar wind: current knowledge and perspective

NASA Astrophysics Data System (ADS)

Ermakov, Vladimir; Zelenyi, Lev; Vaisberg, Oleg; Sementsov, Egor; Dubinin, Eduard

2017-04-01

Solar wind induced atmospheric losses have been studied since earlier 1970th. Several loss channels have been identified including pick-up of exospheric photo-ions and ionospheric ions escape. Measurements performed during several solar cycles showed variation of these losses by about factor of 10, being largest at maximum solar activity. MAVEN spacecraft equipped with comprehensive set of instruments with high temporal and mass resolution operating at Mars since fall 2014 ensures much better investigation of solar wind enforcing Martian environment, Mars atmospheric losses processes and mass loss rate. These issues are very important for understanding of Martian atmospheric evolution including water loss during cosmogonic time. Simultaneous observations by MAVEN and MEX spacecraft open the new perspective in study of Martian environment. In this report we discuss results of past and current missions and preliminary analysis of heavy ions escape using simultaneous measurements of MEX and MAVEN spacecraft.

Asteroid Impact Mission (aim) & Deflection Assessment: AN Opportunity to Understand Impact Dynamics and Modelling

NASA Astrophysics Data System (ADS)

Galvez, A.; Carnelli, I.; Fontaine, M.; Corral Van Damme, C.

2012-09-01

ESA's Future Preparation and Strategic Studies Office has carried out the Asteroid Impact Mission (AIM) study with the objective of defining an affordable and fully independent mission element that ESA could contribute to an Asteroid Impact Deflection Assessment campaign (AIDA), a joint effort of ESA, JHU/APL, NASA, OCA and DLR. The mission design foresees two independent spacecraft, one impactor (DART) and one rendezvous probe (AIM). The target of this mission is the binary asteroid system (65803) Didymos (1996 GT): one spacecraft, DART, would impact the secondary of the Didymos binary system while AIM would observe and measure any the change in the relative orbit. For this joint project, the timing of the experiment is set (maximum proximity of the target to Earth allowing for ground-based characterisation of the experiment) but the spacecraft are still able to pursue their missions fully independently. This paper describes in particular the AIM rendezvous mission concept.

Mid-L/D Lifting Body Entry Demise Analysis

NASA Technical Reports Server (NTRS)

Ling, Lisa

2017-01-01

The mid-lift-to-drag ratio (mid-L/D) lifting body is a fully autonomous spacecraft under design at NASA for enabling a rapid return of scientific payloads from the International Space Station (ISS). For contingency planning and risk assessment for the Earth-return trajectory, an entry demise analysis was performed to examine three potential failure scenarios: (1) nominal entry interface conditions with loss of control, (2) controlled entry at maximum flight path angle, and (3) controlled entry at minimum flight path angle. The objectives of the analysis were to predict the spacecraft breakup sequence and timeline, determine debris survival, and calculate the debris dispersion footprint. Sensitivity analysis was also performed to determine the effect of the initial pitch rate on the spacecraft stability and breakup during the entry. This report describes the mid-L/D lifting body and presents the results of the entry demise and sensitivity analyses.

Saturn Apollo Program

NASA Image and Video Library

1974-06-01

This illustration shows the docking configuration of the Apollo-Soyuz Test Project (ASTP). The ASTP was the first international docking of the U.S.'s Apollo spacecraft and the U.S.S.R.'s Soyuz spacecraft in space. A joint engineering team from the two countries met to develop a docking system that permitted the two spacecraft to link in space and allowed the two crews to travel from one spacecraft to the other. This system entailed developing a large habitable Docking Module (DM) to be carried on the Apollo spacecraft to facilitate the joining of two dissimilar spacecraft. The Marshall Space Flight Center was responsible for development and sustaining engineering of the Saturn IB launch vehicle during the mission. The ASTP marked the last use of the Saturn Launch Vehicle.

The 2011 Draconid Shower Risk to Earth-Orbiting Satellites

NASA Technical Reports Server (NTRS)

Cooke, William J.; Moser, Danielle E.

2010-01-01

Current meteor shower forecast models project a strong Draconid outburst, possibly a storm, on October 8, 2011, with a duration of approximately 7 hours and peaking between 19 and 21 hours UT. Predicted rates span an order of magnitude, with maximum Zenithal Hourly Rates (ZHRs) ranging from a few tens to several hundred. Calibration of the NASA MSFC Meteoroid Stream Model 1 to radar and optical observations of past apparitions, particularly the 2005 Draconid outburst 2, suggest that the maximum rate will be several hundreds per hour. Given the high spatial density of the Draconid stream, this implies a maximum meteoroid flux of 5-10 Draconids km(exp -2)/hr (to a limiting diameter of 1 mm), some 25-50 times greater than the normal sporadic flux of 0.2 km(exp -2)/ hr for particles of this size. Total outburst fluence, assuming a maximum ZHR of 750, is 15.5 Draconids km(exp -2), resulting in an overall 10x risk increase to spacecraft surfaces vulnerable to hypervelocity impacts by 1 mm particles. It is now established that a significant fraction of spacecraft anomalies produced by shower meteoroids (e.g. OLYMPUS and LandSat 5) are caused by electrostatic discharges produced by meteoroid impacts. In these cases, the charge generated is roughly proportional to v(exp 3.5(4)), giving a Draconid moving at 20 km/s approximately 1/80th the electrical damage potential of a Leonid of the same mass. In other words, a Draconid outburst with a maximum ZHR of 800 presents the same electrical risk as a normal Leonid shower with a ZHR of 15, assuming the mass indices and shower durations are the same. This is supported by the fact that no spacecraft electrical anomalies were reported during the strong Draconid outbursts of 1985 and 1998. However, the lack of past anomalies should not be taken as carte blanche for satellite operators to ignore the 2011 Draconids, as the upcoming outburst will constitute a period of enhanced risk for vehicles in near-Earth space. Each spacecrft is unique, and components have differing damage thresholds; programs are encouraged to conduct analyses to determine whether or not mitigation strategies are necessary for their vehicles.

How autonomy and the web are taking the people out of TacSat-2

NASA Astrophysics Data System (ADS)

Finley, Charles J.

2006-05-01

One of the most costly components of the on-orbit operation of a spacecraft is the people that execute the mission. Historically, for Air Force Research Laboratory (AFRL) and the Department of Defense Space Test Program (STP) research and development, test and evaluation (RDT&E) space missions, a team of fifteen personnel maintains 24-hour coverage for the three-week Launch and Early Operations (L/EO) phase of the mission and four one-week L/EO rehearsals. During the Nominal Operations phase of the mission, 2.5 "man-days" of support are necessary each day that the spacecraft remains on-orbit, as well as during the two, week-long, nominal operations rehearsals. Therefore, the mission-dedicated personnel contribution to the cost of a one-year mission is more than eleven man-years, and this does not include the personnel that actually operate the antennas at the various remote ground facilities or develop and maintain the mission-specific or shared-use ground network, hardware, and software. In the low-budget RDT&E world, hardware, software, or Concept of Operations (CONOPS) developments that significantly reduce the necessary Operations personnel investment can mean the difference between a mission that does or does not survive. This paper explores the CONOPS and suite of tools that the TacSat-2 program has put together to achieve maximum mission effectiveness at minimum manpower cost.

Report of the committee on a commercially developed space facility

NASA Technical Reports Server (NTRS)

Shea, Joseph F.; Stever, H. Guyford; Cutter, W. Bowman, III; Demisch, Wolfgang H.; Fink, Daniel J.; Flax, Alexander H.; Gatos, Harry C.; Glicksman, Martin E.; Lanzerotti, Louis J.; Logsdon, John M., III

1989-01-01

Major facilities that could support significant microgravity research and applications activity are discussed. The ground-based facilities include drop towers, aircraft flying parabolic trajectories, and sounding rockets. Facilities that are intrinsically tied to the Space Shuttle range from Get-Away-Special canisters to Spacelab long modules. There are also orbital facilities which include recoverable capsules launched on expendable launch vehicles, free-flying spacecraft, and space stations. Some of these existing, planned, and proposed facilities are non-U.S. in origin, but potentially available to U.S. investigators. In addition, some are governmentally developed and operated whereas others are planned to be privately developed and/or operated. Tables are provided to show the facility, developer, duration, estimated gravity level, crew interaction, flight frequency, year available, power to payload, payload volume, and maximum payload mass. The potential of direct and indirect benefits of manufacturing in space are presented.

The Pioneer XI high field fluxgate magnetometer

NASA Technical Reports Server (NTRS)

Acuna, M. A.; Ness, N. F.

1975-01-01

The high field fluxgate magnetometer experiment flown aboard the Pioneer XI spacecraft is described. This extremely simple instrument was used to extend the spacecraft's upper-limit measurement capability by approximately an order of magnitude (from 0.14 mT to 1.00 mT) with minimum power and volume requirements. This magnetometer was designed to complement the low-field measurements provided by a helium vector magnetometer and utilizes magnetic ring core sensors with biaxial orthogonal sense coils. The instrument is a single-range, triaxial-fluxgate magnetometer capable of measuring fields of up to 1 mT along each orthogonal axis, with a maximum resolution of 1 microT.

Create your own science planning tool in 3 days with SOA

NASA Technical Reports Server (NTRS)

Streiffert, Barbara A.; Polanskey, Carol A.; O'Reilly, Taifun

2003-01-01

Scientific discovery and advancement of knowledge has been, and continues to be, the goal for space missions at Jet Propulsion Laboratory. Scientist must plan their observation/experiments to get the maximum data return in order to make those discoveries. However, each mission has different science objectives, a different spacecraft and different instrument payloads, as well as, different routes to different destinations with different spacecraft restrictions and characteristics. In the current reduced cost environment, manageable cost for mission planning software is a must. Science Opportunity Analyzer (SOA), a planning tool for scientists and mission planners, utilizes a simple approach to reduce cost and promote reusability.

Electrostatic-probe measurements of plasma parameters for two reentry flight experiments at 25000 feet per second

NASA Technical Reports Server (NTRS)

Jones, W. L., Jr.; Cross, A. E.

1972-01-01

Unique plasma diagnostic measurements at high altitudes from two geometrically similar blunt body reentry spacecraft using electrostatic probe rakes are presented. The probes measured the positive ion density profiles (shape and magnitude) during the two flights. The probe measurements were made at eight discrete points (1 cm to 7 cm) from the vehicle surface in the aft flow field of the spacecraft over the altitude range of 85.3 to 53.3 km (280,000 to 175,000 ft) with measured densities of 10 to the 8th power to 10 to the 12th power electrons/cu cm, respectively. Maximum reentry velocity for each spacecraft was approximately 7620 meters/second (25,000 ft/sec). In the first flight experiment, water was periodically injected into a flow field which was contaminated by ablation products from the spacecraft nose region. The nonablative nose of the second spacecraft thereby minimized flow field contamination. Comparisons of the probe measured density profiles with theoretical calculations are presented with discussion as to the probable cause of significant disagreement. Also discussed are the correlation of probe measurements with vehicle angle of attack motions and the good high altitude agreement between electron densities inferred from the probe measurements, VHF antenna measurements, and microwave reflectometer diagnostic measurements.

Dissertation Defense Computational Fluid Dynamics Uncertainty Analysis for Payload Fairing Spacecraft Environmental Control Systems

NASA Technical Reports Server (NTRS)

Groves, Curtis Edward

2014-01-01

Spacecraft thermal protection systems are at risk of being damaged due to airflow produced from Environmental Control Systems. There are inherent uncertainties and errors associated with using Computational Fluid Dynamics to predict the airflow field around a spacecraft from the Environmental Control System. This paper describes an approach to quantify the uncertainty in using Computational Fluid Dynamics to predict airflow speeds around an encapsulated spacecraft without the use of test data. Quantifying the uncertainty in analytical predictions is imperative to the success of any simulation-based product. The method could provide an alternative to traditional "validation by test only" mentality. This method could be extended to other disciplines and has potential to provide uncertainty for any numerical simulation, thus lowering the cost of performing these verifications while increasing the confidence in those predictions. Spacecraft requirements can include a maximum airflow speed to protect delicate instruments during ground processing. Computational Fluid Dynamics can be used to verify these requirements; however, the model must be validated by test data. This research includes the following three objectives and methods. Objective one is develop, model, and perform a Computational Fluid Dynamics analysis of three (3) generic, non-proprietary, environmental control systems and spacecraft configurations. Several commercially available and open source solvers have the capability to model the turbulent, highly three-dimensional, incompressible flow regime. The proposed method uses FLUENT, STARCCM+, and OPENFOAM. Objective two is to perform an uncertainty analysis of the Computational Fluid Dynamics model using the methodology found in "Comprehensive Approach to Verification and Validation of Computational Fluid Dynamics Simulations". This method requires three separate grids and solutions, which quantify the error bars around Computational Fluid Dynamics predictions. The method accounts for all uncertainty terms from both numerical and input variables. Objective three is to compile a table of uncertainty parameters that could be used to estimate the error in a Computational Fluid Dynamics model of the Environmental Control System /spacecraft system. Previous studies have looked at the uncertainty in a Computational Fluid Dynamics model for a single output variable at a single point, for example the re-attachment length of a backward facing step. For the flow regime being analyzed (turbulent, three-dimensional, incompressible), the error at a single point can propagate into the solution both via flow physics and numerical methods. Calculating the uncertainty in using Computational Fluid Dynamics to accurately predict airflow speeds around encapsulated spacecraft in is imperative to the success of future missions.

Dissertation Defense: Computational Fluid Dynamics Uncertainty Analysis for Payload Fairing Spacecraft Environmental Control Systems

NASA Technical Reports Server (NTRS)

Groves, Curtis Edward

2014-01-01

Spacecraft thermal protection systems are at risk of being damaged due to airflow produced from Environmental Control Systems. There are inherent uncertainties and errors associated with using Computational Fluid Dynamics to predict the airflow field around a spacecraft from the Environmental Control System. This paper describes an approach to quantify the uncertainty in using Computational Fluid Dynamics to predict airflow speeds around an encapsulated spacecraft without the use of test data. Quantifying the uncertainty in analytical predictions is imperative to the success of any simulation-based product. The method could provide an alternative to traditional validation by test only mentality. This method could be extended to other disciplines and has potential to provide uncertainty for any numerical simulation, thus lowering the cost of performing these verifications while increasing the confidence in those predictions.Spacecraft requirements can include a maximum airflow speed to protect delicate instruments during ground processing. Computational Fluid Dynamics can be used to verify these requirements; however, the model must be validated by test data. This research includes the following three objectives and methods. Objective one is develop, model, and perform a Computational Fluid Dynamics analysis of three (3) generic, non-proprietary, environmental control systems and spacecraft configurations. Several commercially available and open source solvers have the capability to model the turbulent, highly three-dimensional, incompressible flow regime. The proposed method uses FLUENT, STARCCM+, and OPENFOAM. Objective two is to perform an uncertainty analysis of the Computational Fluid Dynamics model using the methodology found in Comprehensive Approach to Verification and Validation of Computational Fluid Dynamics Simulations. This method requires three separate grids and solutions, which quantify the error bars around Computational Fluid Dynamics predictions. The method accounts for all uncertainty terms from both numerical and input variables. Objective three is to compile a table of uncertainty parameters that could be used to estimate the error in a Computational Fluid Dynamics model of the Environmental Control System spacecraft system.Previous studies have looked at the uncertainty in a Computational Fluid Dynamics model for a single output variable at a single point, for example the re-attachment length of a backward facing step. For the flow regime being analyzed (turbulent, three-dimensional, incompressible), the error at a single point can propagate into the solution both via flow physics and numerical methods. Calculating the uncertainty in using Computational Fluid Dynamics to accurately predict airflow speeds around encapsulated spacecraft in is imperative to the success of future missions.

Computational Fluid Dynamics Uncertainty Analysis for Payload Fairing Spacecraft Environmental Control Systems

NASA Technical Reports Server (NTRS)

Groves, Curtis E.

2013-01-01

Spacecraft thermal protection systems are at risk of being damaged due to airflow produced from Environmental Control Systems. There are inherent uncertainties and errors associated with using Computational Fluid Dynamics to predict the airflow field around a spacecraft from the Environmental Control System. This proposal describes an approach to validate the uncertainty in using Computational Fluid Dynamics to predict airflow speeds around an encapsulated spacecraft. The research described here is absolutely cutting edge. Quantifying the uncertainty in analytical predictions is imperative to the success of any simulation-based product. The method could provide an alternative to traditional"validation by test only'' mentality. This method could be extended to other disciplines and has potential to provide uncertainty for any numerical simulation, thus lowering the cost of performing these verifications while increasing the confidence in those predictions. Spacecraft requirements can include a maximum airflow speed to protect delicate instruments during ground processing. Computationaf Fluid Dynamics can be used to veritY these requirements; however, the model must be validated by test data. The proposed research project includes the following three objectives and methods. Objective one is develop, model, and perform a Computational Fluid Dynamics analysis of three (3) generic, non-proprietary, environmental control systems and spacecraft configurations. Several commercially available solvers have the capability to model the turbulent, highly three-dimensional, incompressible flow regime. The proposed method uses FLUENT and OPEN FOAM. Objective two is to perform an uncertainty analysis of the Computational Fluid . . . Dynamics model using the methodology found in "Comprehensive Approach to Verification and Validation of Computational Fluid Dynamics Simulations". This method requires three separate grids and solutions, which quantify the error bars around Computational Fluid Dynamics predictions. The method accounts for all uncertainty terms from both numerical and input variables. Objective three is to compile a table of uncertainty parameters that could be used to estimate the error in a Computational Fluid Dynamics model of the Environmental Control System /spacecraft system. Previous studies have looked at the uncertainty in a Computational Fluid Dynamics model for a single output variable at a single point, for example the re-attachment length of a backward facing step. To date, the author is the only person to look at the uncertainty in the entire computational domain. For the flow regime being analyzed (turbulent, threedimensional, incompressible), the error at a single point can propagate into the solution both via flow physics and numerical methods. Calculating the uncertainty in using Computational Fluid Dynamics to accurately predict airflow speeds around encapsulated spacecraft in is imperative to the success of future missions.

Evolutionary Scheduler for the Deep Space Network

NASA Technical Reports Server (NTRS)

Guillaume, Alexandre; Lee, Seungwon; Wang, Yeou-Fang; Zheng, Hua; Chau, Savio; Tung, Yu-Wen; Terrile, Richard J.; Hovden, Robert

2010-01-01

A computer program assists human schedulers in satisfying, to the maximum extent possible, competing demands from multiple spacecraft missions for utilization of the transmitting/receiving Earth stations of NASA s Deep Space Network. The program embodies a concept of optimal scheduling to attain multiple objectives in the presence of multiple constraints.

International Space Station (ISS) Meteoroid/Orbital Debris Shielding

NASA Technical Reports Server (NTRS)

Christiansen, Eric L.

1999-01-01

Design practices to provide protection for International Space Station (ISS) crew and critical equipment from meteoroid and orbital debris (M/OD) Impacts have been developed. Damage modes and failure criteria are defined for each spacecraft system. Hypervolocity Impact -1 - and analyses are used to develop ballistic limit equations (BLEs) for each exposed spacecraft system. BLEs define Impact particle sizes that result in threshold failure of a particular spacecraft system as a function of Impact velocity, angles and particle density. The BUMPER computer code Is used to determine the probability of no penetration (PNP) that falls the spacecraft shielding based on NASA standard meteoroid/debris models, a spacecraft geometry model, and the BLEs. BUMPER results are used to verify spacecraft shielding requirements Low-weight, high-performance shielding alternatives have been developed at the NASA Johnson Space Center (JSC) Hypervelocity Impact Technology Facility (HITF) to meet spacecraft protection requirements.

Electrical Grounding Architecture for Unmanned Spacecraft

NASA Technical Reports Server (NTRS)

1998-01-01

This handbook is approved for use by NASA Headquarters and all NASA Centers and is intended to provide a common framework for consistent practices across NASA programs. This handbook was developed to describe electrical grounding design architecture options for unmanned spacecraft. This handbook is written for spacecraft system engineers, power engineers, and electromagnetic compatibility (EMC) engineers. Spacecraft grounding architecture is a system-level decision which must be established at the earliest point in spacecraft design. All other grounding design must be coordinated with and be consistent with the system-level architecture. This handbook assumes that there is no one single 'correct' design for spacecraft grounding architecture. There have been many successful satellite and spacecraft programs from NASA, using a variety of grounding architectures with different levels of complexity. However, some design principles learned over the years apply to all types of spacecraft development. This handbook summarizes those principles to help guide spacecraft grounding architecture design for NASA and others.

Saturn Apollo Program

NASA Image and Video Library

1974-01-01

This illustration depicts a comparison of two space vehicles, the U.S.'s Saturn IB launch vehicle and the U.S.S.R.'s Soyuz launch vehicle, for the Apollo-Soyuz Test Project. The ASTP was the first international docking of the U.S.'s Apollo spacecraft and the U.S.S.R.'s Soyuz spacecraft in space. A joint engineering team from the two countries met to develop a docking system that permitted the two spacecraft to link in space and allowed the two crews to travel from one spacecraft to the other. This system entailed developing a large habitable Docking Module (DM) to be carried on the Apollo spacecraft to facilitate the joining of two dissimilar spacecraft. The Marshall Space Flight Center was responsible for development and sustaining engineering of the Saturn IB launch vehicle during the mission.

Rapid Spacecraft Development: Results and Lessons Learned

NASA Technical Reports Server (NTRS)

Watson, William A.

2002-01-01

The Rapid Spacecraft Development Office (RSDO) at NASA's Goddard Space Flight Center is responsible for the management and direction of a dynamic and versatile program for the definition, competition, and acquisition of multiple indefinite delivery and indefinite quantity contracts - resulting in a catalog of spacecraft buses. Five spacecraft delivery orders have been placed by the RSDO and one spacecraft has been launched. Numerous concept and design studies have been performed, most with the intent of leading to a future spacecraft acquisition. A collection of results and lessons learned is recorded to highlight management techniques, methods and processes employed in the conduct of spacecraft acquisition. Topics include working relationships under fixed price delivery orders, price and value, risk management, contingency reserves, and information restrictions.

Development of electrical test procedures for qualification of spacecraft against EID. Volume 2: Review and specification of test procedures

NASA Technical Reports Server (NTRS)

Wilkenfeld, J. M.; Harlacher, B. L.; Mathews, D.

1982-01-01

A combined experimental and analytical program to develop system electrical test procedures for the qualification of spacecraft against damage produced by space-electron-induced discharges (EID) occurring on spacecraft dielectric outer surfaces is described. A review and critical evaluation of possible approaches to qualify spacecraft against space electron-induced discharges (EID) is presented. A variety of possible schemes to simulate EID electromagnetic effects produced in spacecraft was studied. These techniques form the principal element of a provisional, recommended set of test procedures for the EID qualification spacecraft. Significant gaps in our knowledge about EID which impact the final specification of an electrical test to qualify spacecraft against EID are also identified.

Cooper-Harper Experience Report for Spacecraft Handling Qualities Applications

NASA Technical Reports Server (NTRS)

Bailey, Randall E.; Jackson, E. Bruce; Bilimoria, Karl D.; Mueller, Eric R.; Frost, Chad R.; Alderete, Thomas S.

2009-01-01

A synopsis of experience from the fixed-wing and rotary-wing aircraft communities in handling qualities development and the use of the Cooper-Harper pilot rating scale is presented as background for spacecraft handling qualities research, development, test, and evaluation (RDT&E). In addition, handling qualities experiences and lessons-learned from previous United States (US) spacecraft developments are reviewed. This report is intended to provide a central location for references, best practices, and lessons-learned to guide current and future spacecraft handling qualities RDT&E.

Study of hollow corner retroreflectors for use in a synchronous orbit

NASA Technical Reports Server (NTRS)

Yoder, P. R., Jr.

1975-01-01

The performance of a hollow corner cube retroreflector made up of three mutually perpendicular optically flat mirrors when undergoing the thermal-mechanical strains induced by a spacecraft environment was studied. Of particular interest was a device of 200 square centimeter optical aperture used on a satellite in a synchronous orbit. It was assumed that the reflector always faces the earth. The effects of direct solar irradiance, earthshine, and albedo were considered. The results included the maximum mirror surface temperature during the orbit as well as the worst-case loss of optical performance due to thermally-induced mirror distortions. It was concluded that a device made of three suitably coated flat ULE mirrors, optically contacted to each other and supported mechanically in a nonrigid mount, would be expected to concentrate over 80 percent of the theoretical maximum energy in the central of the far field diffraction pattern. Continued development of the device through a detailed design, fabrication, and test phase was recommended.

The Voyager spacecraft /James Watt International Gold Medal Lecture/

NASA Technical Reports Server (NTRS)

Heacock, R. L.

1980-01-01

The Voyager Project background is reviewed with emphasis on selected features of the Voyager spacecraft. Investigations by the Thermo-electric Outer Planets Spacecraft Project are discussed, including trajectories, design requirements, and the development of a Self Test and Repair computer, and a Computer Accessed Telemetry System. The design and configuration of the spacecraft are described, including long range communications, attitude control, solar independent power, sequencing and control data handling, and spacecraft propulsion. The development program, maintained by JPL, experienced a variety of problems such as design deficiencies, and process control and manufacturing problems. Finally, the spacecraft encounter with Jupiter is discussed, and expectations for the Saturn encounter are expressed.

Vibration Testing of Stirling Power Convertors

NASA Technical Reports Server (NTRS)

Hughes, Bill; Goodnight, Thomas; McNelis, Mark E.; Suarez, Vicente J.; Schreiber, Jeff; Samorezov, Sergey

2003-01-01

The NASA John H. Glenn Research Center (GRC) and the U.S. Department of Energy (DOE) are currently developing a high efficient, long life, free piston Stirling convertor for use as an advanced spacecraft power system for future NASA missions. As part of this development, a Stirling Technology Demonstrator Convertor (TDC), developed by Stirling Technology Company (STC) for DOE, was vibration tested at GRC s Structural Dynamics Laboratory (SDU7735) in November- December 1999. This testing demonstrated that the Stirling TDC is able to withstand the harsh random vibration (20 to 2000 Hertz) seen during a typical spacecraft launch and survive with no structural damage or functional power performance degradation, thereby enabling its usage in future spacecraft power systems. The Stirling Vibration Test Team at NASA GRC and STC personnel conducted tests on a single 55 electric watt TDC. The purpose was to characterize the TDC s structural response to vibration and determine if the TDC could survive the vibration criteria established by the Jet Propulsion Laboratory (JPL) for launch environments. The TDC was operated at full-stroke and full power conditions during the vibration testing. The TDC was tested in two orientations, with the direction of vibration parallel and perpendicular to the TDC s moving components (displacer and piston). The TDC successfully passed a series of sine and random vibration tests. The most severe test was a 12.3 Grms random vibration test (peak vibration level of 0.2 g2/Hz from 50 to 250 Hertz) with test durations of 3 minutes per axis. The random vibration test levels were chosen to simulate, with margin, the maximum anticipated launch vibration conditions. As a result of this very successful vibration testing and successful evaluations in other key technical readiness areas, the Stirling power system is now considered a viable technology for future application for NASA spacecraft missions. Possible usage of the Stirling power system would be to supply on- board electric spacecraft power for future NASA Deep-Space Missions, performing as an attractive alternative to Radioisotope Thermoelectric Generators (RTG). Usage of the Stirling technology is also being considered as the electric power source for future Mars rovers, whose mission profiles may exclude the use of photovoltaic power systems (such as exploring at high Martian latitudes or for missions of lengthy durations). GRC s Thermo-Mechanical Systems Branch (5490) provides Stirling technology expertise under a Space Act Agreement with the DOE. Additional vibration testing, by GRC s Structural Systems Dynamics Branch (7733, is planned to continue to demonstrate the Stirling power system s vibration capability as its technology and flight system designs progress.

Test Telemetry And Command System (TTACS)

NASA Technical Reports Server (NTRS)

Fogel, Alvin J.

1994-01-01

The Jet Propulsion Laboratory has developed a multimission Test Telemetry and Command System (TTACS) which provides a multimission telemetry and command data system in a spacecraft test environment. TTACS reuses, in the spacecraft test environment, components of the same data system used for flight operations; no new software is developed for the spacecraft test environment. Additionally, the TTACS is transportable to any spacecraft test site, including the launch site. The TTACS is currently operational in the Galileo spacecraft testbed; it is also being provided to support the Cassini and Mars Surveyor Program projects. Minimal personnel data system training is required in the transition from pre-launch spacecraft test to post-launch flight operations since test personnel are already familiar with the data system's operation. Additionally, data system components, e.g. data display, can be reused to support spacecraft software development; and the same data system components will again be reused during the spacecraft integration and system test phases. TTACS usage also results in early availability of spacecraft data to data system development and, as a result, early data system development feedback to spacecraft system developers. The TTACS consists of a multimission spacecraft support equipment interface and components of the multimission telemetry and command software adapted for a specific project. The TTACS interfaces to the spacecraft, e.g., Command Data System (CDS), support equipment. The TTACS telemetry interface to the CDS support equipment performs serial (RS-422)-to-ethernet conversion at rates between 1 bps and 1 mbps, telemetry data blocking and header generation, guaranteed data transmission to the telemetry data system, and graphical downlink routing summary and control. The TTACS command interface to the CDS support equipment is nominally a command file transferred in non-real-time via ethernet. The CDS support equipment is responsible for metering the commands to the CDS; additionally for Galileo, TTACS includes a real-time-interface to the CDS support equipment. The TTACS provides the basic functionality of the multimission telemetry and command data system used during flight operations. TTACS telemetry capabilities include frame synchronization, Reed-Solomon decoding, packet extraction and channelization, and data storage/query. Multimission data display capabilities are also available. TTACS command capabilities include command generation verification, and storage.

The other side of the safety coin. [aerospace operations

NASA Technical Reports Server (NTRS)

Roth, Gilbert L.

1986-01-01

The development, inspection and testing requirements for successful production and launch and safe operation of spaceflight hardware are discussed. Emphasis is placed on paying acute attention to malfunctions, which could be caused by contaminants (particles in docking rings), insufficiently durable materials (Orbiter brakes), etc. Generic and specific problems which occur in propulsion, avionics, mechanical and computer systems and in configuration management, manufacturing and process control efforts are explored. Case histories of deficiencies found in LOX fuel lines, contaminated hydraulic control systems, the Solar Maximum Mission thermal insulation grommets, are summarized. Thorough inspection and testing procedures and design change recording during manufacture of spacecraft components are identified as requisites for successful space missions.

Imaging Flash Lidar for Safe Landing on Solar System Bodies and Spacecraft Rendezvous and Docking

NASA Technical Reports Server (NTRS)

Amzajerdian, Farzin; Roback, Vincent E.; Bulyshev, Alexander E.; Brewster, Paul F.; Carrion, William A; Pierrottet, Diego F.; Hines, Glenn D.; Petway, Larry B.; Barnes, Bruce W.; Noe, Anna M.

2015-01-01

NASA has been pursuing flash lidar technology for autonomous, safe landing on solar system bodies and for automated rendezvous and docking. During the final stages of the landing from about 1 kilometer to 500 meters above the ground, the flash lidar can generate 3-Dimensional images of the terrain to identify hazardous features such as craters, rocks, and steep slopes. The onboard flight computer can then use the 3-D map of terrain to guide the vehicle to a safe location. As an automated rendezvous and docking sensor, the flash lidar can provide relative range, velocity, and bearing from an approaching spacecraft to another spacecraft or a space station. NASA Langley Research Center has developed and demonstrated a flash lidar sensor system capable of generating 16,000 pixels range images with 7 centimeters precision, at 20 Hertz frame rate, from a maximum slant range of 1800 m from the target area. This paper describes the lidar instrument and presents the results of recent flight tests onboard a rocket-propelled free-flyer vehicle (Morpheus) built by NASA Johnson Space Center. The flights were conducted at a simulated lunar terrain site, consisting of realistic hazard features and designated landing areas, built at NASA Kennedy Space Center specifically for this demonstration test. This paper also provides an overview of the plan for continued advancement of the flash lidar technology aimed at enhancing its performance to meet both landing and automated rendezvous and docking applications.

NASA Centers and Universities Collaborate Through Smallsat Technology Partnerships

NASA Technical Reports Server (NTRS)

Cockrell, James

2018-01-01

The Small Spacecraft Technology (SST) Program within the NASA Space Technology Mission Directorate is chartered develop and demonstrate the capabilities that enable small spacecraft to achieve science and exploration missions in "unique" and "more affordable" ways. Specifically, the SST program seeks to enable new mission architectures through the use of small spacecraft, to expand the reach of small spacecraft to new destinations, and to make possible the augmentation existing assets and future missions with supporting small spacecraft. The SST program sponsors smallsat technology development partnerships between universities and NASA Centers in order to engage the unique talents and fresh perspectives of the university community and to share NASA experience and expertise in relevant university projects to develop new technologies and capabilities for small spacecraft. These partnerships also engage NASA personnel in the rapid, agile and cost-conscious small spacecraft approaches that have evolved in the university community, as well as increase support to university efforts and foster a new generation of innovators for NASA and the nation.

High Power Microwave Tube Reliability Study

DTIC Science & Technology

1976-08-01

Factors . . . . . ................ 67 1. Environmental Factors . . . . . . . . . a. Ground Fixed ...... .......... 67 b. Ground Mobile ...including cube structure and operating parameters as factors in the models but also environment and aplication . Initially, the tubes to be included in...instLllations. Mobile ground based and seagoing systems have minimum restrictions, spacecraft systems the maximum and airborne system Y 6 .*.. restrictionts

Sonic-boom measurements in the focus region during the ascent of Apollo 17. [maximum positive overpressure, positive impulse, signature duration, and bow-shock rise time

NASA Technical Reports Server (NTRS)

Henderson, H. R.; Hilton, D. A.

1974-01-01

Sonic-boom pressure signatures recorded during the ascent phase of Apollo 17 are presented. The measurements were obtained onboard six U.S. Navy ships positioned along the ground track of the spacecraft vehicle in the area of expected focus resulting from the flight path and acceleration of the vehicle. Tracings of the measured signatures are presented along with values of the maximum positive overpressure, positive impulse, signature duration, and bowshock rise time. Also included are brief descriptions of the ships and their location, the deployment of the sonic-boom instrumentation, flight profiles and operating conditions for the launch vehicle and spacecraft, surface-weather and sea-state information at the measuring sites, and high-altitude weather information for the general measurement areas. Comparisons of the measured and predicted sonic-boom overpressures for the Apollo 17 mission are presented. The measured data are also compared with data from the Apollo 15 and 16 missions and data from flight test programs of various aircraft.

Spacecraft attitude and velocity control system

NASA Technical Reports Server (NTRS)

Paluszek, Michael A. (Inventor); Piper, Jr., George E. (Inventor)

1992-01-01

A spacecraft attitude and/or velocity control system includes a controller which responds to at least attitude errors to produce command signals representing a force vector F and a torque vector T, each having three orthogonal components, which represent the forces and torques which are to be generated by the thrusters. The thrusters may include magnetic torquer or reaction wheels. Six difference equations are generated, three having the form ##EQU1## where a.sub.j is the maximum torque which the j.sup.th thruster can produce, b.sub.j is the maximum force which the j.sup.th thruster can produce, and .alpha..sub.j is a variable representing the throttling factor of the j.sup.th thruster, which may range from zero to unity. The six equations are summed to produce a single scalar equation relating variables .alpha..sub.j to a performance index Z: ##EQU2## Those values of .alpha. which maximize the value of Z are determined by a method for solving linear equations, such as a linear programming method. The Simplex method may be used. The values of .alpha..sub.j are applied to control the corresponding thrusters.

Cluster Inter-Spacecraft Communications

NASA Technical Reports Server (NTRS)

Cox, Brian

2008-01-01

A document describes a radio communication system being developed for exchanging data and sharing data-processing capabilities among spacecraft flying in formation. The system would establish a high-speed, low-latency, deterministic loop communication path connecting all the spacecraft in a cluster. The system would be a wireless version of a ring bus that complies with the Institute of Electrical and Electronics Engineers (IEEE) standard 1393 (which pertains to a spaceborne fiber-optic data bus enhancement to the IEEE standard developed at NASA's Jet Propulsion Laboratory). Every spacecraft in the cluster would be equipped with a ring-bus radio transceiver. The identity of a spacecraft would be established upon connection into the ring bus, and the spacecraft could be at any location in the ring communication sequence. In the event of failure of a spacecraft, the ring bus would reconfigure itself, bypassing a failed spacecraft. Similarly, the ring bus would reconfigure itself to accommodate a spacecraft newly added to the cluster or newly enabled or re-enabled. Thus, the ring bus would be scalable and robust. Reliability could be increased by launching, into the cluster, spare spacecraft to be activated in the event of failure of other spacecraft.

Astronaut Alan Shepard - U.S.S. Champlain - Post-Recovery Mercury Capsule

NASA Image and Video Library

1961-05-05

S61-02727 (5 May 1961) --- Astronaut Alan B. Shepard is seen on the deck of the USS Lake Champlain after the recovery of his Mercury capsule in the western Atlantic Ocean. Shepard and the Mercury spacecraft designated the ?Freedom 7? were flown to the deck of the recovery ship within 11 minutes of splashdown. MR-3 was the United States? first manned space mission. The spacecraft attained a maximum speed of 5,180 miles per hour, reached an altitude of 116 1/2 statute miles, and landed 302 statute miles downrange from Cape Canaveral, Florida. The suborbital mission lasted 15 minutes and 22 seconds. Photo credit: NASA or National Aeronautics and Space Administration

Solar array study for solar electric propulsion spacecraft for the Encke rendezvous mission

NASA Technical Reports Server (NTRS)

Sequeira, E. A.; Patterson, R. E.

1974-01-01

The work is described which was performed on the design, analysis and performance of a 20 kW rollup solar array capable of meeting the design requirements of a solar electric spacecraft for the 1980 Encke rendezvous mission. To meet the high power requirements of the proposed electric propulsion mission, solar arrays on the order of 186.6 sq m were defined. Because of the large weights involved with arrays of this size, consideration of array configurations is limited to lightweight, large area concepts with maximum power-to-weight ratios. Items covered include solar array requirements and constraints, array concept selection and rationale, structural and electrical design considerations, and reliability considerations.

Study and development of a cryogenic heat exchanger for life support systems

NASA Technical Reports Server (NTRS)

Soliman, M. M.

1973-01-01

A prototype cryogenic heat exchanger for removal of waste heat from a spacecraft environmental control life support system was developed. The heat exchanger uses the heat sink capabilities of the cryogenic propellants and, hence, can operate over all mission phases from prelaunch to orbit, to post landing, with quiescent periods during orbit. A survey of candidate warm fluids resulted in the selection of E-2, a fluorocarbon compound, because of its low freezing point and high boiling point. The final design and testing of the heat exchanger was carried out, however, using Freon-21, which is similar to E-2 except for its low boiling point. This change was motivated by the desire for cost effectiveness of the experimental program. The transient performance of the heat exchanger was demonstrated by an analog simulation of the heat sink system. Under the realistic transient heat load conditions (20 sec ramp from minimum to maximum Freon-21 inlet temperature), the control system was able to maintain the warm fluid outlet temperature within + or - 3 F. For a 20-sec ramp from 0 F to -400 F in the hydrogen inlet temperature, at maximum heat load, the warm fluid outlet temperature was maintained within + or - 7 F.

Heliocentric phasing performance of electric sail spacecraft

NASA Astrophysics Data System (ADS)

Mengali, Giovanni; Quarta, Alessandro A.; Aliasi, Generoso

2016-10-01

We investigate the heliocentric in-orbit repositioning problem of a spacecraft propelled by an Electric Solar Wind Sail. Given an initial circular parking orbit, we look for the heliocentric trajectory that minimizes the time required for the spacecraft to change its azimuthal position, along the initial orbit, of a (prescribed) phasing angle. The in-orbit repositioning problem can be solved using either a drift ahead or a drift behind maneuver and, in general, the flight times for the two cases are different for a given value of the phasing angle. However, there exists a critical azimuthal position, whose value is numerically found, which univocally establishes whether a drift ahead or behind trajectory is superior in terms of flight time it requires for the maneuver to be completed. We solve the optimization problem using an indirect approach for different values of both the spacecraft maximum propulsive acceleration and the phasing angle, and the solution is then specialized to a repositioning problem along the Earth's heliocentric orbit. Finally, we use the simulation results to obtain a first order estimate of the minimum flight times for a scientific mission towards triangular Lagrangian points of the Sun-[Earth+Moon] system.

Trajectory Control of Rendezvous with Maneuver Target Spacecraft

NASA Technical Reports Server (NTRS)

Zhou, Zhinqiang

2012-01-01

In this paper, a nonlinear trajectory control algorithm of rendezvous with maneuvering target spacecraft is presented. The disturbance forces on the chaser and target spacecraft and the thrust forces on the chaser spacecraft are considered in the analysis. The control algorithm developed in this paper uses the relative distance and relative velocity between the target and chaser spacecraft as the inputs. A general formula of reference relative trajectory of the chaser spacecraft to the target spacecraft is developed and applied to four different proximity maneuvers, which are in-track circling, cross-track circling, in-track spiral rendezvous and cross-track spiral rendezvous. The closed-loop differential equations of the proximity relative motion with the control algorithm are derived. It is proven in the paper that the tracking errors between the commanded relative trajectory and the actual relative trajectory are bounded within a constant region determined by the control gains. The prediction of the tracking errors is obtained. Design examples are provided to show the implementation of the control algorithm. The simulation results show that the actual relative trajectory tracks the commanded relative trajectory tightly. The predicted tracking errors match those calculated in the simulation results. The control algorithm developed in this paper can also be applied to interception of maneuver target spacecraft and relative trajectory control of spacecraft formation flying.

Thermal Analysis of Iodine Satellite (iSAT)

NASA Technical Reports Server (NTRS)

Mauro, Stephanie

2015-01-01

This paper presents the progress of the thermal analysis and design of the Iodine Satellite (iSAT). The purpose of the iSAT spacecraft (SC) is to demonstrate the ability of the iodine Hall Thruster propulsion system throughout a one year mission in an effort to mature the system for use on future satellites. The benefit of this propulsion system is that it uses a propellant, iodine, that is easy to store and provides a high thrust-to-mass ratio. The spacecraft will also act as a bus for an earth observation payload, the Long Wave Infrared (LWIR) Camera. Four phases of the mission, determined to either be critical to achieving requirements or phases of thermal concern, are modeled. The phases are the Right Ascension of the Ascending Node (RAAN) Change, Altitude Reduction, De-Orbit, and Science Phases. Each phase was modeled in a worst case hot environment and the coldest phase, the Science Phase, was also modeled in a worst case cold environment. The thermal environments of the spacecraft are especially important to model because iSAT has a very high power density. The satellite is the size of a 12 unit cubesat, and dissipates slightly more than 75 Watts of power as heat at times. The maximum temperatures for several components are above their maximum operational limit for one or more cases. The analysis done for the first Design and Analysis Cycle (DAC1) showed that many components were above or within 5 degrees Centigrade of their maximum operation limit. The battery is a component of concern because although it is not over its operational temperature limit, efficiency greatly decreases if it operates at the currently predicted temperatures. In the second Design and Analysis Cycle (DAC2), many steps were taken to mitigate the overheating of components, including isolating several high temperature components, removal of components, and rearrangement of systems. These changes have greatly increased the thermal margin available.

STATISTICAL CHARACTERISTICS OF ELEMENTAL ABUNDANCE RATIOS: OBSERVATIONS FROM THE ACE SPACECRAFT

DOE Office of Scientific and Technical Information (OSTI.GOV)

Zhao, L.-L.; Zhang, H.

We statistically analyze the elemental galactic cosmic ray (GCR) composition measurements of elements 5 ≤ Z ≤ 28 within the energy range 30–500 MeV/nucleon from the CRIS instrument on board the ACE spacecraft in orbit about the L1 Lagrange point during the period from 1997 to 2014. Similarly to the last unusual solar minimum, the elevated elemental intensities of all heavy nuclei during the current weak solar maximum in 2014 are ∼40% higher than that of the previous solar maximum in 2002, which has been attributed to the weak modulation associated with low solar activity levels during the ongoing weakestmore » solar maximum since the dawn of space age. In addition, the abundance ratios of heavy nuclei with respect to elemental oxygen are generally independent of kinetic energy per nucleon in the energy region 60–200 MeV/nuc, in good agreement with previous experiments. Furthermore, the abundance ratios of most relatively abundant species, except carbon, exhibit considerable solar-cycle variation, which are obviously positively correlated with the sunspot numbers with about one-year time lag. We also find that the percentage variation of abundance ratios for most elements are approximately identical. These preliminary results provide valuable insights into the characteristics of elemental heavy nuclei composition and place new and significant constraints on future GCR heavy nuclei propagation and modulation models.« less

Spacecraft 2000

NASA Technical Reports Server (NTRS)

1986-01-01

The objective of the Workshop was to focus on the key technology area for 21st century spacecraft and the programs needed to facilitate technology development and validation. Topics addressed include: spacecraft systems; system development; structures and materials; thermal control; electrical power; telemetry, tracking, and control; data management; propulsion; and attitude control.

The Apollo spacecraft: A chronology. Volume 2: 8 November 1962 - 30 September 1964

NASA Technical Reports Server (NTRS)

Morse, M. L.; Bays, J. K.

1973-01-01

A chronology of the Apollo spacecraft development and production program is presented. The subjects discussed are: (1) defining contractural relations, (2) developing hardware distinctions, and (3) developing software ground rules. Illustrations, drawings, and photographs are used extensively to supplement the technical writing. Descriptions of life support systems, communication equipment, propulsion systems, control devices, and spacecraft components are provided.

The natural space environment: Effects on spacecraft

NASA Technical Reports Server (NTRS)

James, Bonnie F.; Norton, O. W. (Compiler); Alexander, Margaret B. (Editor)

1994-01-01

The effects of the natural space environments on spacecraft design, development, and operation are the topic of a series of NASA Reference Publications currently being developed by the Electromagnetics and Environments Branch, Systems Analysis and Integration Laboratory, Marshall Space Flight Center. This primer provides an overview of the natural space environments and their effect on spacecraft design, development, and operations, and also highlights some of the new developments in science and technology for each space environment. It is hoped that a better understanding of the space environment and its effect on spacecraft will enable program management to more effectively minimize program risks and costs, optimize design quality, and successfully achieve mission objectives.

Structures and materials technology needs for communications and remote sensing spacecraft

NASA Technical Reports Server (NTRS)

Gronet, M. J.; Jensen, G. A.; Hoskins, J. W.

1995-01-01

This report documents trade studies conducted from the perspective of a small spacecraft developer to determine and quantify the structures and structural materials technology development needs for future commercial and NASA small spacecraft to be launched in the period 1999 to 2005. Emphasis is placed on small satellites weighing less than 1800 pounds for two focus low-Earth orbit missions: commercial communications and remote sensing. The focus missions are characterized in terms of orbit, spacecraft size, performance, and design drivers. Small spacecraft program personnel were interviewed to determine their technology needs, and the results are summarized. A systems-analysis approach for quantifying the benefits of inserting advanced state-of-the-art technologies into a current reference, state-of-the-practice small spacecraft design is developed and presented. This approach is employed in a set of abbreviated trade studies to quantify the payoffs of using a subset of 11 advanced technologies selected from the interview results The 11 technology development opportunities are then ranked based on their relative payoff. Based on the strong potential for significant benefits, recommendations are made to pursue development of 8 and the 11 technologies. Other important technology development areas identified are recommended for further study.

A new environment for multiple spacecraft power subsystem mission operations

NASA Technical Reports Server (NTRS)

Bahrami, K. A.

1990-01-01

The engineering analysis subsystem environment (EASE) is being developed to enable fewer controllers to monitor and control power and other spacecraft engineering subsystems. The EASE prototype has been developed to support simultaneous real-time monitoring of several spacecraft engineering subsystems. It is being designed to assist with offline analysis of telemetry data to determine trends, and to help formulate uplink commands to the spacecraft. An early version of the EASE prototype has been installed in the JPL Space Flight Operations Facility for online testing. The EASE prototype is installed in the Galileo Mission Support Area. The underlying concept, development, and testing of the EASE prototype and how it will aid in the ground operations of spacecraft power subsystems are discussed.

EDOS operations concept and development approach

NASA Technical Reports Server (NTRS)

Knoble, G.; Garman, C.; Alcott, G.; Ramchandani, C.; Silvers, J.

1994-01-01

The Earth Observing System (EOS) Data and Operations System (EDOS) is being developed by the National Aeronautics and Space Administration (NASA) Goddard Space Flight Center (GSFC) for the capture, level zero processing, distribution, and backup archiving of high speed telemetry data received from EOS spacecraft. All data received will conform to the Consultative Committee for Space Data Standards (CCSDS) recommendations. The major EDOS goals are to: (1) minimize EOS program costs to implement and operate EDOS; (2) respond effectively to EOS growth requirements; and (3) maintain compatibility with existing and enhanced versions of NASA institutional systems required to support EOS spacecraft. In order to meet these goals, the following objectives have been defined for EDOS: (1) standardize EDOS interfaces to maximize utility for future requirements; (2) emphasize life-cycle cost (LCC) considerations (rather than procurement costs) in making design decisions and meeting reliability, maintainability, availability (RMA) and upgradability requirements; (3) implement data-driven operations to the maximum extent possible to minimize staffing requirements and to maximize system responsiveness; (4) provide a system capable of simultaneously supporting multiple spacecraft, each in different phases of their life-cycles; (5) provide for technology insertion features to accommodate growth and future LCC reductions during the operations phase; and (6) provide a system that is sufficiently robust to accommodate incremental performance upgrades while supporting operations. Operations concept working group meetings were facilitated to help develop the EDOS operations concept. This provided a cohesive concept that met with approval of responsible personnel from the start. This approach not only speeded up the development process by reducing review cycles, it also provided a medium for generating good ideas that were immediately molded into feasible concepts. The operations concept was then used as a basis for the EDOS specification. When it was felt that concept elements did not support detailed requirements, the facilitator process was used to resolve discrepancies or to add new concept elements to support the specification. This method provided an ongoing revisal of the operations concept and prevented large revisions at the end of the requirement analysis phase of system development.

Tropical Rainfall Measuring Mission (TRMM) project. VI - Spacecraft, scientific instruments, and launching rocket. Part 1 - Spacecraft

NASA Technical Reports Server (NTRS)

Keating, Thomas; Ihara, Toshio; Miida, Sumio

1990-01-01

A cooperative United States/Japan study was made for one year from 1987 to 1988 regarding the feasibility of the Tropical Rainfall Measuring Mission (TRMM). As part of this study a phase-A-level design of spacecraft for TRMM was developed by NASA/GSFC, and the result was documented in a feasibility study. The phase-A-level design is developed for the TRMM satellite utilizing a multimission spacecraft.

Observations of a Unique Type of ULF Wave by Low-Altitude Space Technology 5 Satellites

NASA Technical Reports Server (NTRS)

Le, G.; Chi, P. J.; Strangeway, R. J.; Slavin, J. A.

2011-01-01

We report a unique type of ULF waves observed by low-altitude Space Technology 5 (ST-5) constellation mission. ST-5 is a three-microsatellite constellation deployed into a 300 x 4500 km dawn-dusk and Sun-synchronous polar orbit with 105.6deg inclination angle. Because of the Earth's rotation and the dipole tilt effect, the spacecraft's dawn-dusk orbit track can reach as low as subauroral latitudes during the course of a day. Whenever the spacecraft traverse the dayside closed field line region at subauroral latitudes, they frequently observe strong transverse oscillations at 30-200 mHz, or in the Pc2-3 frequency range. These Pc2-3 waves appear as wave packets with durations in the order of 5-10 min. As the maximum separations of the ST-5 spacecraft are in the order of 10 min, the three ST-5 satellites often observe very similar wave packets, implying these wave oscillations occur in a localized region. The coordinated ground-based magnetic observations at the spacecraft footprints, however, do not see waves in the Pc2-3 band; instead, the waves appear to be the common Pc4-5 waves associated with field line resonances. We suggest that these unique Pc2-3 waves seen by ST-5 are in fact the Doppler-shifted Pc4-5 waves as a result of rapid traverse of the spacecraft across the resonant field lines azimuthally at low altitudes. The observations with the unique spacecraft dawn-dusk orbits at proper altitudes and magnetic latitudes reveal the azimuthal characteristics of field line resonances.

Attitude ground support system for the solar maximum mission spacecraft

NASA Technical Reports Server (NTRS)

Nair, G.

1980-01-01

The SMM attitude ground support system (AGSS) supports the acquisition of spacecraft roll attitude reference, performs the in-flight calibration of the attitude sensor complement, supports onboard control autonomy via onboard computer data base updates, and monitors onboard computer (OBC) performance. Initial roll attitude acquisition is accomplished by obtaining a coarse 3 axis attitude estimate from magnetometer and Sun sensor data and subsequently refining it by processing data from the fixed head star trackers. In-flight calibration of the attitude sensor complement is achieved by processing data from a series of slew maneuvers designed to maximize the observability and accuracy of the appropriate alignments and biases. To ensure autonomy of spacecraft operation, the AGSS selects guide stars and computes sensor occultation information for uplink to the OBC. The onboard attitude control performance is monitored on the ground through periodic attitude determination and processing of OBC data in downlink telemetry. In general, the control performance has met mission requirements. However, software and hardware problems have resulted in sporadic attitude reference losses.

Optimum solar electric interplanetary mission opportunities from 1975 to 1990

NASA Technical Reports Server (NTRS)

Mann, F. I.; Horsewood, J. L.

1971-01-01

A collection of optimum trajectory and spacecraft data is presented for unmanned interplanetary missions from 1975 to 1990 using solar electric propulsion. Data are presented for one-way flyby and orbiter missions from Earth to Venus, Mars, Jupiter, Saturn, Uranus, Neptune, and Pluto. The solar system model assumes planetary ephemerides which very closely approximate the true motion of the planets. Direct and indirect flight profiles are investigated. Data are presented for two representative flight times for each mission. The launch vehicle is the Titan 3 B (core)/Centaur, and a constant jet exhaust speed solar electric propulsion system having a specific mass of 30 kg/kw is completely optimized in terms of power level and jet exhaust speed to yield maximum net spacecraft mass. The hyperbolic excess speeds at departure and arrival and the launch date are optimized for each mission. For orbiter missions, a chemical retro stage is used to brake the spacecraft into a highly eccentric capture orbit about the target planet.

Radio range measurements of coronal electron densities at 13 and 3.6 centimeter wavelengths during the 1988 solar conjunction of Voyager 2

NASA Technical Reports Server (NTRS)

Krisher, T. P.; Anderson, J. D.; Morabito, D. D.; Asmar, S. W.; Borutzki, S. E.; Delitsky, M. L.; Densmore, A. C.; Eshe, P. M.; Lewis, G. D.; Maurer, M. J.

1991-01-01

Radio range measurements of total solar plasma delay obtained during the solar conjunction of the Voyager 2 spacecraft in December 1988, which occurred near solar maximum activity in the 11 yr cycle are reported. The radio range measurements were generated by the Deep Space Network at two wavelengths on the downlink from the spacecraft: 3.6 and 13 cm. A direct measurement of the integrated electron density along the ray path between the earth stations and the spacecraft was obtained by differencing the range at the two wavelengths. Coronal electron density profiles have been derived during ingress and egress of the ray path, which approached the sun to within 5 solar radii. At 10 solar radii, the derived density profiles yield 34079 + or - 611/cu cm on ingress and 49688 + or - 983/cu cm on egress. These density levels are significantly higher than observed near previous solar maxima.

Maximum Torque and Momentum Envelopes for Reaction Wheel Arrays

NASA Technical Reports Server (NTRS)

Reynolds, R. G.; Markley, F. Landis

2001-01-01

Spacecraft reaction wheel maneuvers are limited by the maximum torque and/or angular momentum which the wheels can provide. For an n-wheel configuration, the torque or momentum envelope can be obtained by projecting the n-dimensional hypercube, representing the domain boundary of individual wheel torques or momenta, into three dimensional space via the 3xn matrix of wheel axes. In this paper, the properties of the projected hypercube are discussed, and algorithms are proposed for determining this maximal torque or momentum envelope for general wheel configurations. Practical implementation strategies for specific wheel configurations are also considered.

History of Hubble Space Telescope (HST)

NASA Image and Video Library

1985-01-01

This is a view of a solar cell blanket deployed on a water table during the Solar Array deployment test. The Hubble Space Telescope (HST) Solar Arrays provide power to the spacecraft. The arrays are mounted on opposite sides of the HST, on the forward shell of the Support Systems Module. Each array stands on a 4-foot mast that supports a retractable wing of solar panels 40-feet (12.1-meters) long and 8.2-feet (2.5-meters) wide, in full extension. The arrays rotate so that the solar cells face the Sun as much as possible to harness the Sun's energy. The Space Telescope Operations Control Center at the Goddard Space Center operates the array, extending the panels and maneuvering the spacecraft to focus maximum sunlight on the arrays. The purpose of the HST, the most complex and sensitive optical telescope ever made, is to study the cosmos from a low-Earth orbit. By placing the telescope in space, astronomers are able to collect data that is free of the Earth's atmosphere. The HST Solar Array was designed by the European Space Agency and built by British Aerospace. The Marshall Space Flight Center had overall responsibility for design, development, and construction of the HST.

Outflow structure of the quiet sun corona probed by spacecraft radio scintillations in strong scattering

DOE Office of Scientific and Technical Information (OSTI.GOV)

Imamura, Takeshi; Ando, Hiroki; Toda, Tomoaki

Radio scintillation observations have been unable to probe flow speeds in the low corona where the scattering of radio waves is exceedingly strong. Here we estimate outflow speeds continuously from the vicinity of the Sun to the outer corona (heliocentric distances of 1.5-20.5 solar radii) by applying the strong scattering theory to radio scintillations for the first time, using the Akatsuki spacecraft as the radio source. Small, nonzero outflow speeds were observed over a wide latitudinal range in the quiet-Sun low corona, suggesting that the supply of plasma from closed loops to the solar wind occurs over an extended area.more » The existence of power-law density fluctuations down to the scale of 100 m was suggested, which is indicative of well-developed turbulence which can play a key role in heating the corona. At higher altitudes, a rapid acceleration typical of radial open fields is observed, and the temperatures derived from the speed profile show a distinct maximum in the outer corona. This study opened up a possibility of observing detailed flow structures near the Sun from a vast amount of existing interplanetary scintillation data.« less

Development of CCDs for REXIS on OSIRIS-REx

NASA Astrophysics Data System (ADS)

Ryu, Kevin K.; Burke, Barry E.; Clark, Harry R.; Lambert, Renee D.; O'Brien, Peter; Suntharalingam, Vyshnavi; Ward, Christopher M.; Warner, Keith; Bautz, Mark W.; Binzel, Richard P.; Kissel, Steven E.; Masterson, Rebecca A.

2014-07-01

The Regolith x-ray Imaging Spectrometer (REXIS) is a coded-aperture soft x-ray imaging instrument on the OSIRIS-REx spacecraft to be launched in 2016. The spacecraft will fly to and orbit the near-Earth asteroid Bennu, while REXIS maps the elemental distribution on the asteroid using x-ray fluorescence. The detector consists of a 2×2 array of backilluminated 1k×1k frame transfer CCDs with a flight heritage to Suzaku and Chandra. The back surface has a thin p+-doped layer deposited by molecular-beam epitaxy (MBE) for maximum quantum efficiency and energy resolution at low x-ray energies. The CCDs also feature an integrated optical-blocking filter (OBF) to suppress visible and near-infrared light. The OBF is an aluminum film deposited directly on the CCD back surface and is mechanically more robust and less absorptive of x-rays than the conventional free-standing aluminum-coated polymer films. The CCDs have charge transfer inefficiencies of less than 10-6, and dark current of 1e-/pixel/second at the REXIS operating temperature of -60 °C. The resulting spectral resolution is 115 eV at 2 KeV. The extinction ratio of the filter is ~1012 at 625 nm.

The terminal Velocity of the Deep Impact dust Ejecta

NASA Astrophysics Data System (ADS)

Rengel, M.; Küppers, M.; Keller, H. U.; Gutierrez, P.; Hviid, S. F.

2009-05-01

The collision of the projectile released from NASA Deep Impact spacecraft on the nucleus of comet 9P/Tempel 1 generated a hot plume. Afterwards ejecta were created, and material moved slowly in a form of a dust cloud, which dissipated during several days after the impact. Here we report a study about the distribution of terminal velocities of the particles ejected by the impact. This is performed by the development and application of an ill-conditioned inverse problem approach. We model the light-curves as seen by the Narrow Angle Camera (NAC) of OSIRIS onboard the ESA spacecraft Rosetta, and we compare them with the OSIRIS observations. Terminal velocities are derived using a maximum likelihood estimator. The dust velocity distribution is well constrained, and peaks at around 220 m s^{-1}, which is in good agreement with published estimates of the expansion velocities of the dust cloud. Measured and modeled velocity of the dust cloud suggests that the impact ejecta were quickly accelerated by the gas in the cometary coma. This analysis provides a more thorough understanding of the properties (velocity and mass of dust) of the Deep Impact dust cloud.

Spacecraft cryogenic gas storage systems

NASA Technical Reports Server (NTRS)

Rysavy, G.

1971-01-01

Cryogenic gas storage systems were developed for the liquid storage of oxygen, hydrogen, nitrogen, and helium. Cryogenic storage is attractive because of the high liquid density and low storage pressure of cryogens. This situation results in smaller container sizes, reduced container-strength levels, and lower tankage weights. The Gemini and Apollo spacecraft used cryogenic gas storage systems as standard spacecraft equipment. In addition to the Gemini and Apollo cryogenic gas storage systems, other systems were developed and tested in the course of advancing the state of the art. All of the cryogenic storage systems used, developed, and tested to date for manned-spacecraft applications are described.

Realtime Knowledge Management (RKM): From an International Space Station (ISS) Point of View

NASA Technical Reports Server (NTRS)

Robinson, Peter I.; McDermott, William; Alena, Richard L.

2004-01-01

We are developing automated methods to provide realtime access to spacecraft domain knowledge relevant a spacecraft's current operational state. The method is based upon analyzing state-transition signatures in the telemetry stream. A key insight is that documentation relevant to a specific failure mode or operational state is related to the structure and function of spacecraft systems. This means that diagnostic dependency and state models can provide a roadmap for effective documentation navigation and presentation. Diagnostic models consume the telemetry and derive a high-level state description of the spacecraft. Each potential spacecraft state description is matched against the predictions of models that were developed from information found in the pages and sections in the relevant International Space Station (ISS) documentation and reference materials. By annotating each model fragment with the domain knowledge sources from which it was derived we can develop a system that automatically selects those documents representing the domain knowledge encapsulated by the models that compute the current spacecraft state. In this manner, when the spacecraft state changes, the relevant documentation context and presentation will also change.

A Study of Learning Curve Impact on Three Identical Small Spacecraft

NASA Technical Reports Server (NTRS)

Chen, Guangming; McLennan, Douglas D.

2003-01-01

With an eye to the future strategic needs of NASA, the New Millennium Program is funding the Space Technology 5 (ST-5) project to address the future needs in the area of small satellites in constellation missions. The ST-5 project, being developed at Goddard Space Flight Center, involves the development and simultaneous launch of three small, 20-kilogram-class spacecraft. ST-5 is only a test drive and future NASA science missions may call for fleets of spacecraft containing tens of smart and capable satellites in an intelligent constellation. The objective of ST-5 project is to develop three such pioneering small spacecraft for flight validation of several critical new technologies. The ST-5 project team at Goddard Space Flight Center has completed the spacecraft design, is now building and testing the three flight units. The launch readiness date (LRD) is in December 2005. A critical part of ST-5 mission is to prove that it is possible to build these small but capable spacecraft with recurring cost low enough to make future NASA s multi- spacecraft constellation missions viable from a cost standpoint.

Solution for Direct Solar Impingement Problem on Landsat-7 ETM+ Cooler Door During Cooler Outgas in Flight

NASA Technical Reports Server (NTRS)

Choi, Michael K.

1999-01-01

There was a thermal anomaly of the Landsat-7 Enhanced Thematic Mapper Plus (ETM+) radiative cooler cold stage during the cooler outgas phase in flight. With the cooler door in the outgas position and the outgas heaters enabled, the cold stage temperature increased to a maximum of 323 K when the spacecraft was in the sunlight, which was warmer than the 316.3 K upper set point of the outgas heater controller on the cold stage. Also, the outgas heater cycled off when the cold stage was warming up to 323 K. A corrective action was taken before the attitude of the spacecraft was changed during the first week in flight. One orbit before the attitude was changed, the outgas heaters were disabled to cool off the cold stage. The cold stage temperature increase was strongly dependent on the spacecraft roll and yaw. It provided evidence that direct solar radiation entered the gap between the cooler door and cooler shroud. There was a concern that the direct solar radiation could cause polymerization of hydrocarbons, which could contaminate the cooler and lead to a thermal short. After outgas with the cooler door in the outgas position for seven days, the cooler door was changed to the fully open position. With the cooler door fully open, the maximum cold stage temperature was 316.3 K when the spacecraft was in the sunlight, and the duty cycle of the outgas heater in the eclipse was the same as that in the sunlight. It provided more evidence that direct solar radiation had entered the gap between the cooler door and cooler shroud. Cooler outgas continued for seven more days, with the cooler door fully open. The corrective actions had prevented overheating of the cold stage and cold focal plane array (CFPA), which could damage these two components. They also minimized the risk of contamination on the cold stage, which could lead to a thermal short.

Science and software support for spacecraft solar occultation experiments

NASA Technical Reports Server (NTRS)

Hessameddin, G.; Becher, J.

1982-01-01

The temperature dependence of absorption coefficients of ozone was studied between 7567 A and 3630 A. When the gas was cooled from room temperature to -108 C, an overall increase in the absorption coefficients was noticed. The maximum increase of 5% occurred at lambda = 6020 A. In general, the absorption is linearly dependent on temperature.

Thermoelectric Outer Planets Spacecraft (TOPS)

NASA Technical Reports Server (NTRS)

1973-01-01

The research and advanced development work is reported on a ballistic-mode, outer planet spacecraft using radioisotope thermoelectric generator (RTG) power. The Thermoelectric Outer Planet Spacecraft (TOPS) project was established to provide the advanced systems technology that would allow the realistic estimates of performance, cost, reliability, and scheduling that are required for an actual flight mission. A system design of the complete RTG-powered outer planet spacecraft was made; major technical innovations of certain hardware elements were designed, developed, and tested; and reliability and quality assurance concepts were developed for long-life requirements. At the conclusion of its active phase, the TOPS Project reached its principal objectives: a development and experience base was established for project definition, and for estimating cost, performance, and reliability; an understanding of system and subsystem capabilities for successful outer planets missions was achieved. The system design answered long-life requirements with massive redundancy, controlled by on-board analysis of spacecraft performance data.

Parabolic Flight Evaluation of Bacterial Adhesion on Multiple Antimicrobial Surface Treatments

NASA Technical Reports Server (NTRS)

Birmele, Michele

2011-01-01

This report describes the development of a test method and the evaluation of the effectiveness of antimicrobial technologies in reduced gravity based on parabolic flight experiments. Microbial growth is a common occurrence on fully immersed wetted surfaces in spacecraft environmental control and life support systems despite the use of chemical and/or physical \\disinfection. Many materials and surface treatments with antimicrobial properties are commercially available but none have been vetted for spaceflight applications. Herein a test method is explained that included ground and reduced gravity parabolic flight experiments with a standard microorganism recovered from spacecraft, Pseudomonas aeruginosa, added at a concentration of 1 x 10(exp 5) cells per milliliter (mL) onto challenge material coupon surfaces. Several experimental materials were observed to slightly reduce microbial attachment in reduced gravity flight experiments, but none were capable of eliminating all challenge bacteria. Lunar gravity had an increased antimicrobial effect in 28 out of 36 test coupons compared to microgravity when provided otherwise identical conditions for growth, suggesting trace .amounts of gravity may be required for maximum antimicrobial performance. Bacterial cells exposed to variable gravity had more than twice as ,much intracellular adenosine triphosphate (ATP) when compared to control cells exposed only to Earth gravity due to a short duration response to environmental stress. An ATP luminescence assay was the method most amenable to development of an in-flight microbial monitoring assay

A Comparison of Learning Technologies for Teaching Spacecraft Software Development

ERIC Educational Resources Information Center

Straub, Jeremy

2014-01-01

The development of software for spacecraft represents a particular challenge and is, in many ways, a worst case scenario from a design perspective. Spacecraft software must be "bulletproof" and operate for extended periods of time without user intervention. If the software fails, it cannot be manually serviced. Software failure may…

Update of the ISTP Solar Maximum Mission: ISTP Project Scientist for Theory and Ground-Based Observations

NASA Technical Reports Server (NTRS)

Curtis, Steve

1999-01-01

Building upon the numerous successes of the pre-solar maximum International Solar Terrestrial Physics (ISTP) mission, the ISTP Solar Maximum Mission is expected to produce new insights into global flow of energy, momentum, and mass, from the Sun, through the heliosphere, into the magnetosphere and to their final deposition in the terrestrial upper atmosphere/ionosphere system. Of particular interest is the determination of the geo-effectiveness of solar events, principally Coronal Mass Ejections (CMEs). Given the expected increased frequency and strength of CMEs during the Solar Maximum period, a major advance in our understanding of nature of the coupling of CMEs to the magnetosphere-ionosphere-atmosphere system is expected. The roles during this time of the various ISTP assets will be discussed. These assets will include the SOHO, Wind, Polar, and Geotail spacecraft, the ground-based observing networks and the theory tools.

Analytical Modeling of Pressure Wall Hole Size and Maximum Tip-to-Tip Crack Length for Perforating Normal and Oblique Orbital Debris Impacts

NASA Technical Reports Server (NTRS)

Schonberg, William P.; Mohamed, Essam

1997-01-01

This report presents the results of a study whose objective was to develop first-principles-based models of hole size and maximum tip-to-tip crack length for a spacecraft module pressure wall that has been perforated in an orbital debris particle impact. The hole size and crack length models are developed by sequentially characterizing the phenomena comprising the orbital debris impact event, including the initial impact, the creation and motion of a debris cloud within the dual-wall system, the impact of the debris cloud on the pressure wall, the deformation of the pressure wall due to debris cloud impact loading prior to crack formation, pressure wall crack initiation, propagation, and arrest, and finally pressure wall deformation following crack initiation and growth. The model development has been accomplished through the application of elementary shock physics and thermodynamic theory, as well as the principles of mass, momentum, and energy conservation. The predictions of the model developed herein are compared against the predictions of empirically-based equations for hole diameters and maximum tip-to-tip crack length for three International Space Station wall configurations. The ISS wall systems considered are the baseline U.S. Lab Cylinder, the enhanced U.S. Lab Cylinder, and the U.S. Lab Endcone. The empirical predictor equations were derived from experimentally obtained hole diameters and crack length data. The original model predictions did not compare favorably with the experimental data, especially for cases in which pressure wall petalling did not occur. Several modifications were made to the original model to bring its predictions closer in line with the experimental results. Following the adjustment of several empirical constants, the predictions of the modified analytical model were in much closer agreement with the experimental results.

Materials and structures technology insertion into spacecraft systems: Successes and challenges

NASA Astrophysics Data System (ADS)

Rawal, Suraj

2018-05-01

Over the last 30 years, significant advancements have led to the use of multifunctional materials and structures technologies in spacecraft systems. This includes the integration of adaptive structures, advanced composites, nanotechnology, and additive manufacturing technologies. Development of multifunctional structures has been directly influenced by the implementation of processes and tools for adaptive structures pioneered by Prof. Paolo Santini. Multifunctional materials and structures incorporating non-structural engineering functions such as thermal, electrical, radiation shielding, power, and sensors have been investigated. The result has been an integrated structure that offers reduced mass, packaging volume, and ease of integration for spacecraft systems. Current technology development efforts are being conducted to develop innovative multifunctional materials and structures designs incorporating advanced composites, nanotechnology, and additive manufacturing. However, these efforts offer significant challenges in the qualification and acceptance into spacecraft systems. This paper presents a brief overview of the technology development and successful insertion of advanced material technologies into spacecraft structures. Finally, opportunities and challenges to develop and mature next generation advanced materials and structures are presented.

Distributed Autonomous Control of Multiple Spacecraft During Close Proximity Operations

DTIC Science & Technology

2007-12-01

programs may be the XSS-11. The AFRL Space Vehicle Directorate at Kirtland Air Force Base in New Mexico developed the XSS-11 in order to exhibit the...the LQR/APF algorithm appears to be a promising new development for the field of multiple spacecraft close proximity maneuver control. Monte...dissertation reports the development of an autonomous distributed control algorithm for multiple spacecraft during close proximity operations

Importance of temperature control for HEFLEX, a biological experiment for Spacelab 1. [plant gravitational physiology study

NASA Technical Reports Server (NTRS)

Chapman, D. K.; Brown, A. H.

1979-01-01

The importance of temperature control to HEFLEX, a Spacelab experiment designed to measure kinetic properties of Helianthis nutation in a low-g environment, is discussed. It is argued that the development of the HEFLEX experiment has been severely hampered by the inadequate control of ambient air temperature provided by the spacecraft module design. A worst case calculation shows that delivery of only 69% of the maximum yield of useful data from the HEFLEX system is guaranteed; significant data losses from inadequate temperature control are expected. The magnitude of the expected data losses indicates that the cost reductions associated with imprecise temperature controls may prove to be a false economy in the long term.

Studies of the gas tori of Titan and Triton

NASA Technical Reports Server (NTRS)

Smyth, William H.

1995-01-01

Progress in the development of the model for the circumplanetary distribution of atomic hydrogen in the Saturn system produced by a Titan source is discussed. Because of the action of the solar radiation acceleration and the obliquity of Saturn, the hydrogen distribution is shown to undergo seasonal changes as the planet moves about the Sun. Preliminary model calculations show that for a continuous Titan source, the H distribution is highly asymmetric about the planet and has a density maximum near the dusk side of Saturn, qualitatively similar to the pattern recently deduced by Shemansky and Hall from observations acquired by the UVS instruments aboard the Voyager spacecrafts. The investigation of these Voyager data will be undertaken in the next project year.

PSAW/MicroSWIS [Microminiature Surface Acoustic Wave (SAW) based Wirelesss Instrumentation System

NASA Technical Reports Server (NTRS)

Heermann, Doug; Krug, Eric

2004-01-01

This Final Report for the PSAW/MicroSWIS Program is provided in compliance with contract number NAS3-01118. This report documents the overall progress of the program and presents project objectives, work carried out, and results obtained. Program Conceptual Design Package stated the following objectives: To develop a sensor/transceiver network that can support networking operations within spacecraft with sufficient bandwidth so that (1) flight control data, (2) avionics data, (3) payload/experiment data, and (4) prognostic health monitoring sensory information can flow to appropriate locations at frequencies that contain the maximum amount of information content but require minimum interconnect and power: a very high speed, low power, programmable modulation, spread-spectrum radio sensor/transceiver.

TOPEX orbital radiation study

NASA Technical Reports Server (NTRS)

Stassinopoulos, E. G.; Barth, J. M.

1984-01-01

The space radiation environment of the TOPEX spacecraft is investigated. A single trajectory was considered. The external (surface incident) charged particle radiation, predicted for the satellite, is determined by orbital flux integration for the specified trajectory. The latest standard models of the environment are used in the calculations. The evaluation is performed for solar maximum conditions. The spacecraft exposure to cosmic rays of galactic origin is evaluated over its flight path through the magnetosphere in terms of geomagnetic shielding effects, both for surface incident heavy ions and for particles emerging behind different material thickness. Limited shielding and dose evaluations are performed for simple infinite slab and spherical geometries. Results, given in graphical and tabular form, are analyzed, explained, and discussed. Conclusions are presented and commented on.

Optimum detection of tones transmitted by a spacecraft

NASA Technical Reports Server (NTRS)

Simon, M. K.; Shihabi, M. M.; Moon, T.

1995-01-01

The performance of a scheme proposed for automated routine monitoring of deep-space missions is presented. The scheme uses four different tones (sinusoids) transmitted from the spacecraft (S/C) to a ground station with the positive identification of each of them used to indicate different states of the S/C. Performance is measured in terms of detection probability versus false alarm probability with detection signal-to-noise ratio as a parameter. The cases where the phase of the received tone is unknown and where both the phase and frequency of the received tone are unknown are treated separately. The decision rules proposed for detecting the tones are formulated from average-likelihood ratio and maximum-likelihood ratio tests, the former resulting in optimum receiver structures.

New model framework and structure and the commonality evaluation model. [concerning unmanned spacecraft projects

NASA Technical Reports Server (NTRS)

1977-01-01

The development of a framework and structure for shuttle era unmanned spacecraft projects and the development of a commonality evaluation model is documented. The methodology developed for model utilization in performing cost trades and comparative evaluations for commonality studies is discussed. The model framework consists of categories of activities associated with the spacecraft system's development process. The model structure describes the physical elements to be treated as separate identifiable entities. Cost estimating relationships for subsystem and program-level components were calculated.

A Reconfigurable Testbed Environment for Spacecraft Autonomy

NASA Technical Reports Server (NTRS)

Biesiadecki, Jeffrey; Jain, Abhinandan

1996-01-01

A key goal of NASA's New Millennium Program is the development of technology for increased spacecraft on-board autonomy. Achievement of this objective requires the development of a new class of ground-based automony testbeds that can enable the low-cost and rapid design, test, and integration of the spacecraft autonomy software. This paper describes the development of an Autonomy Testbed Environment (ATBE) for the NMP Deep Space I comet/asteroid rendezvous mission.

Design of a power management and distribution system for a thermionic-diode powered spacecraft

NASA Technical Reports Server (NTRS)

Kimnach, Greg L.

1996-01-01

The Electrical Systems Development Branch of the Power Technology Division at the NASA Lewis Research Center in Cleveland, Ohio is designing a Power Management and Distribution (PMAD) System for the Air Force's Integrated Solar Upper Stage (ISUS) Engine Ground Test Demonstration (EGD). The ISUS program uses solar-thermal propulsion to perform orbit transfers from Low Earth Orbit (LEO) to Geosynchronous Orbit (GEO) and from LEO to Molnya. The ISUS uses the same energy conversion receiver to perform the LEO to High Earth Orbit (HEO) transfer and to generate on-orbit electric power for the payloads. On-orbit power generation is accomplished via two solar concentrators heating a dual-cavity graphite-core which has Thermionic Diodes (TMD's) encircling each cavity. The graphite core and concentrators together are called the Receiver and Concentrator (RAC). The TDM-emitters reach peak temperatures of approximately 2200K, and the TID-collectors are run at approximately 1000K. Because of the high Specific Impulse (I(sup sp)) of solar thermal propulsion relative to chemical propulsion, and because a common bus is used for communications, GN&C, power, etc., a substantial increase in payload weight is possible. This potentially allows for a stepdown in the required launch vehicle size or class for similar payload weight using conventional chemical propulsion and a separate spacecraft bus. The ISUS power system is to provide 1000W(sub e) at 28+/-6V(sub dc) to the payload/spacecraft from a maximum TID generation capability of 1070W(sub e) at 2200K. Producing power with this quality, protecting the spacecraft from electrical faults and accommodating operational constraints of the TID's are the responsibilities of the PMAD system. The design strategy and system options examined along with the proposed designs for the Flight and EGD configurations are discussed herein.

Underexpanded Supersonic Plume Surface Interactions: Applications for Spacecraft Landings on Planetary Bodies

NASA Technical Reports Server (NTRS)

Mehta, M.; Sengupta, A.; Renno, N. O.; Norman, J. W.; Gulick, D. S.

2011-01-01

Numerical and experimental investigations of both far-field and near-field supersonic steady jet interactions with a flat surface at various atmospheric pressures are presented in this paper. These studies were done in assessing the landing hazards of both the NASA Mars Science Laboratory and Phoenix Mars spacecrafts. Temporal and spatial ground pressure measurements in conjunction with numerical solutions at altitudes of approx.35 nozzle exit diameters and jet expansion ratios (e) between 0.02 and 100 are used. Data from steady nitrogen jets are compared to both pulsed jets and rocket exhaust plumes at Mach approx.5. Due to engine cycling, overpressures and the plate shock dynamics are different between pulsed and steady supersonic impinging jets. In contrast to highly over-expanded (e <1) and underexpanded exhaust plumes, results show that there is a relative ground pressure load maximum for moderately underexpanded (e approx.2-5) jets which demonstrate a long collimated plume shock structure. For plumes with e much >5 (lunar atmospheric regime), the ground pressure is minimal due to the development of a highly expansive shock structure. We show this is dependent on the stability of the plate shock, the length of the supersonic core and plume decay due to shear layer instability which are all a function of the jet expansion ratio. Asymmetry and large gradients in the spatial ground pressure profile and large transient overpressures are predominantly linked to the dynamics of the plate shock. More importantly, this study shows that thruster plumes exhausting into martian environments possess the largest surface pressure loads and can occur at high spacecraft altitudes in contrast to the jet interactions at terrestrial and lunar atmospheres. Theoretical and analytical results also show that subscale supersonic cold gas jets adequately simulate the flow field and loads due to rocket plume impingement provided important scaling parameters are in agreement. These studies indicate the critical importance of testing and modeling plume-surface interactions for descent and ascent of spacecraft and launch vehicles.

Atmospheric drag model calibrations for spacecraft lifetime prediction

NASA Technical Reports Server (NTRS)

Binebrink, A. L.; Radomski, M. S.; Samii, M. V.

1989-01-01

Although solar activity prediction uncertainty normally dominates decay prediction error budget for near-Earth spacecraft, the effect of drag force modeling errors for given levels of solar activity needs to be considered. Two atmospheric density models, the modified Harris-Priester model and the Jacchia-Roberts model, to reproduce the decay histories of the Solar Mesosphere Explorer (SME) and Solar Maximum Mission (SMM) spacecraft in the 490- to 540-kilometer altitude range were analyzed. Historical solar activity data were used in the input to the density computations. For each spacecraft and atmospheric model, a drag scaling adjustment factor was determined for a high-solar-activity year, such that the observed annual decay in the mean semimajor axis was reproduced by an averaged variation-of-parameters (VOP) orbit propagation. The SME (SMM) calibration was performed using calendar year 1983 (1982). The resulting calibration factors differ by 20 to 40 percent from the predictions of the prelaunch ballistic coefficients. The orbit propagations for each spacecraft were extended to the middle of 1988 using the calibrated drag models. For the Jaccia-Roberts density model, the observed decay in the mean semimajor axis of SME (SMM) over the 4.5-year (5.5-year) predictive period was reproduced to within 1.5 (4.4) percent. The corresponding figure for the Harris-Priester model was 8.6 (20.6) percent. Detailed results and conclusions regarding the importance of accurate drag force modeling for lifetime predictions are presented.

Catalog of lunar mission data

NASA Technical Reports Server (NTRS)

Mantel, E. J. (Editor); Miller, E. R. (Editor)

1977-01-01

Several series of spacecraft were developed, designed, built and launched to determine different characteristics of the lunar surface and environment for a manned landing. Both unmanned and manned spacecrafts, spacecraft equipment and lunar missions are documented.

Assembly Platform For Use In Outer Space

NASA Technical Reports Server (NTRS)

Rao, Niranjan S.; Buddington, Patricia A.

1995-01-01

Report describes conceptual platform or framework for use in assembling other structures and spacecraft in outer space. Consists of three fixed structural beams comprising central beam and two cross beams. Robotic manipulators spaced apart on platform to provide telerobotic operation of platform by either space-station or ground crews. Platform and attached vehicles function synergistically to achieve maximum performance for intended purposes.

Low nutation-rate dampers

NASA Technical Reports Server (NTRS)

Tossman, B. E.

1971-01-01

Mission requirements plus spacecraft weight and power constraints often reduce the excitation frequency of a nutation damper below 1 cpm. Since attitude stability is determined by damper performance, maximum effectiveness at low rates is demanded. Presented are design considerations that low-frequency dampers require, along with descriptions of two low-frequency systems: the Direct Measurement Explorer 1 and the Small Astronomy Satellite A (SAS-A).

ART CONCEPTS - APOLLO IX

NASA Image and Video Library

1969-02-20

S69-19796 (February 1969) --- Composite of six artist's concepts illustrating key events, tasks and activities on the fifth day of the Apollo 9 mission, including vehicles undocked, Lunar Module burns for rendezvous, maximum separation, ascent propulsion system burn, formation flying and docking, and Lunar Module jettison ascent burn. The Apollo 9 mission will evaluate spacecraft lunar module systems performance during manned Earth-orbital flight.

Thermal Design to Meet Stringent Temperature Gradient/Stability Requirements of SWIFT BAT Detectors

NASA Technical Reports Server (NTRS)

Choi, Michael K.

2000-01-01

The Burst Alert Telescope (BAT) is an instrument on the National Aeronautics and Space Administration (NASA) SWIFT spacecraft. It is designed to detect gamma ray burst over a broad region of the sky and quickly align the telescopes on the spacecraft to the gamma ray source. The thermal requirements for the BAT detector arrays are very stringent. The maximum allowable temperature gradient of the 256 cadmium zinc telluride (CZT) detectors is PC. Also, the maximum allowable rate of temperature change of the ASICs of the 256 Detector Modules (DMs) is PC on any time scale. The total power dissipation of the DMs and Block Command & Data Handling (BCDH) is 180 W. This paper presents a thermal design that uses constant conductance heat pipes (CCHPs) to minimize the temperature gradient of the DMs, and loop heat pipes (LHPs) to transport the waste heat to the radiator. The LHPs vary the effective thermal conductance from the DMs to the radiator to minimize heater power to meet the heater power budget, and to improve the temperature stability. The DMs are cold biased, and active heater control is used to meet the temperature gradient and stability requirements.

Design and testing of a high power spacecraft thermal management system

NASA Technical Reports Server (NTRS)

Mccabe, Michael E., Jr.; Ku, Jentung; Benner, Steve

1988-01-01

The design and test results are presented of an ammonia hybrid capillary pumped loop thermal control system which could be used for heat acquisition and transport on future large space platforms and attached payloads, such as those associated with the NASA Space Station. The High Power Spacecraft Thermal Management System (HPSTM) can operate as either a passive, capillary pumped two phase thermal control system, or, when additional pressure head is required, as a mechanically pumped loop. Testing has shown that in the capillary mode, the HPSTM evaporators can acquire a total heat load of between 600 W and 24 kW, transported over 10 meters, at a maximum heat flux density of 4.3 W/sq cm. With the mechanical pump circulating the ammonia, a heat acquisition potential of 52 kW was demonstrated for 15 minutes without an evaporator failure. These results represent a significant improvement over the maximum transport capability previously displayed in other capillary systems. The HPSTM system still retains the proven capillary capabilities of heat load sharing and flow control between evaporator plates, rapid power cycling, and nonuniform heating in both the capillary and hybrid operating modes.

Qualification of a Multi-Channel Infrared Laser Absorption Spectrometer for Monitoring CO, HCl, HCN, HF, and CO2 Aboard Manned Spacecraft

NASA Technical Reports Server (NTRS)

Briggs, Ryan M.; Frez, Clifford; Forouhar, Siamak; May, Randy D.; Meyer, Marit E.; Kulis, Michael J.; Berger, Gordon M.

2015-01-01

Monitoring of specific combustion products can provide early-warning detection of accidental fires aboard manned spacecraft and also identify the source and severity of combustion events. Furthermore, quantitative in situ measurements are important for gauging levels of exposure to hazardous gases, particularly on long-duration missions where analysis of returned samples becomes impractical. Absorption spectroscopy using tunable laser sources in the 2 to 5 micrometer wavelength range enables accurate, unambiguous detection of CO, HCl, HCN, HF, and CO2, which are produced in varying amounts through the heating of electrical components and packaging materials commonly used aboard spacecraft. Here, we report on calibration and testing of a five-channel laser absorption spectrometer designed to accurately monitor ambient gas-phase concentrations of these five compounds, with low-level detection limits based on the Spacecraft Maximum Allowable Concentrations. The instrument employs a two-pass absorption cell with a total optical pathlength of 50 cm and a dedicated infrared semiconductor laser source for each target gas. We present results from testing the five-channel sensor in the presence of trace concentrations of the target compounds that were introduced using both gas sources and oxidative pyrolysis (non-flaming combustion) of solid material mixtures.

Electron holes in phase space: What they are and why they matter

NASA Astrophysics Data System (ADS)

Hutchinson, I. H.

2017-05-01

This is a tutorial and selective review explaining the fundamental concepts and some currently open questions concerning the plasma phenomenon of the electron hole. The widespread occurrence of electron holes in numerical simulations, space-craft observations, and laboratory experiments is illustrated. The elementary underlying theory is developed of a one-dimensional electron hole as a localized potential maximum, self-consistently sustained by a deficit of trapped electron phase-space density. The spatial extent of a hole is typically a few Debye lengths; what determines the minimum and maximum possible lengths is explained, addressing the key aspects of the as yet unsettled dispute between the integral and differential approaches to hole structure. In multiple dimensions, holes tend to form less readily; they generally require a magnetic field and distribution-function anisotropy. The mechanisms by which they break up are explained, noting that this transverse instability is not fully understood. Examples are given of plasma circumstances where holes play an important role, and of recent progress on understanding their holistic kinematics and self-acceleration.

Observations of a Unique Type of ULF Waves by Low-Latitude Space Technology Five Mission

NASA Technical Reports Server (NTRS)

Le, G.; Chi, P.; Strangeway, R. J.; Slavin, J. A.

2011-01-01

We report a unique type of ULF waves observed by low-altitude Space Technology 5 (ST-5) constellation mission. ST-5 is a three micro-satellite constellation deployed into a 300 x 4500 km, dawn-dusk, and sun synchronous polar orbit with 105.6deg inclination angle. Due to the Earth s rotation and the dipole tilt effect, the spacecraft s dawn-dusk orbit track can reach as low as subauroral latitudes during the course of a day. Whenever the spacecraft traverse across the dayside closed field line region at subauroral latitudes, they frequently observe strong transverse oscillations at 30-200 mHz, or in the Pc 2-3 frequency range. These Pc 2-3 waves appear as wave packets with durations in the order of 5-10 minutes. As the maximum separations of the ST-5 spacecraft are in the order of 10 minutes, the three ST-5 satellites often observe very similar wave packets, implying these wave oscillations occur in a localized region. The coordinated ground-based magnetic observations at the spacecraft footprints, however, do not see waves in the Pc 2-3 band; instead, the waves appear to be the common Pc 4-5 waves associated with field line resonances. We suggest that these unique Pc 2-3 waves seen by ST-5 are in fact the Doppler-shifted Pc 4-5 waves as a result of rapid traverse of the spacecraft across the resonant field lines azimuthally at low altitudes. The observations with the unique spacecraft dawn-disk orbits at proper altitudes and magnetic latitudes reveal the azimuthal characteristics of field-aligned resonances.

Metabolism of spacecraft cleaning reagents by Mars Odyssey and Phoenix-associated Acinetobacter

NASA Astrophysics Data System (ADS)

Mogul, Rakesh; Barding, Gregory; Baki, Ryan; Perkins, Nicole; Lee, Sooji; Lalla, Sid; Campos, Alexa; Sripong, Kimberly; Madrid, Steve

2016-07-01

The metabolomic and proteomic properties that promote microbial survival in spacecraft assembly facilities are important aspects to planetary protection and astrobiology. In this presentation, we will provide molecular and biological evidence that the spacecraft-associated Acinetobacter metabolize/degrade spacecraft cleaning reagents such as ethanol, 2-propanol, and Kleenol-30. Gas chromatography-mass spectrometry (GC-MS) studies on A. radioresistens 50v1 (Mars Odyssey) show that the metabolome is dependent upon growth conditions and that ^{13}C-labeled ethanol is incorporated into metabolites such as TCA/glyoxylate cycle intermediates, amino acids, monosaccharides, and disaccharides (e.g., trehalose). In fact, plate count assays show that ethanol is a sole carbon source under minimal conditions for several Mars Phoenix and Odyssey-associated Acinetobacter strains, which may explain why the Acinetobacter are among the most abundant genera found in spacecraft assembly facilities. Biochemical analyses support the enzymatic oxidation of ethanol and 2-propanol by a membrane-bound and NAD+/PQQ-dependent alcohol dehydrogenase, with current kinetic data providing similar apparent K _{M} and maximum growth rate values of ˜5 and 8 mM ethanol, respectively. Preliminary GC-MS analysis also suggests that Kleenol-30 is degraded by A. radioresistens 50v1 when grown in ethanol mixtures. Under minimal conditions, A. radioresistens 50v1 (˜10 ^{8} cfu/mL) also displays a remarkable oxidative extremotolerance (˜2-log reduction in 10 mM hydrogen peroxide), which suggests crucial roles for metabolites associated with oxidative stress (e.g., trehalose) and the observed appreciable catalase specific activities. In conclusion, these results provide key insights into the survival strategies of spacecraft-associated Acinetobacter and emphasize the importance of characterizing the carbon metabolism of forward contaminants.

Observations of a Unique Type of ULF Waves by Low-Latitude Space Technology 5 Satellites

NASA Technical Reports Server (NTRS)

Le, G.; Chi, P. J.; Strangeway, R. J.; Slavin, J. A.

2011-01-01

We report a unique type of ULF waves observed by low-altitude Space Technology 5 (ST-5) constellation mission. ST-5 is a three micro-satellite constellation deployed into a 300 x 4500 km, dawn-dusk, and sun synchronous polar orbit with 105.6deg inclination angle. Due to the Earth s rotation and the dipole tilt effect, the spacecraft s dawn-dusk orbit track can reach as low as subauroral latitudes during the course of a day. Whenever the spacecraft traverse across the dayside closed field line region at subauroral latitudes, they frequently observe strong transverse oscillations at 30-200 mHz, or in the Pc 2-3 frequency range. These Pc 2-3 waves appear as wave packets with durations in the order of 5-10 minutes. As the maximum separations of the ST-5 spacecraft are in the order of 10 minutes, the three ST-5 satellites often observe very similar wave packets, implying these wave oscillations occur in a localized region. The coordinated ground-based magnetic observations at the spacecraft footprints, however, do not see waves in the Pc 2-3 band; instead, the waves appear to be the common Pc 4-5 waves associated with field line resonances. We suggest that this unique Pc 2-3 waves seen by ST-5 are in fact the Doppler-shifted Pc 4-5 waves as a result of rapid traverse of the spacecraft across the resonant field lines azimuthally at low altitudes. The observations with the unique spacecraft dawn-disk orbits at proper altitudes and magnetic latitudes reveal the azimuthal characteristics of field-aligned resonances.

Observations of a Unique Type of ULF Waves by Low-Latitude Space Technology 5 Mission

NASA Technical Reports Server (NTRS)

Le, Guan; Chi, P.; Strangeway, R. J.; Slavin, J. A.

2011-01-01

We report a unique type of ULF waves observed by low-altitude Space Technology 5 (ST-5) constellation mission. ST-5 is a three micro-satellite constellation deployed into a 300 x 4500 km, dawn-dusk, and sun synchronous polar orbit with 105.6 inclination angle. Due to the Earth's rotation and the dipole tilt effect, the spacecraft's dawn-dusk orbit track can reach as low as sub auroral latitudes during the course of a day. Whenever the spacecraft traverse across the dayside closed field line region at sub auroral latitudes, they frequently observe strong transverse oscillations at 30-200 mHz, or in the Pc 2-3 frequency range. These Pc 2-3 waves appear as wave packets with durations in the order of 5-10 minutes. As the maximum separations of the ST-5 spacecraft are in the order of 10 minutes, the three ST -5 satellites often observe very similar wave packets, implying these wave oscillations occur in a localized region. The coordinated ground-based magnetic observations at the spacecraft footprints, however, do not see waves in the Pc 2-3 band; instead, the waves appear to be the common Pc 4-5 waves associated with field line resonances. We suggest that these unique Pc 2-3 waves seen by ST-5 are in fact the Doppler-shifted Pc 4-5 waves as a result of rapid traverse of the spacecraft across the resonant field lines azimuthally at low altitudes. The observations with the unique spacecraft dawn-disk orbits at proper altitudes and magnetic latitudes reveal the azimuthal characteristics of field-aligned resonances.

Observations of a Unique Type of ULF Waves by Low-Latitude Space Technology 5 Mission

NASA Technical Reports Server (NTRS)

Le, G.; Chi, P.; Strangeway, R. J.; Slavin, J. A.

2010-01-01

We report a unique type of ULF waves observed by low-altitude Space Technology 5 (ST-5) constellation mission. ST-5 is a three micro-satellite constellation deployed into a 300 x 4500 km, dawn-dusk, and sun synchronous polar orbit with 105.6 degree inclination angle. Due to the Earth's rotation and the dipole tilt effect, the spacecraft's dawn-dusk orbit track can reach as low as subauroral latitudes during the course of a day. Whenever the spacecraft traverse across the dayside closed field line region at sub auroral latitudes, they frequently observe strong transverse oscillations at 30-200 mHz, or in the Pc 2-3 frequency range. These Pc 2-3 waves appear as wave packets with durations in the order of 5-10 minutes. As the maximum separations of the ST-5 spacecraft are in the order of 10 minutes, the three ST-5 satellites often observe very similar wave packets, implying these wave oscillations occur in a localized region. The coordinated ground-based magnetic observations at the spacecraft footprints, however, do not see waves in the Pc 2-3 band; instead, the waves appear to be the common Pc 4-5 waves associated with field line resonances. We suggest that these unique Pc 2-3 waves seen by ST-5 are in fact the Doppler-shifted Pc 4-5 waves as a result of rapid traverse of the spacecraft across the resonant field lines azimuthally at low altitudes. The observations with the unique spacecraft dawn-disk orbits at proper altitudes and magnetic latitudes reveal the azimuthal characteristics of field-aligned resonances.

A shock spectra and impedance method to determine a bound for spacecraft structural loads

NASA Technical Reports Server (NTRS)

Bamford, R.; Trubert, M.

1974-01-01

A method to determine a bound of structural loads for a spacecraft mounted on a launch vehicle is developed. The method utilizes the interface shock spectra and the relative impedance of the spacecraft and launch vehicle. The method is developed for single-degree-of-freedom models and then generalized to multidegree-of-freedom models.

Spacecraft Environmental Anomalies Handbook

DTIC Science & Technology

1989-08-01

1989 4. TITLE AND SUBTITLE S. FUNDING NUMBERS SPACECRAFT ENVIRONMENTAL ANOMALIES HANDBOOK 282201AA PE: 63410F 6. AUTHOR(S) Paul A. Robinson, Jr 7...engineering solutions for mitigating the effects of environmental anomalies have been developed. Among the causes o, spacecraft anomalies are surface...have been discovered after years of investig!:tion, and engineering solutions for mitigating the effccts of environmental anomalies have been developed

Lean spacecraft avionics trade study

NASA Technical Reports Server (NTRS)

Main, John A.

1994-01-01

Spacecraft design is generally an exercise in design trade-offs: fuel vs. weight, power vs. solar cell area, radiation exposure vs. shield weight, etc. Proper analysis of these trades is critical in the development of lightweight, efficient, 'lean' satellites. The modification of the launch plans for the Magnetosphere Imager (MI) to a Taurus launcher from the much more powerful Delta has forced a reduction in spacecraft weight availability into the mission orbit from 1300 kg to less than 500 kg. With weight now a driving factor it is imperative that the satellite design be extremely efficient and lean. The accuracy of engineering trades now takes on an added importance. An understanding of spacecraft subsystem interactions is critical in the development of a good spacecraft design, yet it is a challenge to define these interactions while the design is immature. This is currently an issue in the development of the preliminary design of the MI. The interaction and interfaces between this spacecraft and the instruments it carries are currently unclear since the mission instruments are still under development. It is imperative, however, to define these interfaces so that avionics requirements ideally suited to the mission's needs can be determined.

Virtual Satellite

NASA Technical Reports Server (NTRS)

Hammrs, Stephan R.

2008-01-01

Virtual Satellite (VirtualSat) is a computer program that creates an environment that facilitates the development, verification, and validation of flight software for a single spacecraft or for multiple spacecraft flying in formation. In this environment, enhanced functionality and autonomy of navigation, guidance, and control systems of a spacecraft are provided by a virtual satellite that is, a computational model that simulates the dynamic behavior of the spacecraft. Within this environment, it is possible to execute any associated software, the development of which could benefit from knowledge of, and possible interaction (typically, exchange of data) with, the virtual satellite. Examples of associated software include programs for simulating spacecraft power and thermal- management systems. This environment is independent of the flight hardware that will eventually host the flight software, making it possible to develop the software simultaneously with, or even before, the hardware is delivered. Optionally, by use of interfaces included in VirtualSat, hardware can be used instead of simulated. The flight software, coded in the C or C++ programming language, is compilable and loadable into VirtualSat without any special modifications. Thus, VirtualSat can serve as a relatively inexpensive software test-bed for development test, integration, and post-launch maintenance of spacecraft flight software.

Making Human Spaceflight Practical and Affordable: Spacecraft Designs and their Degree of Operability

NASA Technical Reports Server (NTRS)

Crocker, Alan R.

2011-01-01

As we push toward new and diverse space transportation capabilities, reduction in operations cost becomes increasingly important. Achieving affordable and safe human spaceflight capabilities will be the mark of success for new programs and new providers. The ability to perceive the operational implications of design decisions is crucial in developing safe yet cost competitive space transportation systems. Any human spaceflight program - government or commercial - must make countless decisions either to implement spacecraft system capabilities or adopt operational constraints or workarounds to account for the lack of such spacecraft capabilities. These decisions can benefit from the collective experience that NASA has accumulated in building and operating crewed spacecraft over the last five decades. This paper reviews NASA s history in developing and operating human rated spacecraft, reviewing the key aspects of spacecraft design and their resultant impacts on operations phase complexity and cost. Specific examples from current and past programs - including the Space Shuttle and International Space Station - are provided to illustrate design traits that either increase or increase cost and complexity associated with spacecraft operations. These examples address factors such as overall design performance margins, levels of redundancy, degree of automated failure response, type and quantity of command and telemetry interfaces, and the definition of reference scenarios for analysis and test. Each example - from early program requirements, design implementation and resulting real-time operations experience - to tell the end-to-end "story" Based on these experiences, specific techniques are recommended to enable earlier and more effective assessment of operations concerns during the design process. A formal method for the assessment of spacecraft operability is defined and results of such operability assessments for recent spacecraft designs are provided. Recent experience in applying these techniques to Orion spacecraft development is reviewed to highlight the direct benefits of early operational assessment and collaborative development efforts.

Enhanced Formation Flying for the Earth Observing-1 (EO-1) New Millennium Mission

NASA Technical Reports Server (NTRS)

Folta, David; Quinn, David

1997-01-01

With scientific objectives for Earth observation programs becoming more ambitious and spacecraft becoming more autonomous, the need for new technical approaches on the feasibility of achieving and maintaining formations of spacecraft has come to the forefront. The trend to develop small low cost spacecraft has led many scientists to recognize the advantage of flying several spacecraft in formation, an example of which is shown in the figure below, to achieve the correlated instrument measurements formerly possible only by flying many instruments on a single large platform. Yet, formation flying imposes additional complications on orbit maintenance, especially when each spacecraft has its own orbit requirements. However, advances in automation proposed by GSFC Codes 550 and 712 allow more of the burden in maneuver planning and execution to be placed onboard the spacecraft, mitigating some of the associated operational concerns. The purpose of this analysis is to develop the fundamentals of formation flying mechanics, concepts for understanding the relative motion of free flying spacecraft, and an operational control theory for formation maintenance of the Earth Observing-1 (EO-l) spacecraft that is part of the New Millennium. Results of this development can be used to determine the appropriateness of formation flying for a particular case as well as the operational impacts. Applications to the Mission to Planet Earth (MTPE) Earth Observing System (EOS) and New Millennium (NM) were highly considered in analysis and applications. This paper presents the proposed methods for the guidance and control of the EO-1 spacecraft to formation fly with the Landsat-7 spacecraft using an autonomous closed loop three axis navigation control, GPS, and Cross link navigation support. Simulation results using various fidelity levels of modeling, algorithms developed and implemented in MATLAB, and autonomous 'fuzzy logic' control using AutoCon will be presented. The results of these analysis on the ability to meet mission and formation flying requirements will be presented.

Operation and Development Status of the Spacecraft Fire Experiments (Saffire)

NASA Technical Reports Server (NTRS)

Ruff, Gary A.; Urban, David L.

2016-01-01

Since 2012, a series of Spacecraft Fire Experiments (Saffire) have been under development by the Spacecraft Fire Safety Demonstration (SFS Demo) project, funded by NASA's Advanced Exploration Systems Division. The overall objective of this project is to reduce the uncertainty and risk associated with the design of spacecraft fire safety systems for NASA's exploration missions. The approach to achieving this goal has been to define, develop, and conduct experiments that address gaps in spacecraft fire safety knowledge and capabilities identified by NASA's Fire Safety System Maturation Team. The Spacecraft Fire Experiments (Saffire-I, -II, and -III) are material flammability tests at length scales that are realistic for a spacecraft fire in low-gravity. The specific objectives of these three experiments are to (1) determine how rapidly a large scale fire grows in low-gravity and (2) investigate the low-g flammability limits compared to those obtained in NASA's normal gravity material flammability screening test. The experiments will be conducted in Orbital ATK's Cygnus vehicle after it has unberthed from the International Space Station. The tests will be fully automated with the data downlinked at the conclusion of the test before the Cygnus vehicle reenters the atmosphere. This paper discusses the status of the Saffire-I, II, and III experiments followed by a review of the fire safety technology gaps that are driving the development of objectives for the next series of experiments, Saffire-IV, V, and VI.

Implications of arcing due to spacecraft charging on spacecraft EMI margins of immunity

NASA Technical Reports Server (NTRS)

Inouye, G. T.

1981-01-01

Arcing due to spacecraft charging on spacecraft EMI margins of immunity was determined. The configuration of the P78-2 spacecraft of the SCATHA program was analyzed. A brushfire arc discharge model was developed, and a technique for initiating discharges with a spark plug trigger was for data configuration. A set of best estimate arc discharge parameters was defined. The effects of spacecraft potentials in limiting the discharge current blowout component are included. Arc discharge source models were incorporated into a SEMCAP EMI coupling analysis code for the DSP spacecraft. It is shown that with no mission critical circuits will be affected.

Three-Axis Time-Optimal Attitude Maneuvers of a Rigid-Body

NASA Astrophysics Data System (ADS)

Wang, Xijing; Li, Jisheng

With the development trends for modern satellites towards macro-scale and micro-scale, new demands are requested for its attitude adjustment. Precise pointing control and rapid maneuvering capabilities have long been part of many space missions. While the development of computer technology enables new optimal algorithms being used continuously, a powerful tool for solving problem is provided. Many papers about attitude adjustment have been published, the configurations of the spacecraft are considered rigid body with flexible parts or gyrostate-type systems. The object function always include minimum time or minimum fuel. During earlier satellite missions, the attitude acquisition was achieved by using the momentum ex change devices, performed by a sequential single-axis slewing strategy. Recently, the simultaneous three-axis minimum-time maneuver(reorientation) problems have been studied by many researchers. It is important to research the minimum-time maneuver of a rigid spacecraft within onboard power limits, because of potential space application such as surveying multiple targets in space and academic value. The minimum-time maneuver of a rigid spacecraft is a basic problem because the solutions for maneuvering flexible spacecraft are based on the solution to the rigid body slew problem. A new method for the open-loop solution for a rigid spacecraft maneuver is presented. Having neglected all perturbation torque, the necessary conditions of spacecraft from one state to another state can be determined. There is difference between single-axis with multi-axis. For single- axis analytical solution is possible and the switching line passing through the state-space origin belongs to parabolic. For multi-axis, it is impossible to get analytical solution due to the dynamic coupling between the axes and must be solved numerically. Proved by modern research, Euler axis rotations are quasi-time-optimal in general. On the basis of minimum value principles, a research for reorienting an inertial syrnmetric spacecraft with time cost function from an initial state of rest to a final state of rest is deduced. And the solution to it is stated below: Firstly, the essential condition for solving the problem is deduced with the minimum value principle. The necessary conditions for optimality yield a two point boundary-value problem (TPBVP), which, when solved, produces the control history that minimize time performance index. In the nonsingular control, the solution is the' bang-bang maneuver. The control profile is characterized by Saturated controls for the entire maneuver. The singular control maybe existed. It is only singular in mathematics. According to physical principle, the bigger the mode of the control torque is, the shorter the time is. So saturated controls are used in singular control. Secondly, the control parameters are always in maximum, so the key problem is to determine switch point thus original problem is changed to find the changing time. By the use of adjusting the switch on/off time, the genetic algorithm, which is a new robust method is optimized to determine the switch features without the gyroscopic coupling. There is improvement upon the traditional GA in this research. The homotopy method to find the nonlinear algebra is based on rigorous topology continuum theory. Based on the idea of the homotopy, the relaxation parameters are introduced, and the switch point is figured out with simulated annealing. Computer simulation results using a rigid body show that the new method is feasible and efficient. A practical method of computing approximate solutions to the time-optimal control- switch times for rigid body reorientation has been developed.

Design, Development, Testing, and Evaluation: Human Factors Engineering

NASA Technical Reports Server (NTRS)

Adelstein, Bernard; Hobbs, Alan; OHara, John; Null, Cynthia

2006-01-01

While human-system interaction occurs in all phases of system development and operation, this chapter on Human Factors in the DDT&E for Reliable Spacecraft Systems is restricted to the elements that involve "direct contact" with spacecraft systems. Such interactions will encompass all phases of human activity during the design, fabrication, testing, operation, and maintenance phases of the spacecraft lifespan. This section will therefore consider practices that would accommodate and promote effective, safe, reliable, and robust human interaction with spacecraft systems. By restricting this chapter to what the team terms "direct contact" with the spacecraft, "remote" factors not directly involved in the development and operation of the vehicle, such as management and organizational issues, have been purposely excluded. However, the design of vehicle elements that enable and promote ground control activities such as monitoring, feedback, correction and reversal (override) of on-board human and automation process are considered as per NPR8705.2A, Section 3.3.

Spacecraft Maneuvering at the Sun/Earth-Moon L2 Libration Point

NASA Astrophysics Data System (ADS)

Shahid, Kamran

Spacecraft formation flying in the vicinity of the Sun/Earth-Moon libration points offers many promising possibilities for space exploration. The concept of formation flying involves the distribution of the functionality of a single spacecraft among several smaller, cooperative spacecraft. The libration points are locations relative to two large orbiting bodies where a third body with relatively small mass can remain stationary relative to the two larger bodies. The most significant perturbation experienced by a spacecraft at the libration point is effect of solar radiation pressure. This thesis presents the development of nonlinear control techniques for maneuvering control at the Sun-Earth/Moon L2 libration point. A new thruster based formation control technique is presented. We also consider a leader/follower formation architecture, and examine the station keeping control of the leader spacecraft and the formation control of the follower spacecraft using solar radiation pressure. Reference trajectories of the leader spacecraft, halo and Lissajous orbits, are determined using a numerical technique in order to take into account all major gravitational perturbations. The nonlinear controllers are developed based on Lyapunov analysis, including non-adaptive and adaptive designs. Thruster based and solar radiation pressure based control laws for spacecraft maneuvering at the Sun-Earth/Moon libration point are developed. Higher order sliding mode control is utilized to address the non-affine structure of the solar sail control inputs. The reduced input solar radiation pressure problem is properly addressed as an underactuated control problem. The development of adaptive control for solar sail equipped spacecraft is an innovation and represents and advancement in solar sailing control technology. Controller performance is evaluated in a high fidelity ephemeris model to reflect a realistic simulated space environment. The numerical results demonstrate the effectiveness of the proposed control techniques for spacecraft maneuvering using solar radiation pressure at the L2 libration point. Stationkeeping accuracies of 50m and formation maintenance accuracies of less than 1m are possible using solar radiation pressure at a sub-L2 libration point. The benefits of these control techniques include increasing libration point mission lifetimes and doubling payload mass fractions as compared to conventional propulsion methods.

Development and flight history of SERT 2 spacecraft

NASA Technical Reports Server (NTRS)

Kerslake, William R.; Ignaczak, Louis R.

1992-01-01

A 25-year historical review of the Space Electric Rocket Test 2 (SERT 2) mission is presented. The Agena launch vehicle; the SERT 2 spacecraft; and mission-peculiar spacecraft hardware, including two ion thruster systems, are described. The 3 1/2-year development period, from 1966 to 1970, that was needed to design, fabricate, and qualify the ion thruster system and the supporting spacecraft components, is documented. Major testing of two ion thruster systems and related auxiliary experiments that were conducted in space after the 3 Feb. 1970, launch are reviewed. Extended ion thruster restarts from 1973 to 1981 are reported, in addition to cross-neutralization tests. Tests of a reflector erosion experiment were continued in 1989 to 1991. The continuing performance of spacecraft subsystems, including the solar arrays, over the 1970-1991 period is summarized. Finally, the knowledge of thruster-spacecraft interactions learned from SERT 2 is listed.

NASA's Space Environments and Effects Program: Technology for the New Millennium

NASA Technical Reports Server (NTRS)

Hardage, Donna M.; Pearson, Steven D.

2000-01-01

Current trends in spacecraft development include the use of advanced technologies while maintaining the "faster, better, cheaper" philosophy. Spacecraft designers are continually designing with smaller and faster electronics as well as lighter and thinner materials providing better performance, lower weight, and ultimately lower costs. Given this technology trend, spacecraft will become increasingly susceptible to the harsh space environments, causing damaging or even disabling effects on space systems. NASA's Space Environments and Effects (SEE) Program defines the space environments and provides advanced technology development to support the design, development, and operation of spacecraft systems that will accommodate or mitigate effects due to the harsh space environments. This Program provides a comprehensive and focused approach to understanding the space environment, to define the best techniques for both flight and ground-based experimentation, to update the models which predict both the environments and the environmental effects on spacecraft, and finally to ensure that this multitudinous information is properly maintained and inserted into spacecraft design programs. A description of the SEE Program, its accomplishments, and future activities is provided.

Combining Acceleration and Displacement Dependent Modal Frequency Responses Using an MSC/NASTRAN DMAP Alter

NASA Technical Reports Server (NTRS)

Barnett, Alan R.; Widrick, Timothy W.; Ludwiczak, Damian R.

1996-01-01

Solving for dynamic responses of free-free launch vehicle/spacecraft systems acted upon by buffeting winds is commonly performed throughout the aerospace industry. Due to the unpredictable nature of this wind loading event, these problems are typically solved using frequency response random analysis techniques. To generate dynamic responses for spacecraft with statically-indeterminate interfaces, spacecraft contractors prefer to develop models which have response transformation matrices developed for mode acceleration data recovery. This method transforms spacecraft boundary accelerations and displacements into internal responses. Unfortunately, standard MSC/NASTRAN modal frequency response solution sequences cannot be used to combine acceleration- and displacement-dependent responses required for spacecraft mode acceleration data recovery. External user-written computer codes can be used with MSC/NASTRAN output to perform such combinations, but these methods can be labor and computer resource intensive. Taking advantage of the analytical and computer resource efficiencies inherent within MS C/NASTRAN, a DMAP Alter has been developed to combine acceleration- and displacement-dependent modal frequency responses for performing spacecraft mode acceleration data recovery. The Alter has been used successfully to efficiently solve a common aerospace buffeting wind analysis.

Tools Automate Spacecraft Testing, Operation

NASA Technical Reports Server (NTRS)

2010-01-01

"NASA began the Small Explorer (SMEX) program to develop spacecraft to advance astrophysics and space physics. As one of the entities supporting software development at Goddard Space Flight Center, the Hammers Company Inc. (tHC Inc.), of Greenbelt, Maryland, developed the Integrated Test and Operations System to support SMEX. Later, the company received additional Small Business Innovation Research (SBIR) funding from Goddard for a tool to facilitate the development of flight software called VirtualSat. NASA uses the tools to support 15 satellites, and the aerospace industry is using them to develop science instruments, spacecraft computer systems, and navigation and control software."

Tactical Satellite-3 Mission Overview and Initial Lessons Learned (Postprint)

DTIC Science & Technology

2013-03-01

current buses. The spacecraft bus includes the main structure; attitude control system (reaction wheels and torque rods); the thermal protection...Specific key areas are the relatively rapid checkout of the spacecraft and lessons from the responsive space development. 15. SUBJECT TERMS...relatively rapid checkout of the spacecraft and lessons from the responsive space development. INTRODUCTION The Tactical Satellite 3 mission was a

A charging model for three-axis stabilized spacecraft

NASA Technical Reports Server (NTRS)

Massaro, M. J.; Green, T.; Ling, D.

1977-01-01

A charging model was developed for geosynchronous, three-axis stabilized spacecraft when under the influence of a geomagnetic substorm. The differential charging potentials between the thermally coated or blanketed outer surfaces and metallic structure of a spacecraft were determined when the spacecraft was immersed in a dense plasma cloud of energetic particles. The spacecraft-to-environment interaction was determined by representing the charged particle environment by equivalent current source forcing functions and by representing the spacecraft by its electrically equivalent circuit with respect to the plasma charging phenomenon. The charging model included a sun/earth/spacecraft orbit model that simulated the sum illumination conditions of the spacecraft outer surfaces throughout the orbital flight on a diurnal as well as a seasonal basis. Transient and steady-state numerical results for a three-axis stabilized spacecraft are presented.

A thermal scale modeling study for Apollo and Apollo applications, volume 1

NASA Technical Reports Server (NTRS)

Shannon, R. L.

1972-01-01

The program is reported for developing and demonstrating the capabilities of thermal scale modeling as a thermal design and verification tool for Apollo and Apollo Applications Projects. The work performed for thermal scale modeling of STB; cabin atmosphere/spacecraft cabin wall thermal interface; closed loop heat rejection radiator; and docked module/spacecraft thermal interface are discussed along with the test facility requirements for thermal scale model testing of AAP spacecraft. It is concluded that thermal scale modeling can be used as an effective thermal design and verification tool to provide data early in a spacecraft development program.

Influence of Natural Environments in Spacecraft Design, Development, and Operation

NASA Technical Reports Server (NTRS)

Edwards, Dave

2013-01-01

Spacecraft are growing in complexity and sensitivity to environmental effects. The spacecraft engineer must understand and take these effects into account in building reliable, survivable, and affordable spacecraft. Too much protections, however, means unnecessary expense while too little will potentially lead to early mission loss. The ability to balance cost and risk necessitates an understanding of how the environment impacts the spacecraft and is a critical factor in its design. This presentation is intended to address both the space environment and its effects with the intent of introducing the influence of the environment on spacecraft performance.

Influence of Natural Environments in Spacecraft Design, Development, and Operation

NASA Technical Reports Server (NTRS)

Edwards, Dave

2012-01-01

Spacecraft are growing in complexity and sensitivity to environmental effects. The spacecraft engineer must understand and take these effects into account in building reliable, survivable, and affordable spacecraft. Too much protections, however, means unnecessary expense while too little will potentially lead to early mission loss. The ability to balance cost and risk necessitates an understanding of how the environment impacts the spacecraft and is a critical factor in its design. This presentation is intended to address both the space environment and its effects with the intent of introducing the influence of the environment on spacecraft performance.

Autonomous Spacecraft Communication Interface for Load Planning

NASA Technical Reports Server (NTRS)

Dever, Timothy P.; May, Ryan D.; Morris, Paul H.

2014-01-01

Ground-based controllers can remain in continuous communication with spacecraft in low Earth orbit (LEO) with near-instantaneous communication speeds. This permits near real-time control of all of the core spacecraft systems by ground personnel. However, as NASA missions move beyond LEO, light-time communication delay issues, such as time lag and low bandwidth, will prohibit this type of operation. As missions become more distant, autonomous control of manned spacecraft will be required. The focus of this paper is the power subsystem. For present missions, controllers on the ground develop a complete schedule of power usage for all spacecraft components. This paper presents work currently underway at NASA to develop an architecture for an autonomous spacecraft, and focuses on the development of communication between the Mission Manager and the Autonomous Power Controller. These two systems must work together in order to plan future load use and respond to unanticipated plan deviations. Using a nominal spacecraft architecture and prototype versions of these two key components, a number of simulations are run under a variety of operational conditions, enabling development of content and format of the messages necessary to achieve the desired goals. The goals include negotiation of a load schedule that meets the global requirements (contained in the Mission Manager) and local power system requirements (contained in the Autonomous Power Controller), and communication of off-plan disturbances that arise while executing a negotiated plan. The message content is developed in two steps: first, a set of rapid-prototyping "paper" simulations are preformed; then the resultant optimized messages are codified for computer communication for use in automated testing.

The NASA Mission Operations and Control Architecture Program

NASA Technical Reports Server (NTRS)

Ondrus, Paul J.; Carper, Richard D.; Jeffries, Alan J.

1994-01-01

The conflict between increases in space mission complexity and rapidly declining space mission budgets has created strong pressures to radically reduce the costs of designing and operating spacecraft. A key approach to achieving such reductions is through reducing the development and operations costs of the supporting mission operations systems. One of the efforts which the Communications and Data Systems Division at NASA Headquarters is using to meet this challenge is the Mission Operations Control Architecture (MOCA) project. Technical direction of this effort has been delegated to the Mission Operations Division (MOD) of the Goddard Space Flight Center (GSFC). MOCA is to develop a mission control and data acquisition architecture, and supporting standards, to guide the development of future spacecraft and mission control facilities at GSFC. The architecture will reduce the need for around-the-clock operations staffing, obtain a high level of reuse of flight and ground software elements from mission to mission, and increase overall system flexibility by enabling the migration of appropriate functions from the ground to the spacecraft. The end results are to be an established way of designing the spacecraft-ground system interface for GSFC's in-house developed spacecraft, and a specification of the end to end spacecraft control process, including data structures, interfaces, and protocols, suitable for inclusion in solicitation documents for future flight spacecraft. A flight software kernel may be developed and maintained in a condition that it can be offered as Government Furnished Equipment in solicitations. This paper describes the MOCA project, its current status, and the results to date.

Nickel cadmium battery operations on-orbit: Trials, tribulations, and success on the Upper Atmosphere Research Satellite

NASA Technical Reports Server (NTRS)

Rao, Gopalakrishna M.; Miller, Scott D.

1994-01-01

The Upper Atmosphere Research Satellite (UARS), designed, built, integrated, tested, and operated by NASA and Martin Marietta is a low-Earth orbiting, Earth-observing spacecraft which was launched via Space Shuttle Discovery on September 12, 1991 and deployed three days later. The Modular Power Subsystem (MPS) onboard the satellite is equipped with three NASA Standard 50 Ampere-hour (Ah) nickel-cadmium (NiCd) batteries. McDonnell Douglas Electronics Systems Company fabricated the MPS, and batteries from Gates Aerospace Batteries cells. Nominal battery performance was achieved for the first four months of spacecraft operation. First evidence of anomalous battery performance was observed in January 1992, after the first maximum beta angle (low Depth of Discharge) period. Since then, the Flight Operations Team (FOT), under the direction of Goddard Space Flight Center's UARS Project and Space Power Application Branch, has monitored and managed battery performance by adjusting solar array offset angle, conducting periodic deep discharge, and controlling battery recharge ratio. This paper covers a brief overview of the UARS, the FOT's operational battery management, and the observed spacecraft battery performance.

Electrolysis Propulsion Provides High-Performance, Inexpensive, Clean Spacecraft Propulsion

NASA Technical Reports Server (NTRS)

deGroot, Wim A.

1999-01-01

An electrolysis propulsion system consumes electrical energy to decompose water into hydrogen and oxygen. These gases are stored in separate tanks and used when needed in gaseous bipropellant thrusters for spacecraft propulsion. The propellant and combustion products are clean and nontoxic. As a result, costs associated with testing, handling, and launching can be an order of magnitude lower than for conventional propulsion systems, making electrolysis a cost-effective alternative to state-of-the-art systems. The electrical conversion efficiency is high (>85 percent), and maximum thrust-to-power ratios of 0.2 newtons per kilowatt (N/kW), a 370-sec specific impulse, can be obtained. A further advantage of the water rocket is its dual-mode potential. For relatively high thrust applications, the system can be used as a bipropellant engine. For low thrust levels and/or small impulse bit requirements, cold gas oxygen can be used alone. An added innovation is that the same hardware, with modest modifications, can be converted into an energy-storage and power-generation fuel cell, reducing the spacecraft power and propulsion system weight by an order of magnitude.

A graphical approach to electric sail mission design with radial thrust

NASA Astrophysics Data System (ADS)

Mengali, Giovanni; Quarta, Alessandro A.; Aliasi, Generoso

2013-02-01

This paper describes a semi-analytical approach to electric sail mission analysis under the assumption that the spacecraft experiences a purely radial, outward, propulsive acceleration. The problem is tackled by means of the potential well concept, a very effective idea that was originally introduced by Prussing and Coverstone in 1998. Unlike a classical procedure that requires the numerical integration of the equations of motion, the proposed method provides an estimate of the main spacecraft trajectory parameters, as its maximum and minimum attainable distance from the Sun, with the simple use of analytical relationships and elementary graphs. A number of mission scenarios clearly show the effectiveness of the proposed approach. In particular, when the spacecraft parking orbit is either circular or elliptic it is possible to find the optimal performances required to reach an escape condition or a given distance from the Sun. Another example is given by the optimal strategy required to reach a heliocentric Keplerian orbit of prescribed orbital period. Finally the graphical approach is applied to the preliminary design of a nodal mission towards a Near Earth Asteroid.

Determination of Survivable Fires

NASA Technical Reports Server (NTRS)

Dietrich, D. L.; Niehaus, J. E.; Ruff, G. A.; Urban, D. L.; Takahashi, F.; Easton, J. W.; Abbott, A. A.; Graf, J. C.

2012-01-01

At NASA, there exists no standardized design or testing protocol for spacecraft fire suppression systems (either handheld or total flooding designs). An extinguisher's efficacy in safely suppressing any reasonable or conceivable fire is the primary benchmark. That concept, however, leads to the question of what a reasonable or conceivable fire is. While there exists the temptation to over-size' the fire extinguisher, weight and volume considerations on spacecraft will always (justifiably) push for the minimum size extinguisher required. This paper attempts to address the question of extinguisher size by examining how large a fire a crew member could successfully survive and extinguish in the confines of a spacecraft. The hazards to the crew and equipment during an accidental fire include excessive pressure rise resulting in a catastrophic rupture of the vehicle skin, excessive temperatures that burn or incapacitate the crew (due to hyperthermia), carbon dioxide build-up or other accumulation of other combustion products (e.g. carbon monoxide). Estimates of these quantities are determined as a function of fire size and mass of material burned. This then becomes the basis for determining the maximum size of a target fire for future fire extinguisher testing.

Capture of near-Earth objects with low-thrust propulsion and invariant manifolds

NASA Astrophysics Data System (ADS)

Tang, Gao; Jiang, Fanghua

2016-01-01

In this paper, a mission incorporating low-thrust propulsion and invariant manifolds to capture near-Earth objects (NEOs) is investigated. The initial condition has the spacecraft rendezvousing with the NEO. The mission terminates once it is inserted into a libration point orbit (LPO). The spacecraft takes advantage of stable invariant manifolds for low-energy ballistic capture. Low-thrust propulsion is employed to retrieve the joint spacecraft-asteroid system. Global optimization methods are proposed for the preliminary design. Local direct and indirect methods are applied to optimize the two-impulse transfers. Indirect methods are implemented to optimize the low-thrust trajectory and estimate the largest retrievable mass. To overcome the difficulty that arises from bang-bang control, a homotopic approach is applied to find an approximate solution. By detecting the switching moments of the bang-bang control the efficiency and accuracy of numerical integration are guaranteed. By using the homotopic approach as the initial guess the shooting function is easy to solve. The relationship between the maximum thrust and the retrieval mass is investigated. We find that both numerically and theoretically a larger thrust is preferred.

Navigation Challenges in the MAVEN Science Phase

NASA Technical Reports Server (NTRS)

Demcak, Stuart; Young, Brian; Lam, Try; Trawny, Nikolas; Lee, Clifford; Anderson, Rodney; Broschart, Stephen; Ballard, Christopher; Folta, David C.

2012-01-01

The MAVEN spacecraft will explore Mars' upper atmosphere. The primary science phase will last one (Earth) year, during which the spacecraft will be in an elliptical 4.5 hour orbit at an inclination of 75 degrees. The 75 degree inclination results in the orbit periapsis oscillating between +/-75 degrees latitude, thus naturally covering most Mars latitudes during the primary mission. The orbit will be controlled via maneuvers so that the maximum orbit density remains in a density corridor. This results in the MAVEN science phase being in a light aerobraking type orbit of around 160 km for an extended period. In addition, the mission has significantly less tracking data than aerobraking phases of other missions, and even less than other NASA Mars orbiter primary phases. This results in significant challenges for the Navigation Team. They can be summarized as a difficulty in determining the current density profile, which maps into degraded trajectory predictions and less accurate control over the spacecraft location in the targeted density corridor via maneuvers. This paper describes these challenges and the Navigation Team's plans to meet them.

Does a Local B-Minimum Appear in the Tail Current Sheet During a Substorm Growth Phase?

NASA Astrophysics Data System (ADS)

Sergeev, V. A.; Gordeev, E. I.; Merkin, V. G.; Sitnov, M. I.

2018-03-01

Magnetic configurations with dBz/dr > 0 in the midtail current sheet are potentially unstable to various instabilities associated with the explosive substorm onset. Their existence is hard to confirm with observations of magnetospheric spacecraft. Here we use remote sensing by low-altitude spacecraft that measured the loss cone filling rate during electron-rich solar particle event, providing information about magnetic properties of the tail current sheet. We found a latitudinally localized anisotropic 30 keV electron loss cone region embedded inside an extended region of isotropic solar electron precipitation. It was persistently observed for more than 0.5 h during isolated growth phase event by six Polar Operational Environmental Satellites spacecraft, which crossed the premidnight auroral oval. The embedded anisotropic region was observed 1° poleward of the outer radiation belt boundary over 4-5 h wide magnetic local time sector, suggesting a persistent ridge-type Bz2/j maximum in the equatorial plasma sheet at distances 15-20 RE. We discuss infrequent observations of such events taking into account recent results of global magnetohydrodynamic simulations.

Analysis of the relative movement between mirrors and detectors for the next generation x-ray telescopes

NASA Astrophysics Data System (ADS)

Civitani, Marta

2009-08-01

Focusing X-ray telescopes with imaging capabilities, like SIMBOL-X, HEXISAT and IXO, are characterized by very long focal lengths, greater than 10m. The constraints posed by the launchers on the maximum dimensions of a payload, make necessary using alternatives to monolithic telescopes. One possibility is that the mirror and the detectors are carried by two separate spacecrafts that fly in formation. Another is placing the detector module on a bench that will be extended once in final orbit. In both the case the system will be subjected to deformation due the relative movement of the mirrors with respect to detectors. In one case the deformation will be due to the correction on the position and attitude of the detector spacecraft to maintain the formation with the mirror spacecraft, while in the other to oscillations of the detectors on the top of the bench. The aim of this work is to compare the behavior of the system in the two different configurations and to evaluate the performances of the on board metrology systems needed not to degrade the telescope angular resolution.

Astronaut Alan Shepard is rescued by helicopter at end of MR-3 flight

NASA Image and Video Library

1961-05-05

S88-31376 (5 May 1961) --- Astronaut Alan B. Shepard Jr., pilot of the Mercury-Redstone 3 (MR-3) suborbital spaceflight, is retrieved by a helicopter from the USS Lake Champlain during recovery operations in the western Atlantic Ocean. Shepard and the Mercury spacecraft designated the ?Freedom 7? (floating in water below) were flown to the deck of the recovery ship within 11 minutes of splashdown. MR-3 was the United States? first manned space mission. The spacecraft attained a maximum speed of 5,180 miles per hour, reached an altitude of 116 1/2 statute miles, and landed 302 statute miles downrange from Cape Canaveral, Florida. The suborbital mission lasted 15 minutes and 22 seconds. Photo credit: NASA or National Aeronautics and Space Administration

The effect of solar array degradation on electric propulsion spacecraft performance.

NASA Technical Reports Server (NTRS)

Sauer, C. G., Jr.; Bourke, R. D.

1972-01-01

Current estimates of solar-electric-propulsion spacecraft performance are based upon a solar-array output power which is degraded by approximately 10-13% to account for possible losses caused by proton, electron and micrometeorite damage. Past studies have used a worst case analysis in which the maximum degradation was taken to occur at the beginning of the mission. This paper presents a comparison of mission studies using a hypothetical exponential decrease in power with time, with those using a sudden degradation of solar power. These comparisons indicate that the performance gain by using a time varying degradation during the mission is quite small for outbound missions in the solar system. In addition an indication of the power allocation strategy to be followed during a mission is presented.

Relative Frequency Distribution of D125 C Values for Spore Isolates from the Mariner-Mars 1969 Spacecraft

PubMed Central

Bond, W. W.; Favero, M. S.; Petersen, N. J.; Marshall, J. H.

1971-01-01

Bacterial spore crops were prepared from 103 randomly selected aerobic mesophilic isolates collected during a spore assay of Mariner-Mars 1969 spacecraft conducted by the Jet Propulsion Laboratory. D125 c values, which were determined by the fractional-replicate-unit-negative-most-probable number assay method using a forced air oven, ranged from less than 5 min to a maximum of 58 min. Subsequent identification of the 103 isolates indicated that there was no relationship between species and dry-heat resistance. A theoretical dry-heat survival curve of the “population” was nonlinear. The slope of this curve was determined almost exclusively by the more resistant organisms, although they represented only a small portion of the “population.” PMID:16349904

Applications technology satellites battery and power system design

NASA Technical Reports Server (NTRS)

Ford, F. E.; Bemis, B.

1977-01-01

A summary of the ATS battery design which is onboard the Applications Technology Satellite (ATS) is provided. The 15 ampere hour nickel cadmium cells were manufactured by Gulton, 19 series connected cells per battery, and there are two batteries in each spacecraft. The operating design life was two years in a synchronous orbit, and a maximum depth of discharge of 50 percent. The design temperature for the batteries in the spacecraft was 0 to 25 C, and the charge control consisted of 1 volt versus temperature on a constant percentage voltage. Also, C/10 current limit, and a commandable trickle charge rate, using C/20 or C/60. The undervoltage was sent across a 9 cell and a 10 cell group, and it was set at one volt average per group on either group.

KSC-2012-6105

NASA Image and Video Library

2012-11-01

CAPE CANAVERAL, Fla. – The Orion Exploration Flight Test 1 crew module is undergoing proof pressure testing at the Operations and Checkout Building at NASA's Kennedy Space Center in Florida. The test incrementally pressurizes the spacecraft with breathing air and is designed to demonstrate weld strength capability and structural performance at maximum flight operating pressures. Orion is the exploration spacecraft designed to carry crews to space beyond low Earth orbit. It will provide emergency abort capability, sustain the crew during the space travel and provide safe re-entry from deep space return velocities. The first unpiloted test flight of the Orion is scheduled to launch in 2014 atop a Delta IV rocket and in 2017 on a Space Launch System rocket. For more information, visit http://www.nasa.gov/orion Photo credit: NASA/Ben Smegelsky

KSC-2012-6103

NASA Image and Video Library

2012-11-01

CAPE CANAVERAL, Fla. – The Orion Exploration Flight Test 1 crew module is undergoing proof pressure testing at the Operations and Checkout Building at NASA's Kennedy Space Center in Florida. The test incrementally pressurizes the spacecraft with breathing air and is designed to demonstrate weld strength capability and structural performance at maximum flight operating pressures. Orion is the exploration spacecraft designed to carry crews to space beyond low Earth orbit. It will provide emergency abort capability, sustain the crew during the space travel and provide safe re-entry from deep space return velocities. The first unpiloted test flight of the Orion is scheduled to launch in 2014 atop a Delta IV rocket and in 2017 on a Space Launch System rocket. For more information, visit http://www.nasa.gov/orion Photo credit: NASA/Ben Smegelsky

KSC-2012-6104

NASA Image and Video Library

2012-11-01

CAPE CANAVERAL, Fla. – The Orion Exploration Flight Test 1 crew module is undergoing proof pressure testing at the Operations and Checkout Building at NASA's Kennedy Space Center in Florida. The test incrementally pressurizes the spacecraft with breathing air and is designed to demonstrate weld strength capability and structural performance at maximum flight operating pressures. Orion is the exploration spacecraft designed to carry crews to space beyond low Earth orbit. It will provide emergency abort capability, sustain the crew during the space travel and provide safe re-entry from deep space return velocities. The first unpiloted test flight of the Orion is scheduled to launch in 2014 atop a Delta IV rocket and in 2017 on a Space Launch System rocket. For more information, visit http://www.nasa.gov/orion Photo credit: NASA/Ben Smegelsky

AE 6: A model environment of trapped electrons for solar maximum

NASA Technical Reports Server (NTRS)

Teague, M. J.; Chan, K. W.; Vette, J. I.

1976-01-01

A projected inner zone electron model environment, AE 6, for the epoch 1980 is presented. It is intended to provide estimates of mission fluxes that spacecraft will encounter in the coming solar maximum years. AE 6 is presented by graphs of omnidirectional integral flux as a function of L shell, the ambient magnetic field B, and the energy E. Results of orbital integrations for altitudes from 150 n.m. to 18,000 n.m. are given for circular orbits with four different inclinations, using the AE 6 and the AE 4 solar maximum models for the inner and outer zones, respectively. The derivation of AE 6 is described, and a brief comparison is given of the radial profiles of equatorial fluxes from several related models. A short summary of the associated computer programs is included.

Fire resistant resilient foams. [for seat cushions

NASA Technical Reports Server (NTRS)

Gagliani, J.

1976-01-01

Primary program objectives were the formulation, screening, optimization and characterization of open-cell, fire resistant, low-smoke emitting, thermally stable, resilient polyimide foams suitable for seat cushions in commercial aircraft and spacecraft. Secondary program objectives were to obtain maximum improvement of the tension, elongation and tear characteristics of the foams, while maintaining the resiliency, thermal stability, low smoke emission and other desirable attributes of these materials.

Applications technology satellite F&G /ATS F&G/ mobile terminal.

NASA Technical Reports Server (NTRS)

Greenbaum, L. A.; Baker, J. L.

1971-01-01

The mobile terminal is a flexible, easily transportable system. The terminal design incorporates a combination of unique and proven hardware to provide maximum utility consistent with reliability. The flexibility built into the system will make it possible to satisfy the requirements of the applications technology satellite program concerned with the conduction of various spacecraft technology experiments. The terminal includes two parabolic antennas.

System Design under Uncertainty: Evolutionary Optimization of the Gravity Probe-B Spacecraft

NASA Technical Reports Server (NTRS)

Pullen, Samuel P.; Parkinson, Bradford W.

1994-01-01

This paper discusses the application of evolutionary random-search algorithms (Simulated Annealing and Genetic Algorithms) to the problem of spacecraft design under performance uncertainty. Traditionally, spacecraft performance uncertainty has been measured by reliability. Published algorithms for reliability optimization are seldom used in practice because they oversimplify reality. The algorithm developed here uses random-search optimization to allow us to model the problem more realistically. Monte Carlo simulations are used to evaluate the objective function for each trial design solution. These methods have been applied to the Gravity Probe-B (GP-B) spacecraft being developed at Stanford University for launch in 1999, Results of the algorithm developed here for GP-13 are shown, and their implications for design optimization by evolutionary algorithms are discussed.

A geometric model of a V-slit Sun sensor correcting for spacecraft wobble

NASA Technical Reports Server (NTRS)

Mcmartin, W. P.; Gambhir, S. S.

1994-01-01

A V-Slit sun sensor is body-mounted on a spin-stabilized spacecraft. During injection from a parking or transfer orbit to some final orbit, the spacecraft may not be dynamically balanced. This may result in wobble about the spacecraft spin axis as the spin axis may not be aligned with the spacecraft's axis of symmetry. While the widely used models in Spacecraft Attitude Determination and Control, edited by Wertz, correct for separation, elevation, and azimuthal mounting biases, spacecraft wobble is not taken into consideration. A geometric approach is used to develop a method for measurement of the sun angle which corrects for the magnitude and phase of spacecraft wobble. The algorithm was implemented using a set of standard mathematical routines for spherical geometry on a unit sphere.

ANTS: Applying A New Paradigm for Lunar and Planetary Exploration

NASA Technical Reports Server (NTRS)

Clark, P. E.; Curtis, S. A.; Rilee, M. L.

2002-01-01

ANTS (Autonomous Nano- Technology Swarm), a mission architecture consisting of a large (1000 member) swarm of picoclass (1 kg) totally autonomous spacecraft with both adaptable and evolvable heuristic systems, is being developed as a NASA advanced mission concept, and is here examined as a paradigm for lunar surface exploration. As the capacity and complexity of hardware and software, demands for bandwidth, and the sophistication of goals for lunar and planetary exploration have increased, greater cost constraints have led to fewer resources and thus, the need to operate spacecraft with less frequent human contact. At present, autonomous operation of spacecraft systems allows great capability of spacecraft to 'safe' themselves and survive when conditions threaten spacecraft safety. To further develop spacecraft capability, NASA is at the forefront of development of new mission architectures which involve the use of Intelligent Software Agents (ISAs), performing experiments in space and on the ground to advance deliberative and collaborative autonomous control techniques. Selected missions in current planning stages require small groups of spacecraft weighing tens, instead of hundreds, of kilograms to cooperate at a tactical level to select and schedule measurements to be made by appropriate instruments onboard. Such missions will be characterizing rapidly unfolding real-time events on a routine basis. The next level of development, which we are considering here, is in the use of autonomous systems at the strategic level, to explore the remote terranes, potentially involving large surveys or detailed reconnaissance.

Onboard processor technology review

NASA Technical Reports Server (NTRS)

Benz, Harry F.

1990-01-01

The general need and requirements for the onboard embedded processors necessary to control and manipulate data in spacecraft systems are discussed. The current known requirements are reviewed from a user perspective, based on current practices in the spacecraft development process. The current capabilities of available processor technologies are then discussed, and these are projected to the generation of spacecraft computers currently under identified, funded development. An appraisal is provided for the current national developmental effort.

Industry perspectives on Plug-& -Play Spacecraft Avionics

NASA Astrophysics Data System (ADS)

Franck, R.; Graven, P.; Liptak, L.

This paper describes the methodologies and findings from an industry survey of awareness and utility of Spacecraft Plug-& -Play Avionics (SPA). The survey was conducted via interviews, in-person and teleconference, with spacecraft prime contractors and suppliers. It focuses primarily on AFRL's SPA technology development activities but also explores the broader applicability and utility of Plug-& -Play (PnP) architectures for spacecraft. Interviews include large and small suppliers as well as large and small spacecraft prime contractors. Through these “ product marketing” interviews, awareness and attitudes can be assessed, key technical and market barriers can be identified, and opportunities for improvement can be uncovered. Although this effort focuses on a high-level assessment, similar processes can be used to develop business cases and economic models which may be necessary to support investment decisions.

Magnet-Based System for Docking of Miniature Spacecraft

NASA Technical Reports Server (NTRS)

Howard, Nathan; Nguyen, Hai D.

2007-01-01

A prototype system for docking a miniature spacecraft with a larger spacecraft has been developed by engineers at the Johnson Space Center. Engineers working on Mini AERCam, a free-flying robotic camera, needed to find a way to successfully dock and undock their miniature spacecraft to refuel the propulsion and recharge the batteries. The subsystems developed (see figure) include (1) a docking port, designed for the larger spacecraft, which contains an electromagnet, a ball lock mechanism, and a service probe; and (2) a docking cluster, designed for the smaller spacecraft, which contains either a permanent magnet or an electromagnet. A typical docking operation begins with the docking spacecraft maneuvering into position near the docking port on the parent vehicle. The electromagnet( s) are then turned on, and, if necessary, the docking spacecraft is then maneuvered within the capture envelope of the docking port. The capture envelope for this system is approximated by a 5-in. (12.7-cm) cube centered on the front of the docking-port electromagnet and within an angular misalignment of <30 . Thereafter, the magnetic forces draw the smaller spacecraft toward the larger one and this brings the spacecraft into approximate alignment prior to contact. Mechanical alignment guides provide the final rotational alignment into one of 12 positions. Once the docking vehicle has been captured magnetically in the docking port, the ball-lock mechanism is activated, which locks the two spacecraft together. At this point the electromagnet( s) are turned off, and the service probe extended if recharge and refueling are to be performed. Additionally, during undocking, the polarity of one electromagnet can be reversed to provide a gentle push to separate the two spacecraft. This system is currently being incorporated into the design of Mini AERCam vehicle.

The onset of the solar active cycle 22

NASA Technical Reports Server (NTRS)

Ahluwalia, H. S.

1989-01-01

There is a great deal of interest in being able to predict the main characteristics of a solar activity cycle (SAC). One would like to know, for instance, how large the amplitude (R sub m) of a cycle is likely to be, i.e., the annual mean of the sunspot numbers at the maximum of SAC. Also, how long a cycle is likely to last, i.e., its period. It would also be interesting to be able to predict the details, like how steep the ascending phase of a cycle is likely to be. Questions like these are of practical importance to NASA in planning the launch schedule for the low altitude, expensive spacecrafts like the Hubble Space Telescope, the Space Station, etc. Also, one has to choose a proper orbit, so that once launched the threat of an atmospheric drag on the spacecraft is properly taken into account. Cosmic ray data seem to indicate that solar activity cycle 22 will surpass SAC 21 in activity. The value of R sub m for SAC 22 may approach that of SAC 19. It would be interesting to see whether this prediction is borne out. Researchers are greatly encouraged to proceed with the development of a comprehensive prediction model which includes information provided by cosmic ray data.

Post-flight BET products for the 2nd discovery entry, STS-19 (51-A)

NASA Technical Reports Server (NTRS)

Kelly, G. M.; Mcconnell, J. G.; Heck, M. L.; Troutman, P. A.; Waters, L. A.; Findlay, J. T.

1985-01-01

The post-flight products for the second Discovery flight, STS-19 (51-A), are summarized. The inertial best estimate trajectory (BET), BT19D19/UN=169750N, was developed using spacecraft dynamic measurements from Inertial Measurement Unit 2 (IMU2) in conjunction with the best tracking coverage available for any of the earlier Shuttle entries. As a consequence of the latter, an anchor epoch was selected which conforms to an initial altitude of greater than a million feet. The Extended BET, ST19BET/UN=274885C, incorporated the previously mentioned inertial reconstructed state information and the Langley Atmospheric Information Retrieval System (LAIRS) atmosphere, ST19MET/UN=712662N, with some minor exceptions. Primary and back-up AEROBET reels are NK0165 and NK0201, respectively. This product was only developed over the lowermost 360 kft altitude range due to atmosphere problems but this relates to altitudes well above meaningful signal in the IMUs. Summary results generated from the AEROBET for this flight are presented with meaningful configuration and statistical comparisons from the previous thirteen flights. Modified maximum likelihood estimation (MMLE) files were generated based on IMU2 and the Rate Gyro Assembly/Accelerometer Assembly (RGA/AA), respectively. Appendices attached define spacecraft and physical constants utilized, show plots of the final tracking data residuals from the post-flight fit, list relevant parameters from the BET at a two second spacing, and retain for archival purpose all relevant input and output tapes and files generated.

JEOS. The JANUS earth observation satellite

NASA Astrophysics Data System (ADS)

Molette, P.; Jouan, J.

The JANUS multimission platform has been designed to minimize the cost of the satellite (by a maximum reuse of equipment from other proprogrammes) and of its associated launch by Aŕiane (by a piggy-back configuration optimized for Ariane 4). The paper describes the application of the JANUS platform to an Earth observation mission with the objective to provide a given country with a permanent monitoring of its earth resources by exploitation of spaceborne imagery. According to this objective, and to minimize the overall system and operational cost, the JANUS Earth Observation Satellite (JEOS) will provide a limited coverage with real time transmission of image data, thus avoiding need for on-board storage and simplifying operations. The JEOS operates on a low earth, near polar sun synchronous orbit. Launched in a piggy-back configuration on Ariane 4, with a SPOT or ERS spacecraft, it reaches its operational orbit after a drift orbit of a few weeks maximum. In its operational mode, the JEOS is 3-axis stabilised, earth pointed. After presentation of the platform, the paper describes the solid state push-broom camera which is composed of four optical lenses mounted on a highly stable optical bench. Each lens includes an optics system, reused from an on-going development, and two CCD linear arrays of detectors. The camera provides four registered channels in visible and near IR bands. The whole optical bench is supported by a rotating mechanism which allows rotation of the optical axis in the across-track direction. The JEOS typical performance for a 700 km altitude is then summarized: spatial resolution 30 m, swath width 120 km, off-track capability 325 km,… The payload data handling and transmission electronics, derived from the French SPOT satellite, realizes the processing, formatting, and transmission to the ground; this allows reuse of the standard SPOT receiving stations. The camera is only operated when the spacecraft is within the visibility of the ground station, and image data are directly transmitted to the ground station by the spacecraft X-band transmitter. Finally, the paper presents a set of typical Earth observation missions which can be realized with JEOS, for countries which wish to have their own observation system, possibly also as a complement to the SPOT and/or LANDSAT observation data.

1997 Spacecraft Contamination and Coatings Workshop

NASA Technical Reports Server (NTRS)

Chen, Philip T. (Compiler); Benner, Steve M. (Compiler)

1997-01-01

This volume contains the presentation charts of talks given at the "1997 Spacecraft Contamination and Coatings Workshop," held July 9-10, 1997, in Annapolis, Maryland. The workshop was attended by representatives from NASA, Jet Propulsion Laboratory, Department of Defense, industry, and universities concerned with the the spacecraft contamination engineering and thermal control coatings. The workshop provided a forum for exchanging new developments in spacecraft contamination and coatings.

Task four report: Telemetry, command, and data handling. [communication systems for ATS, SMS, OSO, and ERTS satellites

NASA Technical Reports Server (NTRS)

1973-01-01

An overview of the telemetry, command, and data handling features of four spacecraft developed under GSFC management is presented. Two of these spacecraft ATS and SMS, are designed for geostationary orbit; the other two OSO and ERTS, are designed for low earth orbits. The program time spans for these spacecraft are as shown. The programs are seen to be near contemporary, especially in the 1973, 1974 period. All of the spacecraft listed were developed under GSFC control and are thus subject to the standards set forth in the Aerospace Data System Standard developed by GSFC. These standards must be adhered to by all spacecraft programs under GSFC control or utilizing STDN unless waivers have been granted. The standards were developed to maximize the utilization of the large amount of standard equipment at each STDN ground facility. The standards impose bounds on both the command and telemetry formats to be compatible with the STDN ground station unless valid and acceptable reasons are raised to deviate from these restraints.

Survey of DMSP Charging Events During the Period Preceding Cycle 23 Solar Maximum

NASA Technical Reports Server (NTRS)

Parker, Linda Neergaard; Minow, Joseph I.

2013-01-01

It has been well established that POLAR orbiting satellites can see mild to severe charging levels during solar minimum conditions (Frooninckx and Sojka, 1992, Anderson and Koons, 1996, Anderson, 2012). However, spacecraft operations during solar maximum cannot be considered safe from auroral charging. Recently, we have seen examples of high level charging during the recent approach to solar maximum. We present here a survey of charging events seen by the Defense Meteorological Satellite Program (DMSP) satellites (F16, F17) during the solstices of 2011 and 2012. In this survey, we summarize the condition necessary for charging to occur in this environment, we describe how the lower than normal maximum conditions are conducive to the environment conditions necessary for charging in the POLAR orbit, and we show examples of the more extreme charging events, sometimes exceeding 1 kV, during this time period. We also show examples of other interesting phenomenological events seen in the DMSP data, but which are not considered surface charging events, and discuss the differences.

Survey of DMSP Charging During the Period Preceding Cycle 24 Solar Maximum

NASA Technical Reports Server (NTRS)

NeergaardParker, L.; Minow, Joseph I.

2013-01-01

It has been well established that polar orbiting satellites can see mild to severe charging levels during solar minimum conditions (Frooninckx and Sojka, 1992, Anderson and Koons, 1996, Anderson, 2012). However, spacecraft operations during solar maximum cannot be considered safe from auroral charging. Recently, we have seen examples of high level charging during the recent approach to solar maximum. We present here a survey of charging events seen by the Defense Meteorological Satellite Program (DMSP) satellites (F16, F17) during the solstices of 2011 and 2012. In this survey, we summarize the condition necessary for charging to occur in this environment, we describe how the lower than normal maximum conditions are conducive to the environment conditions necessary for charging in the polar orbit, and we show examples of the more extreme charging events, sometimes exceeding 1 kV, during this time period. We also show examples of other interesting phenomenological events seen in the DMSP data, but which are not considered surface charging events, and discuss the differences.

Solar maximum: Solar array degradation

NASA Technical Reports Server (NTRS)

Miller, T.

1985-01-01

The 5-year in-orbit power degradation of the silicon solar array aboard the Solar Maximum Satellite was evaluated. This was the first spacecraft to use Teflon R FEP as a coverglass adhesive, thus avoiding the necessity of an ultraviolet filter. The peak power tracking mode of the power regulator unit was employed to ensure consistent maximum power comparisons. Telemetry was normalized to account for the effects of illumination intensity, charged particle irradiation dosage, and solar array temperature. Reference conditions of 1.0 solar constant at air mass zero and 301 K (28 C) were used as a basis for normalization. Beginning-of-life array power was 2230 watts. Currently, the array output is 1830 watts. This corresponds to a 16 percent loss in array performance over 5 years. Comparison of Solar Maximum Telemetry and predicted power levels indicate that array output is 2 percent less than predictions based on an annual 1.0 MeV equivalent election fluence of 2.34 x ten to the 13th power square centimeters space environment.

Cost Optimization and Technology Enablement COTSAT-1

NASA Technical Reports Server (NTRS)

Spremo, Stevan; Lindsay, Michael C.; Klupar, Peter Damian; Swank, Aaron J.

2010-01-01

Cost Optimized Test of Spacecraft Avionics and Technologies (COTSAT-1) is an ongoing spacecraft research and development project at NASA Ames Research Center (ARC). The space industry was a hot bed of innovation and development at its birth. Many new technologies were developed for and first demonstrated in space. In the recent past this trend has reversed with most of the new technology funding and research being driven by the private industry. Most of the recent advances in spaceflight hardware have come from the cell phone industry with a lag of about 10 to 15 years from lab demonstration to in space usage. NASA has started a project designed to address this problem. The prototype spacecraft known as Cost Optimized Test of Spacecraft Avionics and Technologies (COTSAT-1) and CheapSat work to reduce these issues. This paper highlights the approach taken by NASA Ames Research center to achieve significant subsystem cost reductions. The COSTAT-1 research system design incorporates use of COTS (Commercial Off The Shelf), MOTS (Modified Off The Shelf), and GOTS (Government Off The Shelf) hardware for a remote sensing spacecraft. The COTSAT-1 team demonstrated building a fully functional spacecraft for $500K parts and $2.0M labor. The COTSAT-1 system, including a selected science payload, is described within this paper. Many of the advancements identified in the process of cost reduction can be attributed to the use of a one-atmosphere pressurized structure to house the spacecraft components. By using COTS hardware, the spacecraft program can utilize investments already made by commercial vendors. This ambitious project development philosophy/cycle has yielded the COTSAT-1 flight hardware. This paper highlights the advancements of the COTSAT-1 spacecraft leading to the delivery of the current flight hardware that is now located at NASA Ames Research Center. This paper also addresses the plans for COTSAT-2.

Spacecraft Formation Flying near Sun-Earth L2 Lagrange Point: Trajectory Generation and Adaptive Full-State Feedback Control

NASA Technical Reports Server (NTRS)

Wong, Hong; Kapila, Vikram

2004-01-01

In this paper, we present a method for trajectory generation and adaptive full-state feedback control to facilitate spacecraft formation flying near the Sun-Earth L2 Lagrange point. Specifically, the dynamics of a spacecraft in the neighborhood of a Halo orbit reveals that there exist quasi-periodic orbits surrounding the Halo orbit. Thus, a spacecraft formation is created by placing a leader spacecraft on a desired Halo orbit and placing follower spacecraft on desired quasi-periodic orbits. To produce a formation maintenance controller, we first develop the nonlinear dynamics of a follower spacecraft relative to the leader spacecraft. We assume that the leader spacecraft is on a desired Halo orbit trajectory and the follower spacecraft is to track a desired quasi-periodic orbit surrounding the Halo orbit. Then, we design an adaptive, full-state feedback position tracking controller for the follower spacecraft providing an adaptive compensation for the unknown mass of the follower spacecraft. The proposed control law is simulated for the case of the leader and follower spacecraft pair and is shown to yield global, asymptotic convergence of the relative position tracking errors.

DMSP Auroral Charging at Solar Cycle 24 Maximum

NASA Technical Reports Server (NTRS)

Chandler, M.; Parker, L. Neergaard; Minow, J. I.

2013-01-01

It has been well established that polar orbiting satellites can experience mild to severe auroral charging levels (on the order of a few hundred volts to few kilovolts negative frame potentials) during solar minimum conditions. These same studies have shown a strong reduction in charging during the rising and declining phases of the past few solar cycles with a nearly complete suppression of auroral charging at solar maximum. Recently, we have observed examples of high level charging during the recent approach to Solar Cycle 24 solar maximum conditions not unlike those reported by Frooninckx and Sojka. These observations demonstrate that spacecraft operations during solar maximum cannot be considered safe from auroral charging when solar activity is low. We present a survey of auroral charging events experienced by the Defense Meteorological Satellite Program (DMSP) F16 satellite during Solar Cycle 24 maximum conditions. We summarize the auroral energetic particle environment and the conditions necessary for charging to occur in this environment, we describe how the lower than normal solar activity levels for Solar Cycle 24 maximum conditions are conducive to charging in polar orbits, and we show examples of the more extreme charging events, sometimes exceeding 1 kV, during this time period.

DARPA Orbital Express program: effecting a revolution in space-based systems

NASA Astrophysics Data System (ADS)

Whelan, David A.; Adler, E. A.; Wilson, Samuel B., III; Roesler, Gordon M., Jr.

2000-11-01

A primary goal of the Defense Advanced Research Projects Agency is to develop innovative, high-risk technologies with the potential of a revolutionary impact on missions of the Department of Defense. DARPA is developing a space experiment to prove the feasibility of autonomous on- orbit servicing of spacecraft. The Orbital Express program will demonstrate autonomous on-orbit refueling, as well as autonomous delivery of a small payload representing an avionics upgrade package. The maneuverability provided to spacecraft from a ready refueling infrastructure will enable radical new capabilities for the military, civil and commercial spacecraft. Module replacement has the potential to extend bus lifetimes, and to upgrade the performance of key subsystems (e.g. processors) at the pace of technology development. The Orbital Express technology development effort will include the necessary autonomy for a viable servicing infrastructure; a universal interface for docking, refueling and module transfers; and a spacecraft bus design compatible with this servicing concept. The servicer spacecraft of the future may be able to act as a host platform for microsatellites, extending their capabilities while reducing risk. An infrastructure based on Orbital Express also benefits from, and stimulates the development of, lower-cost launch strategies.

Passive radiative cooling of a HTS coil for attitude orbit control in micro-spacecraft

NASA Astrophysics Data System (ADS)

Inamori, Takaya; Ozaki, Naoya; Saisutjarit, Phongsatorn; Ohsaki, Hiroyuki

2015-02-01

This paper proposes a novel radiative cooling system for a high temperature superconducting (HTS) coil for an attitude orbit control system in nano- and micro-spacecraft missions. These days, nano-spacecraft (1-10 kg) and micro-spacecraft (10-100 kg) provide space access to a broader range of spacecraft developers and attract interest as space development applications. In planetary and high earth orbits, most previous standard-size spacecraft used thrusters for their attitude and orbit control, which are not available for nano- and micro-spacecraft missions because of the strict power consumption, space, and weight constraints. This paper considers orbit and attitude control methods that use a superconducting coil, which interacts with on-orbit space plasmas and creates a propulsion force. Because these spacecraft cannot use an active cooling system for the superconducting coil because of their mass and power consumption constraints, this paper proposes the utilization of a passive radiative cooling system, in which the superconducting coil is thermally connected to the 3 K cosmic background radiation of deep space, insulated from the heat generation using magnetic holders, and shielded from the sun. With this proposed cooling system, the HTS coil is cooled to 60 K in interplanetary orbits. Because the system does not use refrigerators for its cooling system, the spacecraft can achieve an HTS coil with low power consumption, small mass, and low cost.

Control of nonlinear systems with applications to constrained robots and spacecraft attitude stabilization

NASA Technical Reports Server (NTRS)

Krishnan, Hariharan

1993-01-01

This thesis is organized in two parts. In Part 1, control systems described by a class of nonlinear differential and algebraic equations are introduced. A procedure for local stabilization based on a local state realization is developed. An alternative approach to local stabilization is developed based on a classical linearization of the nonlinear differential-algebraic equations. A theoretical framework is established for solving a tracking problem associated with the differential-algebraic system. First, a simple procedure is developed for the design of a feedback control law which ensures, at least locally, that the tracking error in the closed loop system lies within any given bound if the reference inputs are sufficiently slowly varying. Next, by imposing additional assumptions, a procedure is developed for the design of a feedback control law which ensures that the tracking error in the closed loop system approaches zero exponentially for reference inputs which are not necessarily slowly varying. The control design methodologies are used for simultaneous force and position control in constrained robot systems. The differential-algebraic equations are shown to characterize the slow dynamics of a certain nonlinear control system in nonstandard singularly perturbed form. In Part 2, the attitude stabilization (reorientation) of a rigid spacecraft using only two control torques is considered. First, the case of momentum wheel actuators is considered. The complete spacecraft dynamics are not controllable. However, the spacecraft dynamics are small time locally controllable in a reduced sense. The reduced spacecraft dynamics cannot be asymptotically stabilized using continuous feedback, but a discontinuous feedback control strategy is constructed. Next, the case of gas jet actuators is considered. If the uncontrolled principal axis is not an axis of symmetry, the complete spacecraft dynamics are small time locally controllable. However, the spacecraft attitude cannot be asymptotically stabilized using continuous feedback, but a discontinuous stabilizing feedback control strategy is constructed. If the uncontrolled principal axis is an axis of symmetry, the complete spacecraft dynamics cannot be stabilized. However, the spacecraft dynamics are small time locally controllable in a reduced sense. The reduced spacecraft dynamics cannot be asymptotically stabilized using continuous feedback, but again a discontinuous feedback control strategy is constructed.

Evaluation of a Liquid Amine System for Spacecraft Carbon Dioxide Control

NASA Technical Reports Server (NTRS)

Breaux, D. K.; Friedel, P.; Hwang, K. C.; Probert, G.; Ruder, J. M.; Sawamura, L.

1974-01-01

The analytical and experimental studies are described which were directed toward the acquisition of basic information on utilizing a liquid amine sorbent for in use in a CO2 removal system for manned spacecraft. Liquid amine systems are successfully used on submarines for control of CO2 generated by the crew, but liquid amines were not previously considered for spacecraft applications due to lack of development of satisfactory rotary phase separators. Developments in this area now make consideration of liquid amines practical for spacecraft system CO2 removal. The following major tasks were performed to evaluate liquid amine systems for spacecraft: (1) characterization, through testing, of the basic physical and thermodynamic properties of the amine solution; (2) determination of the dynamic characteristics of a cocurrent flow absorber; and (3) evaluation, synthesis, and selection of a liquid amine system concept oriented toward low power requirements. A low weight, low power system concept was developed. Numerical and graphical data are accompanied by pertinent observations.

NASA's Space Environments and Effects (SEE) Program: Contamination Engineering Technology Development

NASA Technical Reports Server (NTRS)

Pearson, Steven D.; Clifton, K. Stuart

1999-01-01

ABSTRACT The return of the Long Duration Exposure Facility (LDEF) in 1990 brought a wealth of space exposure data on materials, paints, solar cells, etc. and data on the many space environments. The effects of the harsh space environments can provide damaging or even disabling effects on spacecraft, its materials, and its instruments. In partnership with industry, academia, and other government agencies, National Aeronautics & Space Administration's (NASA's) Space Environments & Effects (SEE) Program defines the space environments and provides technology development to accommodate or mitigate these harmful environments on the spacecraft. This program provides a very comprehensive and focused approach to understanding the space environment, to define the best techniques for both flight and ground-based experimentation, to update the models which predict both the environments and the environmental effects on spacecraft, and finally to ensure that this information is properly maintained and inserted into spacecraft design programs. This paper will describe the current SEE Program and will present SEE contamination engineering technology development and risk mitigation for future spacecraft design.

NASA's Space Environments and Effects (SEE) program: contamination engineering technology development

NASA Astrophysics Data System (ADS)

Pearson, Steven D.; Clifton, K. Stuart

1999-10-01

The return of the Long Duration Exposure Facility (LDEF) in 1990 brought a wealth of space exposure data on materials, paints, solar cells, etc. and data on the many space environments. The effects of the harsh space environments can provide damaging or even disabling effects on spacecraft, its materials, and its instruments. In partnership with industry, academia, and other government agencies, National Aeronautics & Space Administration's (NASA's) Space Environments & Effects (SEE) Program defines the space environments and provides technology development to accommodate or mitigate these harmful environments on the spacecraft. This program provides a very comprehensive and focused approach to understanding the space environment, to define the best techniques for both flight and ground-based experimentation, to update the models which predict both the environments and the environmental effects on spacecraft, and finally to ensure that this information is properly maintained and inserted into spacecraft design programs. This paper will describe the current SEE Program and will present SEE contamination engineering technology development and risk mitigation for future spacecraft design.

Development of a space universal modular architecture (SUMO)

NASA Astrophysics Data System (ADS)

Collins, Bernie F.

This concept paper proposes that the space community should develop and implement a universal standard for spacecraft modularity - to improve interoperability of spacecraft components. Pursuing a global industry consensus standard for open and modular spacecraft architecture will encourage trade, remove standards-related market barriers, and in the long run increase both value provided to customers and profitability of the space industrial sector. This concept paper sets out: (1) the goals for a SUMO standard and how it will benefit the space community; (2) background on spacecraft modularity and existing related standards; (3) the proposed technical scope of the current standardization effort; and (4) an approach for creating a SUMO standard.

Ka-Band Autonomous Formation Flying Sensor

NASA Technical Reports Server (NTRS)

Tien, Jeffrey; Purcell, George, Jr.; Srinivasan, Jeffrey; Ciminera, Michael; Srinivasan, Meera; Meehan, Thomas; Young, Lawrence; Aung, MiMi; Amaro, Luis; Chong, Yong;

Dynamic Modeling of Ascent Abort Scenarios for Crewed Launches

NASA Technical Reports Server (NTRS)

Bigler, Mark; Boyer, Roger L.

2015-01-01

For the last 30 years, the United States's human space program has been focused on low Earth orbit exploration and operations with the Space Shuttle and International Space Station programs. After nearly 50 years, the U.S. is again working to return humans beyond Earth orbit. To do so, NASA is developing a new launch vehicle and spacecraft to provide this capability. The launch vehicle is referred to as the Space Launch System (SLS) and the spacecraft is called Orion. The new launch system is being developed with an abort system that will enable the crew to escape launch failures that would otherwise be catastrophic as well as probabilistic design requirements set for probability of loss of crew (LOC) and loss of mission (LOM). In order to optimize the risk associated with designing this new launch system, as well as verifying the associated requirements, NASA has developed a comprehensive Probabilistic Risk Assessment (PRA) of the integrated ascent phase of the mission that includes the launch vehicle, spacecraft and ground launch facilities. Given the dynamic nature of rocket launches and the potential for things to go wrong, developing a PRA to assess the risk can be a very challenging effort. Prior to launch and after the crew has boarded the spacecraft, the risk exposure time can be on the order of three hours. During this time, events may initiate from either of the spacecraft, the launch vehicle, or the ground systems, thus requiring an emergency egress from the spacecraft to a safe ground location or a pad abort via the spacecraft's launch abort system. Following launch, again either the spacecraft or the launch vehicle can initiate the need for the crew to abort the mission and return to the home. Obviously, there are thousands of scenarios whose outcome depends on when the abort is initiated during ascent as to how the abort is performed. This includes modeling the risk associated with explosions and benign system failures that require aborting a spacecraft under very dynamic conditions, particularly in the lower atmosphere, and returning the crew home safely. This paper will provide an overview of the PRA model that has been developed of this new launch system, including some of the challenges that are associated with this effort. Key Words: PRA, space launches, human space program, ascent abort, spacecraft, launch vehicles

How Emerging Technologies are Changing the Rules of Spacecraft Ground Support

NASA Technical Reports Server (NTRS)

Boland, Dillard; Steger, Warren; Weidow, David; Yakstis, Lou

1996-01-01

As part of its effort to develop the flight dynamics distributed system (FDDS), NASA established a program for the continual monitoring of the developments in computer and software technologies, and for assessing the significance of constructing and operating spacecraft ground data systems. In relation to this, technology trends in the computing industry are reviewed, exploring their significance for the spacecraft ground support industry. The technologies considered are: hardware; object computing; Internet; automation, and software development. The ways in which these technologies have affected the industry are considered.

Characteristics of trapped proton anisotropy at Space Station Freedom altitudes

NASA Technical Reports Server (NTRS)

Armstrong, T. W.; Colborn, B. L.; Watts, J. W.

1990-01-01

The ionizing radiation dose for spacecraft in low-Earth orbit (LEO) is produced mainly by protons trapped in the Earth's magnetic field. Current data bases describing this trapped radiation environment assume the protons to have an isotropic angular distribution, although the fluxes are actually highly anisotropic in LEO. The general nature of this directionality is understood theoretically and has been observed by several satellites. The anisotropy of the trapped proton exposure has not been an important practical consideration for most previous LEO missions because the random spacecraft orientation during passage through the radiation belt 'averages out' the anisotropy. Thus, in spite of the actual exposure anisotropy, cumulative radiation effects over many orbits can be predicted as if the environment were isotropic when the spacecraft orientation is variable during exposure. However, Space Station Freedom will be gravity gradient stabilized to reduce drag, and, due to this fixed orientation, the cumulative incident proton flux will remain anisotropic. The anisotropy could potentially influence several aspects of Space Station design and operation, such as the appropriate location for radiation sensitive components and experiments, location of workstations and sleeping quarters, and the design and placement of radiation monitors. Also, on-board mass could possible be utilized to counteract the anisotropy effects and reduce the dose exposure. Until recently only omnidirectional data bases for the trapped proton environment were available. However, a method to predict orbit-average, angular dependent ('vector') trapped proton flux spectra has been developed from the standard omnidirectional trapped proton data bases. This method was used to characterize the trapped proton anisotropy for the Space Station orbit (28.5 degree inclination, circular) in terms of its dependence on altitude, solar cycle modulation (solar minimum vs. solar maximum), shielding thickness, and radiation effect (silicon rad and rem dose).

SmallSat Innovations for Planetary Science

NASA Astrophysics Data System (ADS)

Weinberg, Jonathan; Petroy, Shelley; Roark, Shane; Schindhelm, Eric

2017-10-01

As NASA continues to look for ways to fly smaller planetary missions such as SIMPLEX, MoO, and Venus Bridge, it is important that spacecraft and instrument capabilities keep pace to allow these missions to move forward. As spacecraft become smaller, it is necessary to balance size with capability, reliability and payload capacity. Ball Aerospace offers extensive SmallSat capabilities matured over the past decade, utilizing our broad experience developing mission architecture, assembling spacecraft and instruments, and testing advanced enabling technologies. Ball SmallSats inherit their software capabilities from the flight proven Ball Configurable Platform (BCP) line of spacecraft, and may be tailored to meet the unique requirements of Planetary Science missions. We present here recent efforts in pioneering both instrument miniaturization and SmallSat/sensorcraft development through mission design and implementation. Ball has flown several missions with small, but capable spacecraft. We also have demonstrated a variety of enhanced spacecraft/instrument capabilities in the laboratory and in flight to advance autonomy in spaceflight hardware that can enable some small planetary missions.

Nonlinear Dynamic Behavior in the Cassini Spacecraft Modal Survey

NASA Technical Reports Server (NTRS)

Carney, Kelly S.

1997-01-01

In October 1997, the 6-ton robotic spacecraft, Cassini, will lift off from Cape Canaveral atop a Titan IV B rocket, beginning a 7-year journey to Saturn. Upon completion of that voyage, Cassini will send the Huygens probe into the atmosphere of Saturn's largest moon, Titan. Cassini will then spend years studying Saturn's vast realm of rings, icy moons, and magnetic fields. The size and complexity of this endeavor mandates the involvement of many organizations. The Jet Propulsion Laboratory (JPL) manages the project for NASA and is responsible for the spacecraft design, development, and assembly. The NASA Lewis Research Center is the launch system integrator. As is typical for such a spacecraft, a test-verified finite element model is required for loads analysis. JPL had responsibility for the Cassini modal survey and the development of the spacecraft test-verified finite element model. Test verification is a complex and sometimes subjective process. Because of this, NASA Lewis independently verified and validated the Cassini spacecraft modal survey.

Full-Scale Spacecraft Simulator Design for a 2D Zero Gravity Small Body Surface Sampling Validation

NASA Astrophysics Data System (ADS)

Mongelli, Marco

NASA is developing several Touch-And-Go (TAG) classes of missions. These types of missions like the OSIRIS-REx asteroid sample return [1] or a comet sample return mission (CSSR)[2], consist usually in three phases: propulsive approach to the target, sampling and propulsion to move the spacecraft away from the target. The development of TAG mission, from concept to realization, is usually divided in two phases: Phase I discusses the major trades that could affect the mission architecture; Phase II focuses in detail on the design. This project of a spacecraft emulator fits into phase II and specifically on the way the spacecraft could react in absence of gravity while the Sample Acquisition System (SAS) is collecting the sample. A full-scale spacecraft on a 2D zero-friction environment has been designed. Also a propulsion system has been implemented to re-create the full dynamics of a spacecraft in space.

The Spacecraft Fire Experiment (Saffire) - Objectives, Development and Status

NASA Technical Reports Server (NTRS)

Schoren, William; Ruff, Gary A.; Urban, David L.

2016-01-01

Since 2012, the Spacecraft Fire Experiment (Saffire) has been under development by the Spacecraft Fire Safety Demonstration (SFS Demo) project that is funded by NASA's Advanced Exploration Systems Division in the Human Exploration and Operations Mission Directorate. The overall objective of this project is to reduce the uncertainty and risk associated with the design of spacecraft fire safety systems for NASA's exploration missions. This is accomplished by defining, developing, and conducting experiments that address gaps in spacecraft fire safety knowledge and capabilities identified by NASA's Fire Safety System Maturation Team. This paper describes the three Spacecraft Fire Experiments (Saffire-I, -II, and -III) that were developed at NASA-GRC and that will conduct a series of material flammability tests in low-gravity and at length scales that are realistic for a spacecraft fire. The experiments will be conducted in Orbital ATK's Cygnus vehicle after it has unberthed from the International Space Station. The tests will be fully automated with the data downlinked at the conclusion of the test and before the Cygnus vehicle reenters the atmosphere. The objectives of these experiments are to (1) determine how rapidly a large scale fire grows in low-gravity and (2) investigate the low-g flammability limits compared to those obtained in NASA's normal gravity material flammability screening test. The hardware for these experiments has been completed and is awaiting their respective launches, all planned for 2016. This paper will review the objectives of these experiments and how they address several of the knowledge gaps for NASA's exploration missions. The hardware development will be discussed including several novel approaches that were taken for testing and evaluation of these series payloads. The status of the missions and operational status will also be presented.

Prediction of three sigma maximum dispersed density for aerospace applications

NASA Technical Reports Server (NTRS)

Charles, Terri L.; Nitschke, Michael D.

1993-01-01

Free molecular heating (FMH) is caused by the transfer of energy during collisions between the upper atmosphere molecules and a space vehicle. The dispersed free molecular heating on a surface is an important constraint for space vehicle thermal analyses since it can be a significant source of heating. To reduce FMH to a spacecraft, the parking orbit is often designed to a higher altitude at the expense of payload capability. Dispersed FMH is a function of both space vehicle velocity and atmospheric density, however, the space vehicle velocity variations are insignificant when compared to the atmospheric density variations. The density of the upper atmosphere molecules is a function of altitude, but also varies with other environmental factors, such as solar activity, geomagnetic activity, location, and time. A method has been developed to predict three sigma maximum dispersed density for up to 15 years into the future. This method uses a state-of-the-art atmospheric density code, MSIS 86, along with 50 years of solar data, NASA and NOAA solar activity predictions for the next 15 years, and an Aerospace Corporation correlation to account for density code inaccuracies to generate dispersed maximum density ratios denoted as 'K-factors'. The calculated K-factors can be used on a mission unique basis to calculate dispersed density, and hence dispersed free molecular heating rates. These more accurate K-factors can allow lower parking orbit altitudes, resulting in increased payload capability.

Proceedings of the Spacecraft Charging Technology Conference: Executive Summary

NASA Technical Reports Server (NTRS)

Pike, C. P.; Whipple, E. C., Jr.; Stevens, N. J.; Minges, M. L.; Lehn, W. L.; Bunn, M. H.

1977-01-01

Aerospace environments are reviewed in reference to spacecraft charging. Modelling, a theoretical scheme which can be used to describe the structure of the sheath around the spacecraft and to calculate the charging currents within, is discussed. Materials characterization is considered for experimental determination of the behavior of typical spacecraft materials when exposed to simulated geomagnetic substorm conditions. Materials development is also examined for controlling and minimizing spacecraft charging or at least for distributing the charge in an equipotential manner, using electrical conductive surfaces for materials exposed to space environment.

Power, Propulsion, and Communications for Microspacecraft Missions

NASA Technical Reports Server (NTRS)

deGroot, W. A.; Maloney, T. M.; Vanderaar, M. J.

1998-01-01

The development of small sized, low weight spacecraft should lead to reduced scientific mission costs by lowering fabrication and launch costs. An order of magnitude reduction in spacecraft size can be obtained by miniaturizing components. Additional reductions in spacecraft weight, size, and cost can be obtained by utilizing the synergy that exists between different spacecraft systems. The state-of-the-art of three major systems, spacecraft power, propulsion, and communications is discussed. Potential strategies to exploit the synergy between these systems and/or the payload are identified. Benefits of several of these synergies are discussed.

Concurrent engineering: Spacecraft and mission operations system design

NASA Technical Reports Server (NTRS)

Landshof, J. A.; Harvey, R. J.; Marshall, M. H.

1994-01-01

Despite our awareness of the mission design process, spacecraft historically have been designed and developed by one team and then turned over as a system to the Mission Operations organization to operate on-orbit. By applying concurrent engineering techniques and envisioning operability as an essential characteristic of spacecraft design, tradeoffs can be made in the overall mission design to minimize mission lifetime cost. Lessons learned from previous spacecraft missions will be described, as well as the implementation of concurrent mission operations and spacecraft engineering for the Near Earth Asteroid Rendezvous (NEAR) program.

General Methodology for Designing Spacecraft Trajectories

NASA Technical Reports Server (NTRS)

Condon, Gerald; Ocampo, Cesar; Mathur, Ravishankar; Morcos, Fady; Senent, Juan; Williams, Jacob; Davis, Elizabeth C.

2012-01-01

A methodology for designing spacecraft trajectories in any gravitational environment within the solar system has been developed. The methodology facilitates modeling and optimization for problems ranging from that of a single spacecraft orbiting a single celestial body to that of a mission involving multiple spacecraft and multiple propulsion systems operating in gravitational fields of multiple celestial bodies. The methodology consolidates almost all spacecraft trajectory design and optimization problems into a single conceptual framework requiring solution of either a system of nonlinear equations or a parameter-optimization problem with equality and/or inequality constraints.

Saturn Apollo Program

NASA Image and Video Library

1973-01-01

This illustration depicts a configuration of the Soyuz spacecraft for the Apollo-Soyuz Test Project (ASTP). The ASTP was the first international docking of the U.S.'s Apollo spacecraft and the U.S.S.R.'s Soyuz spacecraft in space. For this project, the Soviets built another in their continuing series of Soyuz space capsules. The U.S. used the Saturn IB Apollo capsule. A joint engineering team from the two countries met to develop a docking system that permitted the two spacecraft to link in space and allowed the crews to travel from one spacecraft to the other.

Searches for Plumes and Ongoing Geologic Activity on Europa from Galileo and Other Spacecraft

NASA Astrophysics Data System (ADS)

Phillips, C. B.

2014-12-01

The recent discovery of an apparent plume erupting from Europa's surface using data from the Hubble Space Telescope (Roth et al. 2014) has prompted renewed interest in the possibility of recent or ongoing geologic activity on Europa. Here we summarize previous searches for plumes and changes on Europa's surface, and make recommendations for future efforts. During the period of time in which the Galileo spacecraft was in orbit in the Jupiter system, we made a number of comparisons with observations taken 20 years earlier by the Voyager spacecraft to look for surface changes (Phillips et al. 2000). We found no changes which were visible on Europa's surface. These comparisons, however, were necessarily limited by the low resolution of the Voyager images, which had a maximum resolution of about 2 km/pixel. We also used Galileo spacecraft data to search for plumes of material being ejected from Europa's surface. A 30-image observation was taken in 1999 to observe the limb and the dark sky just off the limb in a search for active plumes, but no plumes were observed (Phillips et al. 2000). However, Hoppa et al (1999) suggested that this image sequence occurred under unfavorable tidal stress conditions. Plume searches were also performed in eclipse images, but again no plumes were detected. More recently, we compared global-scale images of Europa taken in 2007 by the New Horizons spacecraft during its Jupiter flyby en route to Pluto (Bramson et al. 2011). After a careful search that included the iterative coregistration and ratioing techniques developed by Phillips et al. (2000), again, no changes were found on Europa's surface. If the recent Roth et al. (2014) suggestions of an active plume on Europa prove to be correct, we infer that one of two possibilities must be the case. Either 1) the plume is a recent event and was not active before the 2007 New Horizons flyby; or 2) the plume is intermittent and low-density, consisting primarily of gas and not dust, and therefore results in little to no detectable surface deposits on Europa. References: Roth, L. et al. Science 343, 171-174, 2014. Phillips, C. B., et al. JGR 105, 22579-22598, 2000. Hoppa, G., et al. Icarus 141, 287-298, 1999. Bramson, A. M., et al. LPSC 42, Abs. 1606, 2011.

Space station integrated wall design and penetration damage control

NASA Technical Reports Server (NTRS)

Coronado, A. R.; Gibbins, M. N.; Wright, M. A.; Stern, P. H.

1987-01-01

A methodology was developed to allow a designer to optimize the pressure wall, insulation, and meteoroid/debris shield system of a manned spacecraft for a given spacecraft configuration and threat environment. The threat environment consists of meteoroids and orbital debris, as specified for an arbitrary orbit and expected lifetime. An overall probability of no penetration is calculated, as well as contours of equal threat that take into account spacecraft geometry and orientation. Techniques, tools, and procedures for repairing an impacted and penetrated pressure wall were developed and tested. These techniques are applied from the spacecraft interior and account for the possibility of performing the repair in a vacuum. Hypervelocity impact testing was conducted to: (1) develop and refine appropriate penetration functions, and (2) determine the internal effects of a penetration on personnel and equipment.

Optimization of Closed Loop Eigenvalues: Maneuvering, Vibration Control, and Structure/Control Design Iteration for Flexible Spacecraft.

DTIC Science & Technology

1986-05-31

Nonlinear Feedback Control 8-16 for Spacecraft Attitude Maneuvers" 2. " Spacecraft Attitude Control Using 17-35... nonlinear state feedback control laws are developed for space- craft attitude control using the Euler parameters and conjugate angular momenta. Time... Nonlinear Feedback Control for Spacecraft Attitude Maneuvers," to appear in AIAA J. of Guidance, Control, and Dynamics, (AIAA Paper No. 83-2230-CP,

Use of Model Payload for Europa Mission Development

NASA Technical Reports Server (NTRS)

Lewis, Kari; Klaasan, Ken; Susca, Sara; Oaida, Bogdan; Larson, Melora; Vanelli, Tony; Murray, Alex; Jones, Laura; Thomas, Valerie; Frank, Larry

2016-01-01

This paper discusses the basis for the Model Payload and how it was used to develop the mission design, observation and data acquisition strategy, needed spacecraft capabilities, spacecraft-payload interface needs, mission system requirements and operational scenarios.

Maximum Torque and Momentum Envelopes for Reaction Wheel Arrays

NASA Technical Reports Server (NTRS)

Markley, F. Landis; Reynolds, Reid G.; Liu, Frank X.; Lebsock, Kenneth L.

2009-01-01

Spacecraft reaction wheel maneuvers are limited by the maximum torque and/or angular momentum that the wheels can provide. For an n-wheel configuration, the torque or momentum envelope can be obtained by projecting the n-dimensional hypercube, representing the domain boundary of individual wheel torques or momenta, into three dimensional space via the 3xn matrix of wheel axes. In this paper, the properties of the projected hypercube are discussed, and algorithms are proposed for determining this maximal torque or momentum envelope for general wheel configurations. Practical strategies for distributing a prescribed torque or momentum among the n wheels are presented, with special emphasis on configurations of four, five, and six wheels.

The 1989 Solar Maximum Mission event list

NASA Technical Reports Server (NTRS)

Dennis, B. R.; Licata, J. P.; Tolbert, A. K.

1992-01-01

This document contains information on solar burst and transient activity observed by the Solar Maximum Mission (SMM) during 1989 pointed observations. Data from the following SMM experiments are included: (1) Gamma Ray Spectrometer, (2) Hard X-Ray Burst Spectrometer, (3) Flat Crystal Spectrometer, (4) Bent Crystal Spectrometer, (5) Ultraviolet Spectrometer Polarimeter, and (6) Coronagraph/Polarimeter. Correlative optical, radio, and Geostationary Operational Satellite (GOES) X-ray data are also presented. Where possible, bursts or transients observed in the various wavelengths were grouped into discrete flare events identified by unique event numbers. Each event carries a qualifier denoting the quality or completeness of the observations. Spacecraft pointing coordinates and flare site angular displacement values from sun center are also included.

The 1988 Solar Maximum Mission event list

NASA Technical Reports Server (NTRS)

Dennis, B. R.; Licata, J. P.; Tolbert, A. K.

1992-01-01

Information on solar burst and transient activity observed by the Solar Maximum Mission (SMM) during 1988 pointed observations is presented. Data from the following SMM experiments are included: (1) gamma ray spectrometer; (2) hard x ray burst spectrometer; (3) flat crystal spectrometers; (4) bent crystal spectrometer; (5) ultraviolet spectrometer polarimeter; and (6) coronagraph/polarimeter. Correlative optical, radio, and Geostationary Operational Environmental Satellite (GOES) x ray data are also presented. Where possible, bursts, or transients observed in the various wavelengths were grouped into discrete flare events identified by unique event numbers. Each event carries a qualifier denoting the quality or completeness of the observation. Spacecraft pointing coordinates and flare site angular displacement values from sun center are also included.

The 1984 - 1987 Solar Maximum Mission event list

NASA Technical Reports Server (NTRS)

Dennis, B. R.; Licata, J. P.; Nelson, J. J.; Tolbert, A. K.

1992-01-01

Information on solar burst and transient activity observed by the Solar Maximum Mission (SMM) during 1984-1987 pointed observations is presented. Data from the following SMM experiments are included: (1) gamma ray spectrometer; (2) hard x-ray burst spectrometer; (3) flat crystal spectrometer; (4) bent crystal spectrometer; (5) ultraviolet spectrometer polarimeter; and (6) coronograph/polarimeter. Correlative optical, radio, and Geostationary Operational Environmental Satellite (GOES) x ray data are also presented. Where possible, bursts or transients observed in the various wavelengths were grouped into discrete flare events identified by unique event numbers. Each event carries a qualifier denoting the quality or completeness of the observations. Spacecraft pointing coordinates and flare site angular displacement values from sun center are also included.

Development of a rotary power transformer and inverter drive for spacecraft

NASA Technical Reports Server (NTRS)

Mclyman, W. T.; Bridgeforth, A. O.

1983-01-01

Many future satellites and spacecraft with spun and despun configurations will require the transfer of power across rotating interfaces in lieu of slip-rings and/or flexures. This is particularly true of spacecraft that have to demonstrate a long life expectancy. The rotary transformer has the desirable characteristics of high reliability and low noise, which qualify it as a potential replacement for slip rings. Development of a rotary power transformer follows the successful completion of a task to develop rotary signal-level transformers for the Galileo Spacecraft Project. The physical configuration of a rotary power transformer has a significant effect on its magnetic and electrical characteristics and therefore impacts the design of the dc/ac inverter driver. Important characteristics addressed during this development effort include: operating frequency, efficiency, transformer gap size, leakage inductance, and leakage flux. A breadboard inverter and rotary transformer were designed, fabricated and tested.

Long-Term Solar and Cosmic Radiation Data Bases

DTIC Science & Technology

1991-01-01

determine the magnitude of the variations in the cosmic ray intensity caused by solar activity. Neutron monitors, with their much lower energy threshold...expression that neutron monitors are sensors on spacecraft EARTH. Here we will consider cosmic ray detectors to measure two components of cosmic ...A comparison with the solar cycle as illustrated by the sunspot number in Fig. 1. shows that the maximum cosmic ray intensity occurs near sunspot

Failure of Harmonic Gears During Verification of a Two-Axis Gimbal for the Mars Reconnaissance Orbiter Spacecraft

NASA Technical Reports Server (NTRS)

Johnson, Michael R.; Gehling, Russ; Head, Ray

2006-01-01

The Mars Reconnaissance Orbiter (MRO) spacecraft has three two-axis gimbal assemblies that support and move the High Gain Antenna and two solar array wings. The gimbal assemblies are required to move almost continuously throughout the mission's seven-year lifetime, requiring a large number of output revolutions for each actuator in the gimbal assemblies. The actuator for each of the six axes consists of a two-phase brushless dc motor with a direct drive to the wave generator of a size-32 cup-type harmonic gear. During life testing of an actuator assembly, the harmonic gear teeth failed completely, leaving the size-32 harmonic gear with a maximum output torque capability less than 10% of its design capability. The investigation that followed the failure revealed limitations of the heritage material choices that were made for the harmonic gear components that had passed similar life requirements on several previous programs. Additionally, the methods used to increase the stiffness of a standard harmonic gear component set, while accepted practice for harmonic gears, is limited in its range. The stiffness of harmonic gear assemblies can be increased up to a maximum stiffness point that, if exceeded, compromises the reliability of the gear components for long life applications.

Improved Acoustic Blanket Developed and Tested

NASA Technical Reports Server (NTRS)

1996-01-01

Acoustic blankets are used in the payload fairing of expendable launch vehicles to reduce the fairing's interior acoustics and the subsequent vibration response of the spacecraft. The Cassini spacecraft, to be launched on a Titan IV in October 1997, requires acoustic levels lower than those provided by the standard Titan IV blankets. Therefore, new acoustic blankets were recently developed and tested to reach NASA's goal of reducing the Titan IV acoustic environment to the allowable levels for the Cassini spacecraft.

The Generic Spacecraft Analyst Assistant (GenSAA): A tool for automating spacecraft monitoring with expert systems

NASA Technical Reports Server (NTRS)

Hughes, Peter M.; Luczak, Edward C.

1991-01-01

Flight Operations Analysts (FOAs) in the Payload Operations Control Center (POCC) are responsible for monitoring a satellite's health and safety. As satellites become more complex and data rates increase, FOAs are quickly approaching a level of information saturation. The FOAs in the spacecraft control center for the COBE (Cosmic Background Explorer) satellite are currently using a fault isolation expert system named the Communications Link Expert Assistance Resource (CLEAR), to assist in isolating and correcting communications link faults. Due to the success of CLEAR and several other systems in the control center domain, many other monitoring and fault isolation expert systems will likely be developed to support control center operations during the early 1990s. To facilitate the development of these systems, a project was initiated to develop a domain specific tool, named the Generic Spacecraft Analyst Assistant (GenSAA). GenSAA will enable spacecraft analysts to easily build simple real-time expert systems that perform spacecraft monitoring and fault isolation functions. Lessons learned during the development of several expert systems at Goddard, thereby establishing the foundation of GenSAA's objectives and offering insights in how problems may be avoided in future project, are described. This is followed by a description of the capabilities, architecture, and usage of GenSAA along with a discussion of its application to future NASA missions.

Developing Software for NASA Missions in the New Millennia

NASA Technical Reports Server (NTRS)

Truszkowski, Walt; Rash, James; Rouff, Christopher; Hinchey, Mike

2004-01-01

NASA is working on new mission concepts for exploration of the solar system. The concepts for these missions include swarms of hundreds of cooperating intelligent spacecraft which will be able to work in teams and gather more data than current single spacecraft missions. These spacecraft will not only have to operate independently for long periods of time on their own and in teams, but will also need to have autonomic properties of self healing, self configuring, self optimizing and self protecting for them to survive in the harsh space environment. Software for these types of missions has never been developed before and represents some of the challenges of software development in the new millennia. The Autonomous Nano Technology Swarm (ANTS) mission is an example of one of the swarm missions NASA is considering. The ANTS mission will use a swarm of one thousand pico-spacecraft that weigh less than five pounds. Using an insect colony analog, ANTS will explore the asteroid belt and catalog the mass, density, morphology, and chemical composition of the asteroids. Due to the size of the spacecraft, each will only carry a single miniaturized science instrument which will require them to cooperate in searching for asteroids that are of scientific interest. This article also discusses the ANTS mission, the properties the spacecraft will need and how that will effect future software development.

An Assessment of Educational Benefits from the OpenOrbiter Space Program

ERIC Educational Resources Information Center

Straub, Jeremy; Whalen, David

2013-01-01

This paper analyzes the educational impact of the OpenOrbiter Small Spacecraft Development Initiative, a CubeSat development program underway at the University of North Dakota. OpenOrbiter includes traditional STEM activities (e.g., spacecraft engineering, software development); it also incorporates students from non-STEM disciplines not generally…

COTSAT Small Spacecraft Cost Optimization for Government and Commercial Use

NASA Technical Reports Server (NTRS)

Swank, Aaron J.; Bui, David; Dallara, Christopher; Ghassemieh, Shakib; Hanratty, James; Jackson, Evan; Klupar, Pete; Lindsay, Michael; Ling, Kuok; Mattei, Nicholas;

Solar Sailing is not Science Fiction Anymore

NASA Technical Reports Server (NTRS)

Alhorn, Dean C.

2010-01-01

Over 400 years ago Johannes Kepler envisioned the use of sunlight to propel a spacecraft. Just this year, a solar sail was deployed in orbit for the first time and proved that a spacecraft could effectively use a solar sail for propulsion. NASA's first nano-class solar sail satellite, NanoSail-D was designed and developed in only four months. Although the first unit was lost during the Falcon 1 rocket failure in 2008, the second flight unit has been refurbished and is waiting to be launched later this year. NanoSail-D will further the research into solar sail enabled spacecraft. It will be the first of several more sail enabled spacecraft to be launch in the next few years. FeatherSail is the next generation nano-class sail spacecraft being designed with the goal to prove low earth orbit operational capabilities. Future solar sail spacecraft will require novel ideas and innovative research for the continued development of space systems. One such pioneering idea is the Small Multipurpose Advanced Reconfigurable Technology (SMART) project. The SMART technology has the potential to revolutionize spacecraft avionics. Even though solar sailing is currently in its infancy, the next decade will provide great opportunities for research into sailing in outer space.

Cassini Attitude Control Operations Flight Rules and How They are Enforced

NASA Technical Reports Server (NTRS)

Burk, Thomas; Bates, David

2008-01-01

The Cassini spacecraft was launched on October 15, 1997 and arrived at Saturn on June 30, 2004. It has performed detailed observations and remote sensing of Saturn, its rings, and its satellites since that time. Cassini deployed the European-built Huygens probe which descended through the Titan atmosphere and landed on its surface on January 14, 2005. Operating the Cassini spacecraft is a complex scientific, engineering, and management job. In order to safely operate the spacecraft, a large number of flight rules were developed. These flight rules must be enforced throughout the lifetime of the Cassini spacecraft. Flight rules are defined as any operational limitation imposed by the spacecraft system design, hardware, and software, violation of which would result in spacecraft damage, loss of consumables, loss of mission objectives, loss and/or degradation of science, and less than optimal performance. Flight rules require clear description and rationale. Detailed automated methods have been developed to insure the spacecraft is continuously operated within these flight rules. An overview of all the flight rules allocated to the Cassini Attitude Control and Articulation Subsystem and how they are enforced is presented in this paper.

The Generic Spacecraft Analyst Assistant (gensaa): a Tool for Developing Graphical Expert Systems

NASA Technical Reports Server (NTRS)

Hughes, Peter M.

1993-01-01

During numerous contacts with a satellite each day, spacecraft analysts must closely monitor real-time data. The analysts must watch for combinations of telemetry parameter values, trends, and other indications that may signify a problem or failure. As the satellites become more complex and the number of data items increases, this task is becoming increasingly difficult for humans to perform at acceptable performance levels. At NASA GSFC, fault-isolation expert systems are in operation supporting this data monitoring task. Based on the lessons learned during these initial efforts in expert system automation, a new domain-specific expert system development tool named the Generic Spacecraft Analyst Assistant (GenSAA) is being developed to facilitate the rapid development and reuse of real-time expert systems to serve as fault-isolation assistants for spacecraft analysts. Although initially domain-specific in nature, this powerful tool will readily support the development of highly graphical expert systems for data monitoring purposes throughout the space and commercial industry.

A Hyperbolic Ontology Visualization Tool for Model Application Programming Interface Documentation

NASA Technical Reports Server (NTRS)

Hyman, Cody

2011-01-01

Spacecraft modeling, a critically important portion in validating planned spacecraft activities, is currently carried out using a time consuming method of mission to mission model implementations and integration. A current project in early development, Integrated Spacecraft Analysis (ISCA), aims to remedy this hindrance by providing reusable architectures and reducing time spent integrating models with planning and sequencing tools. The principle objective of this internship was to develop a user interface for an experimental ontology-based structure visualization of navigation and attitude control system modeling software. To satisfy this, a number of tree and graph visualization tools were researched and a Java based hyperbolic graph viewer was selected for experimental adaptation. Early results show promise in the ability to organize and display large amounts of spacecraft model documentation efficiently and effectively through a web browser. This viewer serves as a conceptual implementation for future development but trials with both ISCA developers and end users should be performed to truly evaluate the effectiveness of continued development of such visualizations.

[Development of fixed-base full task space flight training simulator].

PubMed

Xue, Liang; Chen, Shan-quang; Chang, Tian-chun; Yang, Hong; Chao, Jian-gang; Li, Zhi-peng

2003-01-01

Fixed-base full task flight training simulator is a very critical and important integrated training facility. It is mostly used in training of integrated skills and tasks, such as running the flight program of manned space flight, dealing with faults, operating and controlling spacecraft flight, communicating information between spacecraft and ground. This simulator was made up of several subentries including spacecraft simulation, simulating cabin, sight image, acoustics, main controlling computer, instructor and assistant support. It has implemented many simulation functions, such as spacecraft environment, spacecraft movement, communicating information between spacecraft and ground, typical faults, manual control and operating training, training control, training monitor, training database management, training data recording, system detecting and so on.

Microbiological sampling of spacecraft cabling, antennas, solar panels and thermal blankets

NASA Technical Reports Server (NTRS)

Koukol, R. C.

1973-01-01

Sampling procedures and techniques described resulted from various flight project microbiological monitoring programs of unmanned planetary spacecraft. Concurrent with development of these procedures, compatibility evaluations were effected with the cognizant spacecraft subsystem engineers to assure that degradation factors would not be induced during the monitoring program. Of significance were those areas of the spacecraft configuration for which special handling precautions and/or nonstandard sample gathering techniques were evolved. These spacecraft component areas were: cabling, high gain antenna, solar panels, and thermal blankets. The compilation of these techniques provides a historical reference for both the qualification and quantification of sampling parameters as applied to the Mariner Spacecraft of the late 1960's and early 1970's.

Spacecraft Charge Monitor

NASA Astrophysics Data System (ADS)

Goembel, L.

2003-12-01

We are currently developing a flight prototype Spacecraft Charge Monitor (SCM) with support from NASA's Small Business Innovation Research (SBIR) program. The device will use a recently proposed high energy-resolution electron spectroscopic technique to determine spacecraft floating potential. The inspiration for the technique came from data collected by the Atmosphere Explorer (AE) satellites in the 1970s. The data available from the AE satellites indicate that the SCM may be able to determine spacecraft floating potential to within 0.1 V under certain conditions. Such accurate measurement of spacecraft charge could be used to correct biases in space plasma measurements. The device may also be able to measure spacecraft floating potential in the solar wind and in orbit around other planets.

Rockets and spacecraft: Sine qua non of space science

NASA Technical Reports Server (NTRS)

1980-01-01

The evolution of the national launch vehicle stable is presented along with lists of launch vehicles used in NASA programs. A partial list of spacecraft used throughout the world is also given. Scientific spacecraft costs are presented along with an historial overview of project development and funding in NASA.

Model Spacecraft Construction, Units for Secondary School Industrial Arts.

ERIC Educational Resources Information Center

Dean, C. Thomas; And Others

This publication provides twelve model spacecraft construction plans for use by secondary school teachers in industrial arts classes. These models were adopted and developed from plans supplied by the National Aeronautics and Space Administration and are representative selections from the many spacecraft used in space exploration programs. Some…

Method for CFD Simulation of Propellant Slosh in a Spherical Tank

NASA Technical Reports Server (NTRS)

Benson, David J.; Mason, Paul A.

2011-01-01

Propellant sloshing can impart unwanted disturbances to spacecraft, especially if the spacecraft controller is driving the system at the slosh frequency. This paper describes the work performed by the authors in simulating propellant slosh in a spherical tank using computational fluid dynamics (CFD). ANSYS-CFX is the CFD package used to perform the analysis. A 42 in spherical tank is studied with various fill fractions. Results are provided for the forces on the walls and the frequency of the slosh. Snapshots of slosh animation give a qualitative understanding of the propellant slosh. The results show that maximum slosh forces occur at a tank fill fraction of 0.4 and 0.6 due to the amount of mass participating in the slosh and the room available for sloshing to occur. The slosh frequency increases as the tank fill fraction increases.

Long Duration Exposure Facility post-flight data as it influences the Tropical Rainfall Measuring Mission

NASA Technical Reports Server (NTRS)

Straka, Sharon A.

1995-01-01

The Tropical Rainfall Measuring Mission (TRMM) is an earth observing satellite that will be in a low earth orbit (350 kilometers) during the next period of maximum solar activity. The TRMM observatory is expected to experience an atomic oxygen fluence of 8.9 x 10(exp 22) atoms per square centimeter. This fluence is ten times higher than the atomic oxygen impingement incident to the Long Duration Exposure Facility (LDEF). Other environmental concerns on TRMM include: spacecraft glow, silicon oxide contaminant build-up, severe spacecraft material degradation, and contamination deposition resulting from molecular interactions with the dense ambient atmosphere. Because of TRMM's predicted harsh environment, TRMM faces many unique material concerns and subsystem design issues. The LDEF data has influenced the design of TRMM and the TRMM material selection process.

Experimental evaluation of model predictive control and inverse dynamics control for spacecraft proximity and docking maneuvers

NASA Astrophysics Data System (ADS)

Virgili-Llop, Josep; Zagaris, Costantinos; Park, Hyeongjun; Zappulla, Richard; Romano, Marcello

2018-03-01

An experimental campaign has been conducted to evaluate the performance of two different guidance and control algorithms on a multi-constrained docking maneuver. The evaluated algorithms are model predictive control (MPC) and inverse dynamics in the virtual domain (IDVD). A linear-quadratic approach with a quadratic programming solver is used for the MPC approach. A nonconvex optimization problem results from the IDVD approach, and a nonlinear programming solver is used. The docking scenario is constrained by the presence of a keep-out zone, an entry cone, and by the chaser's maximum actuation level. The performance metrics for the experiments and numerical simulations include the required control effort and time to dock. The experiments have been conducted in a ground-based air-bearing test bed, using spacecraft simulators that float over a granite table.

Dynamical Models for Sloshing Dynamics of Helium 2 Under Low-G Conditions

NASA Technical Reports Server (NTRS)

Hung, R. J.; Long, Y. T.

1997-01-01

Coupling of sloshing dynamics within a partially filled rotating dewar of superfluid helium 2 with spacecraft dynamics are investigated in response to the realistic environmental disturbance forces and torques acting on the spacecraft during normal operation. This study investigates: (1) the rotating bubble of superfluid helium 2 reacting to combined environmental disturbances, including gravity gradient, aerodynamic, and magnetic forces and torques; (2) characteristics of slosh reaction forces and torques coupling with spacecraft dynamics; (3) the contribution of slosh dynamics to over-all spacecraft dynamics; and (4) activating of attitude and translation control system. The numerical computation of sloshing dynamics is based on the rotational frame, while the spacecraft dynamics is associated with non-rotational frame. Results show that the contributions of spacecraft dynamics are driven by the environmental disturbances coupling with slosh dynamics. Without considering the effects of environmental disturbances-driven slosh dynamics acting on spacecraft coupling with the spacecraft dynamics may lead to the wrong results for the development of spacecraft system guidance and attitude control techniques.

Deep Impact Sequence Planning Using Multi-Mission Adaptable Planning Tools With Integrated Spacecraft Models

NASA Technical Reports Server (NTRS)

Wissler, Steven S.; Maldague, Pierre; Rocca, Jennifer; Seybold, Calina

2006-01-01

The Deep Impact mission was ambitious and challenging. JPL's well proven, easily adaptable multi-mission sequence planning tools combined with integrated spacecraft subsystem models enabled a small operations team to develop, validate, and execute extremely complex sequence-based activities within very short development times. This paper focuses on the core planning tool used in the mission, APGEN. It shows how the multi-mission design and adaptability of APGEN made it possible to model spacecraft subsystems as well as ground assets throughout the lifecycle of the Deep Impact project, starting with models of initial, high-level mission objectives, and culminating in detailed predictions of spacecraft behavior during mission-critical activities.

Open-systems Architecture of a Standardized Command Interface Chip-set for Switching and Control of a Spacecraft Power Bus

NASA Technical Reports Server (NTRS)

Ruiz, B. Ian; Burke, Gary R.; Lung, Gerald; Whitaker, William D.; Nowicki, Robert M.

2004-01-01

This viewgraph presentation reviews the architecture of the The CIA-AlA chip-set is a set of mixed-signal ASICs that provide a flexible high level interface between the spacecraft's command and data handling (C&DH) electronics and lower level functions in other spacecraft subsystems. Due to the open-systems architecture of the chip-set including an embedded micro-controller a variety of applications are possible. The chip-set was developed for the missions to the outer planets. The chips were developed to provide a single solution for both the switching and regulation of a spacecraft power bus. The Open-Systems Architecture allows for other powerful applications.

AIAA Survivability Technical Committee Draft

NASA Technical Reports Server (NTRS)

Shipman, Jim; Williamson, Joel

1997-01-01

A relatively new area of interest in aerospace systems survivability is the growing threat of spacecraft penetration by orbital debris. Orbital debris, or "space junk", is composed of the man-made remnants of non-functioning spacecraft still orbiting the Earth. NASA estimates that there are currently over 100,000 orbital debris particles 1 centimeter in diameter or larger that cannot be tracked by existing radar, with the population growing at approximately 4% per year in low earth orbits. With an average velocity of over 8.7 km/sec, these projectiles can penetrate and disable many vulnerable spacecraft systems. Since the likelihood of spacecraft penetration increases with spacecraft surface area, large spacecraft (such as the International Space Station) and communication satellite fleets (such as Iridium) have begun to adopt survivability enhancement strategies similar to those employed by combat aircraft. Collision avoidance maneuvers are commonly practiced by the Space Shuttle and are planned by the International Space Station to decrease their susceptibility to impact by trackable orbital debris; likewise, improved shielding, internal equipment placement, and improved crew operations following penetration can reduce the vulnerability of spacecraft to loss following orbital debris impact. Computer simulations such as the Manned Spacecraft and Crew Survivability (MSCSurv) program at the NASA-Marshall Space Flight Center have recently been developed to quantify and reduce the likelihood of crew or spacecraft loss following orbital debris penetration. The AIAA Survivability Technical Committee is working to enable the transfer of military-developed survivability technologies to help the aerospace industry cope with this growing threat.

Using SFOC to fly the Magellan Venus mapping mission

NASA Technical Reports Server (NTRS)

Bucher, Allen W.; Leonard, Robert E., Jr.; Short, Owen G.

1993-01-01

Traditionally, spacecraft flight operations at the Jet Propulsion Laboratory (JPL) have been performed by teams of spacecraft experts utilizing ground software designed specifically for the current mission. The Jet Propulsion Laboratory set out to reduce the cost of spacecraft mission operations by designing ground data processing software that could be used by multiple spacecraft missions, either sequentially or concurrently. The Space Flight Operations Center (SFOC) System was developed to provide the ground data system capabilities needed to monitor several spacecraft simultaneously and provide enough flexibility to meet the specific needs of individual projects. The Magellan Spacecraft Team utilizes the SFOC hardware and software designed for engineering telemetry analysis, both real-time and non-real-time. The flexibility of the SFOC System has allowed the spacecraft team to integrate their own tools with SFOC tools to perform the tasks required to operate a spacecraft mission. This paper describes how the Magellan Spacecraft Team is utilizing the SFOC System in conjunction with their own software tools to perform the required tasks of spacecraft event monitoring as well as engineering data analysis and trending.

Spacecraft System Failures and Anomalies Attributed to the Natural Space Environment

NASA Technical Reports Server (NTRS)

Bedingfield, Keith, L.; Leach, Richard D.; Alexander, Margaret B. (Editor)

1996-01-01

The natural space environment is characterized by many complex and subtle phenomena hostile to spacecraft. The effects of these phenomena impact spacecraft design, development, and operations. Space systems become increasingly susceptible to the space environment as use of composite materials and smaller, faster electronics increases. This trend makes an understanding of the natural space environment essential to accomplish overall mission objectives, especially in the current climate of better/cheaper/faster. This primer provides a brief overview of the natural space environment - definition, related programmatic issues, and effects on various spacecraft subsystems. The primary focus, however, is to catalog, through representative case histories, spacecraft failures and anomalies attributed to the natural space environment. This primer is one in a series of NASA Reference Publications currently being developed by the Electromagnetics and Aerospace Environments Branch, Systems Analysis and Integration Laboratory, Marshall Space Flight Center (MSFC), National Aeronautics and Space Administration (NASA).

Development and experimentation of LQR/APF guidance and control for autonomous proximity maneuvers of multiple spacecraft

NASA Astrophysics Data System (ADS)

Bevilacqua, R.; Lehmann, T.; Romano, M.

2011-04-01

This work introduces a novel control algorithm for close proximity multiple spacecraft autonomous maneuvers, based on hybrid linear quadratic regulator/artificial potential function (LQR/APF), for applications including autonomous docking, on-orbit assembly and spacecraft servicing. Both theoretical developments and experimental validation of the proposed approach are presented. Fuel consumption is sub-optimized in real-time through re-computation of the LQR at each sample time, while performing collision avoidance through the APF and a high level decisional logic. The underlying LQR/APF controller is integrated with a customized wall-following technique and a decisional logic, overcoming problems such as local minima. The algorithm is experimentally tested on a four spacecraft simulators test bed at the Spacecraft Robotics Laboratory of the Naval Postgraduate School. The metrics to evaluate the control algorithm are: autonomy of the system in making decisions, successful completion of the maneuver, required time, and propellant consumption.

Global Precipitation Measurement (GPM) Spacecraft Lithium Ion Battery Micro-Cycling Investigation

NASA Technical Reports Server (NTRS)

Dakermanji, George; Lee, Leonine; Spitzer, Thomas

2016-01-01

The Global Precipitation Measurement (GPM) spacecraft was jointly developed by NASA and JAXA. It is a Low Earth Orbit (LEO) spacecraft launched on February 27, 2014. The power system is a Direct Energy Transfer (DET) system designed to support 1950 watts orbit average power. The batteries use SONY 18650HC cells and consist of three 8s by 84p batteries operated in parallel as a single battery. During instrument integration with the spacecraft, large current transients were observed in the battery. Investigation into the matter traced the cause to the Dual-Frequency Precipitation Radar (DPR) phased array radar which generates cyclical high rate current transients on the spacecraft power bus. The power system electronics interaction with these transients resulted in the current transients in the battery. An accelerated test program was developed to bound the effect, and to assess the impact to the mission.

Formation Algorithms and Simulation Testbed

NASA Technical Reports Server (NTRS)

Wette, Matthew; Sohl, Garett; Scharf, Daniel; Benowitz, Edward

2004-01-01

Formation flying for spacecraft is a rapidly developing field that will enable a new era of space science. For one of its missions, the Terrestrial Planet Finder (TPF) project has selected a formation flying interferometer design to detect earth-like planets orbiting distant stars. In order to advance technology needed for the TPF formation flying interferometer, the TPF project has been developing a distributed real-time testbed to demonstrate end-to-end operation of formation flying with TPF-like functionality and precision. This is the Formation Algorithms and Simulation Testbed (FAST) . This FAST was conceived to bring out issues in timing, data fusion, inter-spacecraft communication, inter-spacecraft sensing and system-wide formation robustness. In this paper we describe the FAST and show results from a two-spacecraft formation scenario. The two-spacecraft simulation is the first time that precision end-to-end formation flying operation has been demonstrated in a distributed real-time simulation environment.

Spacecraft Charging Calculations: NASCAP-2K and SEE Spacecraft Charging Handbook

NASA Technical Reports Server (NTRS)

Davis, V. A.; Neergaard, L. F.; Mandell, M. J.; Katz, I.; Gardner, B. M.; Hilton, J. M.; Minor, J.

2002-01-01

For fifteen years NASA and the Air Force Charging Analyzer Program for Geosynchronous Orbits (NASCAP/GEO) has been the workhorse of spacecraft charging calculations. Two new tools, the Space Environment and Effects (SEE) Spacecraft Charging Handbook (recently released), and Nascap-2K (under development), use improved numeric techniques and modern user interfaces to tackle the same problem. The SEE Spacecraft Charging Handbook provides first-order, lower-resolution solutions while Nascap-2K provides higher resolution results appropriate for detailed analysis. This paper illustrates how the improvements in the numeric techniques affect the results.

Mathematical modeling of a class of multibody flexible spacecraft structures

NASA Technical Reports Server (NTRS)

Kelkar, Atul, G.

1994-01-01

A mathematical model for a general multibody flexible spacecraft is obtained. The generic spacecraft considered consists of a flexible central body to which a number of flexible multibody structures are attached. The coordinate systems used in the derivation allow effective decoupling of the translational motion of the entire spacecraft from its rotational motion about its center of mass. The derivation assumes that the deformations in the bodies are only due to elastic motions. The dynamic model derived is a closed-form vector-matrix differential equation. The model developed can be used for analysis and simulation of many realistic spacecraft configurations.

Electrostatic Return of Contaminants

NASA Technical Reports Server (NTRS)

Rantanen, R.; Gordon, T.

2003-01-01

A Model has been developed capable of calculating the electrostatic return of spacecraft-emitted molecules that are ionized and attracted back to the spacecraft by the spacecraft electric potential on its surfaces. The return of ionized contaminant molecules to charged spacecraft surfaces is very important to all altitudes. It is especially important at geosynchronous and interplanetary environments, since it may be the only mechanism by which contaminants can degrade a surface. This model is applicable to all altitudes and spacecraft geometries. In addition to results of the model will be completed to cover a wide range of potential space systems.

Modular design attitude control system

NASA Technical Reports Server (NTRS)

Chichester, F. D.

1982-01-01

A hybrid multilevel linear quadratic regulator (ML-LQR) approach was developed and applied to the attitude control of models of the rotational dynamics of a prototype flexible spacecraft and of a typical space platform. Three axis rigid body flexible suspension models were developed for both the spacecraft and the space platform utilizing augmented body methods. Models of the spacecraft with hybrid ML-LQR attitude control and with LQR attitude control were simulated and their response with the two different types of control were compared.

Advanced thermal control technology for commercial applications

NASA Technical Reports Server (NTRS)

Swanson, Theodore D.

1991-01-01

A number of the technologies previously developed for the thermal control of spacecraft have found their way into commercial application. Specialized coatings and heat pipes are but two examples. The thermal control of current and future spacecraft is becoming increasingly more demanding, and a variety of new technologies are being developed to meet these needs. Closed two-phase loops are perceived to be the answer to many of the new requirements. All of these technologies are discussed, and their spacecraft and current terrestrial applications are summarized.

Characterization of Magnetospheric Spacecraft Charging Environments Using the LANL Magnetospheric Plasma Analyzer Data Set

NASA Technical Reports Server (NTRS)

Hardage, Donna (Technical Monitor); Davis, V. A.; Mandell, M. J.; Thomsen, M. F.

2003-01-01

An improved specification of the plasma environment has been developed for use in modeling spacecraft charging. It was developed by statistically analyzing a large part of the LANL Magnetospheric Plasma Analyzer (MPA) data set for ion and electron spectral signature correlation with spacecraft charging, including anisotropies. The objective is to identify a relatively simple characterization of the full particle distributions that yield an accurate predication of the observed charging under a wide variety of conditions.

An algorithm for enhanced formation flying of satellites in low earth orbit

NASA Astrophysics Data System (ADS)

Folta, David C.; Quinn, David A.

1998-01-01

With scientific objectives for Earth observation programs becoming more ambitious and spacecraft becoming more autonomous, the need for innovative technical approaches on the feasibility of achieving and maintaining formations of spacecraft has come to the forefront. The trend to develop small low-cost spacecraft has led many scientists to recognize the advantage of flying several spacecraft in formation to achieve the correlated instrument measurements formerly possible only by flying many instruments on a single large platform. Yet, formation flying imposes additional complications on orbit maintenance, especially when each spacecraft has its own orbit requirements. However, advances in automation and technology proposed by the Goddard Space Flight Center (GSFC) allow more of the burden in maneuver planning and execution to be placed onboard the spacecraft, mitigating some of the associated operational concerns. The purpose of this paper is to present GSFC's Guidance, Navigation, and Control Center's (GNCC) algorithm for Formation Flying of the low earth orbiting spacecraft that is part of the New Millennium Program (NMP). This system will be implemented as a close-loop flight code onboard the NMP Earth Orbiter-1 (EO-1) spacecraft. Results of this development can be used to determine the appropriateness of formation flying for a particular case as well as operational impacts. Simulation results using this algorithm integrated in an autonomous `fuzzy logic' control system called AutoCon™ are presented.

Enhancing the Cassini Mission Through FP Applications After Launch

NASA Technical Reports Server (NTRS)

Morgan, Paula S.

2016-01-01

Although rigorous pre-emptive measures are taken to preclude failures and anomalous conditions from occurring in JPL spacecraft missions prior to launch, unforeseeable problems can still surface after liftoff. In the case of the Cassini/Huygens Mission-to-Saturn spacecraft, several problems were observed post-launch: 1) immediately after takeoff, the collected engineering/science data stored on the Solid State Recorders (SSR) contained a significantly higher number of corrupted bits than was expected (considerably over spec) due to human error in the memory mapping of these devices, 2) numerous Solid State Power Switches (SSPS) sporadically tripped off throughout the mission due to cosmic ray bombardment from the unique space environment, and 3) false assumptions in the pressure regulator design in combination with missing heritage test data led to inaccurate design conclusions, causing the issuance of two waivers for the regulator to close properly (a potentially mission catastrophic single-point failure which occurred 24 days after launch) - amongst other problems. For Cassini, some of these anomalies led to arduous work-arounds or required continuous monitoring of telemetry variables by the ground-based Spacecraft Operations Flight Support (SOFS) team in order to detect and fix fault occurrences as they happened. Fortunately, sufficient funding and schedule margin allowed several Fault Protection (FP) solutions to be implemented into post-launch Flight Software (FSW) uploads to help resolve these issues autonomously, reducing SOFS ground support efforts while improving anomaly recovery time in order to preserve maximum science capture. This paper details the FP applications used to resolve the above issues as well as to optimize solutions for several other problems experienced by the Cassini spacecraft during its fight, in order to enhance the spacecraft's overall mission success throughout the 18 years of its 20 year expedition to and within the Saturnian system.

A Comparison of Structurally Connected and Multiple Spacecraft Interferometers

NASA Technical Reports Server (NTRS)

Surka, Derek M.; Crawley, Edward F.

1996-01-01

Structurally connected and multiple spacecraft interferometers are compared in an attempt to establish the maximum baseline (referred to as the "cross-over baseline") for which it is preferable to operate a single-structure interferometer in space rather than an interferometer composed of numerous, smaller spacecraft. This comparison is made using the total launched mass of each configuration as the comparison metric. A framework of study within which structurally connected and multiple spacecraft interferometers can be compared is presented in block diagram form. This methodology is then applied to twenty-two different combinations of trade space parameters to investigate the effects of different orbits, orientations, truss materials, propellants, attitude control actuators, onboard disturbance sources, and performance requirements on the cross-over baseline. Rotating interferometers and the potential advantages of adding active structural control to the connected truss of the structurally connected interferometer are also examined. The minimum mass design of the structurally connected interferometer that meets all performance-requirements and satisfies all imposed constraints is determined as a function of baseline. This minimum mass design is then compared to the design of the multiple spacecraft interferometer. It is discovered that the design of the minimum mass structurally connected interferometer that meets all performance requirements and constraints in solar orbit is limited by the minimum allowable aspect ratio, areal density, and gage of the struts. In the formulation of the problem used in this study, there is no advantage to adding active structural control to the truss for interferometers in solar orbit. The cross-over baseline for missions of practical duration (ranging from one week to thirty years) in solar orbit is approximately 400 m for non-rotating interferometers and 650 m for rotating interferometers.

Modeling and Analysis of Realistic Fire Scenarios in Spacecraft

NASA Technical Reports Server (NTRS)

Brooker, J. E.; Dietrich, D. L.; Gokoglu, S. A.; Urban, D. L.; Ruff, G. A.

2015-01-01

An accidental fire inside a spacecraft is an unlikely, but very real emergency situation that can easily have dire consequences. While much has been learned over the past 25+ years of dedicated research on flame behavior in microgravity, a quantitative understanding of the initiation, spread, detection and extinguishment of a realistic fire aboard a spacecraft is lacking. Virtually all combustion experiments in microgravity have been small-scale, by necessity (hardware limitations in ground-based facilities and safety concerns in space-based facilities). Large-scale, realistic fire experiments are unlikely for the foreseeable future (unlike in terrestrial situations). Therefore, NASA will have to rely on scale modeling, extrapolation of small-scale experiments and detailed numerical modeling to provide the data necessary for vehicle and safety system design. This paper presents the results of parallel efforts to better model the initiation, spread, detection and extinguishment of fires aboard spacecraft. The first is a detailed numerical model using the freely available Fire Dynamics Simulator (FDS). FDS is a CFD code that numerically solves a large eddy simulation form of the Navier-Stokes equations. FDS provides a detailed treatment of the smoke and energy transport from a fire. The simulations provide a wealth of information, but are computationally intensive and not suitable for parametric studies where the detailed treatment of the mass and energy transport are unnecessary. The second path extends a model previously documented at ICES meetings that attempted to predict maximum survivable fires aboard space-craft. This one-dimensional model implies the heat and mass transfer as well as toxic species production from a fire. These simplifications result in a code that is faster and more suitable for parametric studies (having already been used to help in the hatch design of the Multi-Purpose Crew Vehicle, MPCV).

Minimum-Time and Vibration Avoidance Attitude Maneuver for Spacecraft with Torque and Momentum Limit Constraints in Redundant Reaction Wheel Configuration

NASA Technical Reports Server (NTRS)

Ha, Kong Q.; Femiano, Michael D.; Mosier, Gary E.

2004-01-01

In this paper, we present an optimal open-loop slew trajectory algorithm developed at GSFC for the so-called "Yardstick design" of the James Webb Space Telescope (JWST). JWST is an orbiting infrared observatory featuring a lightweight, segmented primary mirror approximately 6 meters in diameter and a sunshield approximately the size of a tennis court. This large, flexible structure will have significant number of lightly damped, dominant flexible modes. With very stringent requirements on pointing accuracy and image quality, it is important that slewing be done within the required time constraint and with minimal induced vibration in order to maximize observing efficiency. With reaction wheels as control actuators, initial wheel speeds as well as individual wheel torque and momentum limits become dominant constraints in slew performance. These constraints must be taken into account when performing slews to ensure that unexpected reaction wheel saturation does not occur, since such saturation leads to control failure in accurately tracking commanded motion and produces high frequency torque components capable of exciting structural modes. A minimum-time constraint is also included and coupled with reaction wheel limit constraints in the optimization to minimize both the effect of the control torque on the flexible body motion and the maneuver time. The optimization is on slew command parameters, such as maximum slew velocity and acceleration, for a given redundant reaction wheel configuration and is based on the dynamic interaction between the spacecraft and reaction wheel motion. Analytical development of the slew algorithm to generate desired slew position, rate, and acceleration profiles to command a feedback/feed forward control system is described. High-fidelity simulation and experimental results are presented to show that the developed slew law achieves the objectives.

Methodology for Developing a Probabilistic Risk Assessment Model of Spacecraft Rendezvous and Dockings

NASA Technical Reports Server (NTRS)

Farnham, Steven J., II; Garza, Joel, Jr.; Castillo, Theresa M.; Lutomski, Michael

2011-01-01

In 2007 NASA was preparing to send two new visiting vehicles carrying logistics and propellant to the International Space Station (ISS). These new vehicles were the European Space Agency s (ESA) Automated Transfer Vehicle (ATV), the Jules Verne, and the Japanese Aerospace and Explorations Agency s (JAXA) H-II Transfer Vehicle (HTV). The ISS Program wanted to quantify the increased risk to the ISS from these visiting vehicles. At the time, only the Shuttle, the Soyuz, and the Progress vehicles rendezvoused and docked to the ISS. The increased risk to the ISS was from an increase in vehicle traffic, thereby, increasing the potential catastrophic collision during the rendezvous and the docking or berthing of the spacecraft to the ISS. A universal method of evaluating the risk of rendezvous and docking or berthing was created by the ISS s Risk Team to accommodate the increasing number of rendezvous and docking or berthing operations due to the increasing number of different spacecraft, as well as the future arrival of commercial spacecraft. Before the first docking attempt of ESA's ATV and JAXA's HTV to the ISS, a probabilistic risk model was developed to quantitatively calculate the risk of collision of each spacecraft with the ISS. The 5 rendezvous and docking risk models (Soyuz, Progress, Shuttle, ATV, and HTV) have been used to build and refine the modeling methodology for rendezvous and docking of spacecrafts. This risk modeling methodology will be NASA s basis for evaluating the addition of future ISS visiting spacecrafts hazards, including SpaceX s Dragon, Orbital Science s Cygnus, and NASA s own Orion spacecraft. This paper will describe the methodology used for developing a visiting vehicle risk model.

Developing planetary protection technology- microbial diversity of the Mars Orbiter Odyssey and the spacecraft assembly and encapsulation facility II

NASA Technical Reports Server (NTRS)

Duc, M. La; Chen, F.; Kern, R.; Koukol, R.; Baker, A.; Venkateswaran, K.

2001-01-01

A study in which several surface samples, retrieved from both the Mars Odyssey Spacecraft and the Kennedy Space Center (KSC) Spacecraft Assembly and Encapsulation Facility II (SAEF-II), were prcesed and evaluated by both molecular and traditional culture-based methods for the microbial diversity.

Hardware Simulations of Spacecraft Attitude Synchronization Using Lyapunov-Based Controllers

NASA Astrophysics Data System (ADS)

Jung, Juno; Park, Sang-Young; Eun, Youngho; Kim, Sung-Woo; Park, Chandeok

2018-04-01

In the near future, space missions with multiple spacecraft are expected to replace traditional missions with a single large spacecraft. These spacecraft formation flying missions generally require precise knowledge of relative position and attitude between neighboring agents. In this study, among the several challenging issues, we focus on the technique to control spacecraft attitude synchronization in formation. We develop a number of nonlinear control schemes based on the Lyapunov stability theorem and considering special situations: full-state feedback control, full-state feedback control with unknown inertia parameters, and output feedback control without angular velocity measurements. All the proposed controllers offer absolute and relative control using reaction wheel assembly for both regulator and tracking problems. In addition to the numerical simulations, an air-bearing-based hardware-in-the-loop (HIL) system is used to verify the proposed control laws in real-time hardware environments. The pointing errors converge to 0.5{°} with numerical simulations and to 2{°} using the HIL system. Consequently, both numerical and hardware simulations confirm the performance of the spacecraft attitude synchronization algorithms developed in this study.

Dynamic Modeling of Ascent Abort Scenarios for Crewed Launches

NASA Technical Reports Server (NTRS)

Bigler, Mark; Boyer, Roger L.

2015-01-01

For the last 30 years, the United States' human space program has been focused on low Earth orbit exploration and operations with the Space Shuttle and International Space Station programs. After over 40 years, the U.S. is again working to return humans beyond Earth orbit. To do so, NASA is developing a new launch vehicle and spacecraft to provide this capability. The launch vehicle is referred to as the Space Launch System (SLS) and the spacecraft is called Orion. The new launch system is being developed with an abort system that will enable the crew to escape launch failures that would otherwise be catastrophic as well as probabilistic design requirements set for probability of loss of crew (LOC) and loss of mission (LOM). In order to optimize the risk associated with designing this new launch system, as well as verifying the associated requirements, NASA has developed a comprehensive Probabilistic Risk Assessment (PRA) of the integrated ascent phase of the mission that includes the launch vehicle, spacecraft and ground launch facilities. Given the dynamic nature of rocket launches and the potential for things to go wrong, developing a PRA to assess the risk can be a very challenging effort. Prior to launch and after the crew has boarded the spacecraft, the risk exposure time can be on the order of three hours. During this time, events may initiate from either the spacecraft, the launch vehicle, or the ground systems, thus requiring an emergency egress from the spacecraft to a safe ground location or a pad abort via the spacecraft's launch abort system. Following launch, again either the spacecraft or the launch vehicle can initiate the need for the crew to abort the mission and return home. Obviously, there are thousands of scenarios whose outcome depends on when the abort is initiated during ascent and how the abort is performed. This includes modeling the risk associated with explosions and benign system failures that require aborting a spacecraft under very dynamic conditions, particularly in the lower atmosphere, and returning the crew home safely. This paper will provide an overview of the PRA model that has been developed of this new launch system, including some of the challenges that are associated with this effort.

On determining fluxgate magnetometer spin axis offsets from mirror mode observations

NASA Astrophysics Data System (ADS)

Plaschke, Ferdinand; Narita, Yasuhito

2016-09-01

In-flight calibration of fluxgate magnetometers that are mounted on spacecraft involves finding their outputs in vanishing ambient fields, the so-called magnetometer offsets. If the spacecraft is spin-stabilized, then the spin plane components of these offsets can be relatively easily determined, as they modify the spin tone content in the de-spun magnetic field data. The spin axis offset, however, is more difficult to determine. Therefore, usually Alfvénic fluctuations in the solar wind are used. We propose a novel method to determine the spin axis offset: the mirror mode method. The method is based on the assumption that mirror mode fluctuations are nearly compressible such that the maximum variance direction is aligned to the mean magnetic field. Mirror mode fluctuations are typically found in the Earth's magnetosheath region. We introduce the method and provide a first estimate of its accuracy based on magnetosheath observations by the THEMIS-C spacecraft. We find that 20 h of magnetosheath measurements may already be sufficient to obtain high-accuracy spin axis offsets with uncertainties on the order of a few tenths of a nanotesla, if offset stability can be assumed.

Top of the World

NASA Image and Video Library

2017-08-28

These turbulent clouds are on top of the world at Saturn. NASA's Cassini spacecraft captured this view of Saturn's north pole on April 26, 2017 - the day it began its Grand Finale -- as it approached the planet for its first daring dive through the gap between the planet and its rings. Although the pole is still bathed in sunlight at present, northern summer solstice on Saturn occurred on May 24, 2017, bringing the maximum solar illumination to the north polar region. Now the Sun begins its slow descent in the northern sky, which eventually will plunge the north pole into Earth-years of darkness. Cassini's long mission at Saturn enabled the spacecraft to see the Sun rise over the north, revealing that region in great detail for the first time. This view looks toward the sunlit side of the rings from about 44 degrees above the ring plane. The image was taken with the Cassini spacecraft wide-angle camera using a spectral filter which preferentially admits wavelengths of near-infrared light centered at 752 nanometers. The view was obtained at a distance of approximately 166,000 miles (267,000 kilometers) from Saturn. Image scale is about 10 miles (16 kilometers) per pixel. https://photojournal.jpl.nasa.gov/catalog/PIA21343

Observation of Three-Dimensional Magnetic Reconnection in the Terrestrial Magnetotail

NASA Astrophysics Data System (ADS)

Zhou, Meng; Ashour-Abdalla, Maha; Deng, Xiaohua; Pang, Ye; Fu, Huishan; Walker, Raymond; Lapenta, Giovanni; Huang, Shiyong; Xu, Xiaojun; Tang, Rongxin

2017-09-01

Study of magnetic reconnection has been focused on two-dimensional geometry in the past decades, whereas three-dimensional structures and dynamics of reconnection X line are poorly understood. In this paper, we report Cluster multispacecraft observations of a three-dimensional magnetic reconnection X line with a weak guide field ( 25% of the upstream magnetic field) in the Earth's magnetotail. We find that the X line not only retreated tailward but also expanded across the tail following the electron flow direction with a maximum average speed of (0.04-0.15) VA,up, where VA,up is the upstream Alfvén speed, or (0.14-0.57) Vde, where Vde is the electron flow speed in the out-of-plane direction. An ion diffusion region was observed by two spacecraft that were separated about 10 ion inertial lengths along the out-of-plane direction; however, these two spacecraft observed distinct magnetic structures associated with reconnection: one spacecraft observed dipolarization fronts, while the other one observed flux ropes. This indicates that reconnection proceeds in drastically different ways in different segments along the X line only a few ion inertial lengths apart.

Evaluation of TDRSS-user orbit determination accuracy using batch least-squares and sequential methods

NASA Technical Reports Server (NTRS)

Oza, D. H.; Jones, T. L.; Hodjatzadeh, M.; Samii, M. V.; Doll, C. E.; Hart, R. C.; Mistretta, G. D.

1991-01-01

The development of the Real-Time Orbit Determination/Enhanced (RTOD/E) system as a prototype system for sequential orbit determination on a Disk Operating System (DOS) based Personal Computer (PC) is addressed. The results of a study to compare the orbit determination accuracy of a Tracking and Data Relay Satellite System (TDRSS) user spacecraft obtained using RTOD/E with the accuracy of an established batch least squares system, the Goddard Trajectory Determination System (GTDS), is addressed. Independent assessments were made to examine the consistencies of results obtained by the batch and sequential methods. Comparisons were made between the forward filtered RTOD/E orbit solutions and definitive GTDS orbit solutions for the Earth Radiation Budget Satellite (ERBS); the maximum solution differences were less than 25 m after the filter had reached steady state.

Apollo Program Summary Report: Synopsis of the Apollo Program Activities and Technology for Lunar Exploration

NASA Technical Reports Server (NTRS)

1975-01-01

Overall program activities and the technology developed to accomplish lunar exploration are discussed. A summary of the flights conducted over an 11-year period is presented along with specific aspects of the overall program, including lunar science, vehicle development and performance, lunar module development program, spacecraft development testing, flight crew summary, mission operations, biomedical data, spacecraft manufacturing and testing, launch site facilities, equipment, and prelaunch operations, and the lunar receiving laboratory. Appendixes provide data on each of the Apollo missions, mission type designations, spacecraft weights, records achieved by Apollo crewmen, vehicle histories, and a listing of anomalous hardware conditions noted during each flight beginning with Apollo 4.

Helioseismic inferences of the solar cycles 23 and 24: GOLF and VIRGO observations

NASA Astrophysics Data System (ADS)

Salabert, D.; García, R. A.; Jiménez, A.

2014-12-01

The Sun-as-a star helioseismic spectrophotometer GOLF and photometer VIRGO instruments onboard the SoHO spacecraft are collecting high-quality, continuous data since April 1996. We analyze here these unique datasets in order to investigate the peculiar and weak on-going solar cycle 24. As this cycle 24 is reaching its maximum, we compare its rising phase with the rising phase of the previous solar cycle 23.

Time Exceedances for High Intensity Solar Proton Fluxes

NASA Technical Reports Server (NTRS)

Xapsos, Michael A.; Stauffer, Craig A.; Jordan, Thomas M.; Adam, James H., Jr.; Dietrich, William F.

2011-01-01

A model is presented for times during a space mission that specified solar proton flux levels are exceeded. This includes both total time and continuous time periods during missions. Results for the solar maximum and solar minimum phases of the solar cycle are presented and compared for a broad range of proton energies and shielding levels. This type of approach is more amenable to reliability analysis for spacecraft systems and instrumentation than standard statistical models.

Periods of High Intensity Solar Proton Flux

NASA Technical Reports Server (NTRS)

Xapsos, Michael A.; Stauffer, Craig A.; Jordan, Thomas M.; Adams, James H.; Dietrich, William F.

2012-01-01

Analysis is presented for times during a space mission that specified solar proton flux levels are exceeded. This includes both total time and continuous time periods during missions. Results for the solar maximum and solar minimum phases of the solar cycle are presented and compared for a broad range of proton energies and shielding levels. This type of approach is more amenable to reliability analysis for spacecraft systems and instrumentation than standard statistical models.

Adaptation of G-TAG Software for Validating Touch-and-Go Comet Surface Sampling Design Methodology

NASA Technical Reports Server (NTRS)

Mandic, Milan; Acikmese, Behcet; Blackmore, Lars

2011-01-01

The G-TAG software tool was developed under the R&TD on Integrated Autonomous Guidance, Navigation, and Control for Comet Sample Return, and represents a novel, multi-body dynamics simulation software tool for studying TAG sampling. The G-TAG multi-body simulation tool provides a simulation environment in which a Touch-and-Go (TAG) sampling event can be extensively tested. TAG sampling requires the spacecraft to descend to the surface, contact the surface with a sampling collection device, and then to ascend to a safe altitude. The TAG event lasts only a few seconds but is mission-critical with potentially high risk. Consequently, there is a need for the TAG event to be well characterized and studied by simulation and analysis in order for the proposal teams to converge on a reliable spacecraft design. This adaptation of the G-TAG tool was developed to support the Comet Odyssey proposal effort, and is specifically focused to address comet sample return missions. In this application, the spacecraft descends to and samples from the surface of a comet. Performance of the spacecraft during TAG is assessed based on survivability and sample collection performance. For the adaptation of the G-TAG simulation tool to comet scenarios, models are developed that accurately describe the properties of the spacecraft, approach trajectories, and descent velocities, as well as the models of the external forces and torques acting on the spacecraft. The adapted models of the spacecraft, descent profiles, and external sampling forces/torques were more sophisticated and customized for comets than those available in the basic G-TAG simulation tool. Scenarios implemented include the study of variations in requirements, spacecraft design (size, locations, etc. of the spacecraft components), and the environment (surface properties, slope, disturbances, etc.). The simulations, along with their visual representations using G-View, contributed to the Comet Odyssey New Frontiers proposal effort by indicating problems and/or benefits of different approaches and designs.

Development of a special purpose spacecraft coating, phase 4

NASA Technical Reports Server (NTRS)

Gillman, H. D.

1980-01-01

Coating formulations based on a fluorocarbon resin were evaluated for use on spacecraft exteriors. Formulations modified with an acrylic resin were found to have excellent offgassing properties. A much less expensive process for increasing to solid content of the fluorocarbon latex was developed.

A methodology to select a wire insulation for use in habitable spacecraft.

PubMed

Paulos, T; Apostolakis, G

1998-08-01

This paper investigates electrical overheating events aboard a habitable spacecraft. The wire insulation involved in these failures plays a major role in the entire event scenario from threat development to detection and damage assessment. Ideally, if models of wire overheating events in microgravity existed, the various wire insulations under consideration could be quantitatively compared. However, these models do not exist. In this paper, a methodology is developed that can be used to select a wire insulation that is best suited for use in a habitable spacecraft. The results of this study show that, based upon the Analytic Hierarchy Process and simplifying assumptions, the criteria selected, and data used in the analysis, Tefzel is better than Teflon for use in a habitable spacecraft.

Six degree of freedom simulation system for evaluating automated rendezvous and docking spacecraft

NASA Technical Reports Server (NTRS)

Rourke, Kenneth H.; Tsugawa, Roy K.

1991-01-01

Future logistics supply and servicing vehicles such as cargo transfer vehicles (CTV) must have full 6 degree of freedom (6DOF) capability in order to perform requisite rendezvous, proximity operations, and capture operations. The design and performance issues encountered when developing a 6DOF maneuvering spacecraft are very complex with subtle interactions which are not immediately obvious or easily anticipated. In order to deal with these complexities and develop robust maneuvering spacecraft designs, a simulation system and associated family of tools are used at TRW for generating and validating spacecraft performance requirements and guidance algorithms. An overview of the simulator and tools is provided. These are used by TRW for autonomous rendezvous and docking research projects including CTV studies.

Design and development of a brushless, direct drive solar array reorientation system

NASA Technical Reports Server (NTRS)

Jessee, R. D.

1972-01-01

This report covers the design and development of the laboratory model, and is essentially a compilation of reports covering the system and its various parts. To enhance completeness, the final report of Phase 1 covering circuit development of the controller is also included. A controller was developed for a brushless, direct-drive, single axis solar array reorientation system for earth-pointed, passively-stabilized spacecraft. A control systems was designed and breadboard circuits were built and tested for performance. The controller is designed to take over automatic control of the array on command after the spacecraft is stabilized in orbit. The controller will orient the solar array to the sun vector and automatically track to maintain proper orientation. So long as the orbit is circular, orientation toward the sun is maintained even though the spacecraft goes into the shadow of the earth. Particular attention was given in the design to limit reaction between the array and the spacecraft.

DMSP Auroral Charging at Solar Cycle 24 Maximum

NASA Technical Reports Server (NTRS)

Chandler, Michael; Parker, Linda Neergaard; Minow, Joseph I.

2013-01-01

It has been well established that polar orbiting satellites can experience mild to severe auroral charging levels (on the order of a few hundred volts to few kilovolts negative frame potentials) during solar minimum conditions (Frooninckx and Sojka, 1992; Anderson and Koons, 1996; Anderson, 2012). These same studies have shown a strong reduction in charging during the rising and declining phases of the past few solar cycles with a nearly complete suppression of auroral charging at solar maximum. Recently, we have observed examples of high level charging during the recent approach to Solar Cycle 24 solar maximum conditions not unlike those reported by Frooninckx and Sojka (1992). These observations demonstrate that spacecraft operations during solar maximum cannot be considered safe from auroral charging when solar activity is low. We present a survey of auroral charging events experienced by the Defense Meteorological Satellite Program (DMSP) F16 satellite during Solar Cycle 24 maximum conditions. We summarize the auroral energetic particle environment and the conditions necessary for charging to occur in this environment, we describe how the lower than normal solar activity levels for Solar Cycle 24 maximum conditions are conducive to charging in polar orbits, and we show examples of the more extreme charging events, sometimes exceeding 1 kV, during this time period.

MODEL CORRELATION STUDY OF A RETRACTABLE BOOM FOR A SOLAR SAIL SPACECRAFT

NASA Technical Reports Server (NTRS)

Adetona, O.; Keel, L. H.; Oakley, J. D.; Kappus, K.; Whorton, M. S.; Kim, Y. K.; Rakpczy, J. M.

2005-01-01

To realize design concepts, predict dynamic behavior and develop appropriate control strategies for high performance operation of a solar-sail spacecraft, we developed a simple analytical model that represents dynamic behavior of spacecraft with various sizes. Since motion of the vehicle is dominated by retractable booms that support the structure, our study concentrates on developing and validating a dynamic model of a long retractable boom. Extensive tests with various configurations were conducted for the 30 Meter, light-weight, retractable, lattice boom at NASA MSFC that is structurally and dynamically similar to those of a solar-sail spacecraft currently under construction. Experimental data were then compared with the corresponding response of the analytical model. Though mixed results were obtained, the analytical model emulates several key characteristics of the boom. The paper concludes with a detailed discussion of issues observed during the study.

Embedded Thermal Control for Spacecraft Subsystems Miniaturization

NASA Technical Reports Server (NTRS)

Didion, Jeffrey R.

2014-01-01

Optimization of spacecraft size, weight and power (SWaP) resources is an explicit technical priority at Goddard Space Flight Center. Embedded Thermal Control Subsystems are a promising technology with many cross cutting NSAA, DoD and commercial applications: 1.) CubeSatSmallSat spacecraft architecture, 2.) high performance computing, 3.) On-board spacecraft electronics, 4.) Power electronics and RF arrays. The Embedded Thermal Control Subsystem technology development efforts focus on component, board and enclosure level devices that will ultimately include intelligent capabilities. The presentation will discuss electric, capillary and hybrid based hardware research and development efforts at Goddard Space Flight Center. The Embedded Thermal Control Subsystem development program consists of interrelated sub-initiatives, e.g., chip component level thermal control devices, self-sensing thermal management, advanced manufactured structures. This presentation includes technical status and progress on each of these investigations. Future sub-initiatives, technical milestones and program goals will be presented.

Expert diagnostics system as a part of analysis software for power mission operations

NASA Technical Reports Server (NTRS)

Harris, Jennifer A.; Bahrami, Khosrow A.

1993-01-01

The operation of interplanetary spacecraft at JPL has become an increasingly complex activity. This complexity is due to advanced spacecraft designs and ambitious mission objectives which lead to operations requirements that are more demanding than those of any previous mission. For this reason, several productivity enhancement measures are underway at JPL within mission operations, particularly in the spacecraft analysis area. These measures aimed at spacecraft analysis include: the development of a multi-mission, multi-subsystem operations environment; the introduction of automated tools into this environment; and the development of an expert diagnostics system. This paper discusses an effort to integrate the above mentioned productivity enhancement measures. A prototype was developed that integrates an expert diagnostics system into a multi-mission, multi-subsystem operations environment using the Galileo Power / Pyro Subsystem as a testbed. This prototype will be discussed in addition to background information associated with it.

Development of the Mars Observer Laser Altimeter (MOLA)

NASA Technical Reports Server (NTRS)

Johnson, Bertrand L., Jr

1993-01-01

The Mars Observer (MO) spacecraft payload scientific mission is to gather data on Martian global topography, gravity, weather, magnetic field and its interaction with the solar flux, surface chemistry, and mineralogy over one Mars year. In mid-1988 the need for a replacement altimeter as part of the payload complement arose. The Mars Observer Laser Altimeter (MOLA) was proposed by GSFC as an in-house effort and shortly afterward was 'conditionally' accepted. Constraints on funding, schedule, power, and mass were imposed with periodic reviews during the instrument development to authorize continuation. MOLA was designed, tested, and delivered in less than 36 months and integrated with the spacecraft. During spacecraft payload testing, the laser failed due to contamination in the laser cavity. In only 6 months, the laser was removed, rebuilt from spare parts, retested, and the instrument reassembled, realigned, requalified, and again delivered for spacecraft integration. Other aspects of the development of the MOLA are presented.

Space Environments and Effects (SEE) Program: Spacecraft Charging Technology Development Activities

NASA Technical Reports Server (NTRS)

Kauffman, Billy; Hardage, Donna; Minor, Jody

2003-01-01

Reducing size and weight of spacecraft, along with demanding increased performance capabilities, introduces many uncertainties in the engineering design community on how materials and spacecraft systems will perform in space. The engineering design community is forever behind on obtaining and developing new tools and guidelines to mitigate the harmful effects of the space environment. Adding to this complexity is the continued push to use Commercial-off-the-shelf (COTS) microelectronics, potential usage of unproven technologies such as large solar sail structures and nuclear electric propulsion. In order to drive down these uncertainties, various programs are working together to avoid duplication, save what resources are available in this technical area and possess a focused agenda to insert these new developments into future mission designs. This paper will introduce the SEE Program, briefly discuss past and currently sponsored spacecraft charging activities and possible future endeavors.

Space Environments and Effects (SEE) Program: Spacecraft Charging Technology Development Activities

NASA Technical Reports Server (NTRS)

Kauffman, B.; Hardage, D.; Minor, J.

2004-01-01

Reducing size and weight of spacecraft, along with demanding increased performance capabilities, introduces many uncertainties in the engineering design community on how materials and spacecraft systems will perform in space. The engineering design community is forever behind on obtaining and developing new tools and guidelines to mitigate the harmful effects of the space environment. Adding to this complexity is the continued push to use Commercial-off-the-Shelf (COTS) microelectronics, potential usage of unproven technologies such as large solar sail structures and nuclear electric propulsion. In order to drive down these uncertainties, various programs are working together to avoid duplication, save what resources are available in this technical area and possess a focused agenda to insert these new developments into future mission designs. This paper will introduce the SEE Program, briefly discuss past and currently sponsored spacecraft charging activities and possible future endeavors.

Results of an analytical study of spacecraft deposition contamination by internal reflection spectroscopy. [(haze on spacecraft windows from space debris)

NASA Technical Reports Server (NTRS)

Mookherji, T.

1976-01-01

Outgassing, deposition, and desorption kinetics of silicone compounds, are examined as examples of optical surface contaminants of spacecraft windows. Their behavior in a space environment after exposure to ultraviolet radiation is also examined. The use of internal reflection spectroscopy is shown to provide a viable means of real-time, in-situ identification of contaminants of orbiting spacecraft. The instrumental techniques are proposed as the basis of further investigations and the development of flight hardware.

Additional application of the NASCAP code. Volume 1: NASCAP extension

NASA Technical Reports Server (NTRS)

Katz, I.; Cassidy, J. J.; Mandell, M. J.; Parks, D. E.; Schnuelle, G. W.; Stannard, P. R.; Steen, P. G.

1981-01-01

The NASCAP computer program comprehensively analyzes problems of spacecraft charging. Using a fully three dimensional approach, it can accurately predict spacecraft potentials under a variety of conditions. Several changes were made to NASCAP, and a new code, NASCAP/LEO, was developed. In addition, detailed studies of several spacecraft-environmental interactions and of the SCATHA spacecraft were performed. The NASCAP/LEO program handles situations of relatively short Debye length encountered by large space structures or by any satellite in low earth orbit (LEO).

A design study for the use of a multiple aperture deployable antenna for soil moisture remote sensing satellite applications

NASA Technical Reports Server (NTRS)

Foldes, P.

1986-01-01

The instrumentation problems associated with the measurement of soil moisture with a meaningful spatial and temperature resolution at a global scale are addressed. For this goal only medium term available affordable technology will be considered. The study while limited in scope, will utilize a large scale antenna structure, which is being developed presently as an experimental model. The interface constraints presented by a singel Space Transportation System (STS) flight will be assumed. Methodology consists of the following steps: review of science requirements; analyze effects of these requirements; present basic system engineering considerations and trade-offs related to orbit parameters, number of spacecraft and their lifetime, observation angles, beamwidth, crossover and swath, coverage percentage, beam quality and resolution, instrument quantities, and integration time; bracket the key system characteristics and develop an electromagnetic design of the antenna-passive radiometer system. Several aperture division combinations and feed array concepts are investigated to achieve maximum feasible performacne within the stated STS constraints.

Active spacecraft potential control system selection for the Jupiter orbiter with probe mission

NASA Technical Reports Server (NTRS)

Beattie, J. R.; Goldstein, R.

1977-01-01

It is shown that the high flux of energetic plasma electrons and the reduced photoemission rate in the Jovian environment can result in the spacecraft developing a large negative potential. The effects of the electric fields produced by this charging phenomenon are discussed in terms of spacecraft integrity as well as charged particle and fields measurements. The primary area of concern is shown to be the interaction of the electric fields with the measuring devices on the spacecraft. The need for controlling the potential of the spacecraft is identified, and a system capable of active control of the spacecraft potential in the Jupiter environment is proposed. The desirability of using this system to vary the spacecraft potential relative to the ambient plasma potential is also discussed. Various charged particle release devices are identified as potential candidates for use with the spacecraft potential control system. These devices are evaluated and compared on the basis of system mass, power consumption, and system complexity and reliability.

Modeling SMAP Spacecraft Attitude Control Estimation Error Using Signal Generation Model

NASA Technical Reports Server (NTRS)

Rizvi, Farheen

2016-01-01

Two ground simulation software are used to model the SMAP spacecraft dynamics. The CAST software uses a higher fidelity model than the ADAMS software. The ADAMS software models the spacecraft plant, controller and actuator models, and assumes a perfect sensor and estimator model. In this simulation study, the spacecraft dynamics results from the ADAMS software are used as CAST software is unavailable. The main source of spacecraft dynamics error in the higher fidelity CAST software is due to the estimation error. A signal generation model is developed to capture the effect of this estimation error in the overall spacecraft dynamics. Then, this signal generation model is included in the ADAMS software spacecraft dynamics estimate such that the results are similar to CAST. This signal generation model has similar characteristics mean, variance and power spectral density as the true CAST estimation error. In this way, ADAMS software can still be used while capturing the higher fidelity spacecraft dynamics modeling from CAST software.

The 1980 solar maximum mission event listing

NASA Technical Reports Server (NTRS)

Speich, D. M.; Nelson, J. J.; Licata, J. P.; Tolbert, A. K.

1991-01-01

Information is contained on solar burst and transient activity observed by the Solar Maximum Mission (SMM) during 1980 pointed observations. Data from the following SMM experiments are included: (1) Gamma Ray Spectrometer, (2) Hard X-Ray Burst Spectrometer, (3) Hard X-Ray Imaging Spectrometer, (4) Flat Crystal Spectrometer, (5) Bent Crystal Spectrometer, (6) Ultraviolet Spectrometer and Polarimeter, and (7) Coronagraph/Polarimeter. Correlative optical, radio, and Geostationary Operational Environmental Satellite (GOES) x ray data are also presented. Where possible, bursts or transients observed in the various wavelengths were grouped into discrete flare events identified by unique event numbers. Each event carries a qualifier denoting the quality or completeness of the observations. Spacecraft pointing coordinates and flare site angular displacement values from Sun center are also included.

Solar radiation pressure effects on the Helios spacecraft

NASA Technical Reports Server (NTRS)

Georgevic, R. M.

1976-01-01

A mathematical model of the solar radiation force and torques, developed for the Mariner 10 Venus/Mercury spacecraft mission, was used for a detailed analysis of the effects of solar light pressure on the Helios spacecraft. Due to the fact that the main body of the Helios spacecraft is a surface of enclosure, inside of which most of the reradiated thermal energy is lost, expressions for the portion of the solar radiation force, produced by the thermal reradiation, had to be given a different form. Hence the need for the derivation of a somewhat different theoretical model for the force acting on the main body of the spacecraft.

Nuclear radiation problems, unmanned thermionic reactor ion propulsion spacecraft

NASA Technical Reports Server (NTRS)

Mondt, J. F.; Sawyer, C. D.; Nakashima, A.

1972-01-01

A nuclear thermionic reactor as the electric power source for an electric propulsion spacecraft introduces a nuclear radiation environment that affects the spacecraft configuration, the use and location of electrical insulators and the science experiments. The spacecraft is conceptually configured to minimize the nuclear shield weight by: (1) a large length to diameter spacecraft; (2) eliminating piping penetrations through the shield; and (3) using the mercury propellant as gamma shield. Since the alumina material is damaged by the high nuclear radiation environment in the reactor it is desirable to locate the alumina insulator outside the reflector or develop a more radiation resistant insulator.

A Technique for the Assessment of Flight Operability Characteristics of Human Rated Spacecraft

NASA Technical Reports Server (NTRS)

Crocker, Alan

2010-01-01

In support of new human rated spacecraft development programs, the Mission Operations Directorate at NASA Johnson Space Center has implemented a formal method for the assessment of spacecraft operability. This "Spacecraft Flight Operability Assessment Scale" defines six key themes of flight operability, with guiding principles and goals stated for each factor. A standardized rating technique provides feedback that is useful to the operations, design and program management communities. Applicability of this concept across the program structure and life cycle is addressed. Examples of operationally desirable and undesirable spacecraft design characteristics are provided, as is a sample of the assessment scale product.

Satellite Ground Operations Automation: Lessons Learned and Future Approaches

NASA Technical Reports Server (NTRS)

Catena, John; Frank, Lou; Saylor, Rick; Weikel, Craig; Obenschain, Arthur F. (Technical Monitor)

2001-01-01

Reducing spacecraft ground system operations costs is a major goal in all missions. The Fast Auroral Snapshot (FAST) flight operations team at the NASA/Goddard Spacecraft Flight Center developed in-house scripts and procedures to automate monitoring of critical spacecraft functions. The initial staffing profile of 16x7 was reduced first to 8x5 and then to 'lights out'. Operations functions became an offline review of system performance and the generation of future science plans for subsequent upload to the spacecraft. Lessons learned will be applied to the challenging Triana mission, where 24x7 contact with the spacecraft will be necessary at all times.

Spacecraft level impacts of integrating concentrator solar arrays

DOE Office of Scientific and Technical Information (OSTI.GOV)

Allen, D.M.; Piszczor, M.F. Jr.

1994-12-31

The paper describes the results of a study to determine the impacts of integrating concentrator solar arrays on spacecraft design and performance. First, concentrator array performance is summarized for the AEC-Able/Entech SCARLET array, the Ioffe refractive and reflective concepts being developed in Russia, the Martin Marietta SLATS system, and other concentrator concepts that have been designed or developed. Concentrator array performance is compared to rigid and flex blanket planar array technologies at the array level. Then other impacts on the spacecraft are quantified. Conclusions highlight the most important results as they relate to recommended approaches in developing concentrator arrays formore » satellites.« less

Numerical modeling of on-orbit propellant motion resulting from an impulsive acceleration

NASA Technical Reports Server (NTRS)

Aydelott, John C.; Mjolsness, Raymond C.; Torrey, Martin D.; Hochstein, John I.

1987-01-01

In-space docking and separation maneuvers of spacecraft that have large fluid mass fractions may cause undesirable spacecraft motion in response to the impulsive-acceleration-induced fluid motion. An example of this potential low gravity fluid management problem arose during the development of the shuttle/Centaur vehicle. Experimentally verified numerical modeling techniques were developed to establish the propellant dynamics, and subsequent vehicle motion, associated with the separation of the Centaur vehicle from the shuttle orbiter cargo bay. Although the shuttle/Centaur development activity was suspended, the numerical modeling techniques are available to predict on-orbit liquid motion resulting from impulsive accelerations for other missions and spacecraft.

Solid-Body Fuse Developed for High- Voltage Space Power Missions

NASA Technical Reports Server (NTRS)

Dolce, James L.; Baez, Anastacio N.

2001-01-01

AEM Incorporated has completed the development, under a NASA Glenn Research Center contract, of a solid-body fuse for high-voltage power systems of satellites and spacecraft systems. High-reliability fuses presently defined by MIL-PRF-23419 do not meet the increased voltage and amperage requirements for the next generation of spacecraft. Solid-body fuses exhibit electrical and mechanical attributes that enable these fuses to perform reliably in the vacuum and high-vibration and -shock environments typically present in spacecraft applications. The construction and screening techniques for solid-body fuses described by MIL-PRF-23419/12 offer an excellent roadmap for the development of high-voltage solid-body fuses.

A unified approach to computer analysis and modeling of spacecraft environmental interactions

NASA Technical Reports Server (NTRS)

Katz, I.; Mandell, M. J.; Cassidy, J. J.

1986-01-01

A new, coordinated, unified approach to the development of spacecraft plasma interaction models is proposed. The objective is to eliminate the unnecessary duplicative work in order to allow researchers to concentrate on the scientific aspects. By streamlining the developmental process, the interchange between theories and experimentalists is enhanced, and the transfer of technology to the spacecraft engineering community is faster. This approach is called the UNIfied Spacecraft Interaction Model (UNISIM). UNISIM is a coordinated system of software, hardware, and specifications. It is a tool for modeling and analyzing spacecraft interactions. It will be used to design experiments, to interpret results of experiments, and to aid in future spacecraft design. It breaks a Spacecraft Ineraction analysis into several modules. Each module will perform an analysis for some physical process, using phenomenology and algorithms which are well documented and have been subject to review. This system and its characteristics are discussed.

The Human as a System - Monitoring Spacecraft Net Habitable Volume throughout the Design Lifecycle

NASA Technical Reports Server (NTRS)

Szabo, Richard; Kallay, Anna; Twyford, Evan; Maida, Jim

2007-01-01

Spacecraft design has historically allocated specific volume and mass "not to exceed" requirements upon individual systems and their accompanying hardware (e.g., life support, avionics) early in their conceptual design in an effort to align the spacecraft with propulsion capabilities. If the spacecraft is too heavy or too wide for the launch stack - it does not get off the ground. This approach has predictably ended with the crew being allocated whatever open, pressurized volume remains. With the recent inauguration of a new human-rated spacecraft - NASA human factors personnel have found themselves in the unique position to redefine the human as a system from the very foundation of design. They seek to develop and monitor a "not to fall below" requirement for crew net habitable volume (NHV) - balanced against the "not to exceed" system volume requirements, with the spacecraft fitting the crew versus the crew having to fit inside the spacecraft.

Star field attitude sensor study for the Pioneer Venus spacecraft

NASA Technical Reports Server (NTRS)

Rudolf, W. P.; Reed, D. R.

1972-01-01

The characteristics of a star field attitude sensor for use with the Pioneer Venus spacecraft are presented. The aspects of technical feasibility, system interface considerations, and cost of flight hardware development are discussed. The tradeoffs which relate to performance, design, cost, and reliability are analyzed. The configuration of the system for installation in the spacecraft is described.

Spacecraft detumbling through energy dissipation

NASA Technical Reports Server (NTRS)

Fitz-Coy, Norman; Chatterjee, Anindya

1993-01-01

The attitude motion of a tumbling, rigid, axisymmetric spacecraft is considered. A methodology for detumbling the spacecraft through energy dissipation is presented. The differential equations governing this motion are stiff, and therefore an approximate solution, based on the variation of constants method, is developed and utilized in the analysis of the detumbling strategy. Stability of the detumbling process is also addressed.

Program For Simulation Of Trajectories And Events

NASA Technical Reports Server (NTRS)

Gottlieb, Robert G.

1992-01-01

Universal Simulation Executive (USE) program accelerates and eases generation of application programs for numerical simulation of continuous trajectories interrupted by or containing discrete events. Developed for simulation of multiple spacecraft trajectories with events as one spacecraft crossing the equator, two spacecraft meeting or parting, or firing rocket engine. USE also simulates operation of chemical batch processing factory. Written in Ada.

Development of a PPT for the EO-1 Spacecraft

NASA Technical Reports Server (NTRS)

Benson, Scott W.; Arrington, Lynn A.; Hoskins, W. Andrew; Meckel, Nicole J.

2000-01-01

A Pulsed Plasma Thruster (PPT) has been developed for use in a technology demonstration flight experiment on the Earth Observing 1 (EO-1) New Millennium Program mission. The thruster replaces the spacecraft pitch axis momentum wheel for control and momentum management during an experiment of a minimum three-day duration. The EO-1 PPT configuration is a combination of new technology and design heritage from similar systems flown in the 1970's and 1980's. Acceptance testing of the protoflight unit has validated readiness for flight, and integration with the spacecraft, including initial combined testing, has been completed. The thruster provides a range of capability from 90 microN-sec impulse bit at 650 sec specific impulse for 12 W input power, through 860 microN-sec impulse bit at 1400 see specific impulse for 70 W input power. Development of this thruster reinitiates technology research and development and re-establishes an industry base for production of flight hardware. This paper reviews the EO-1 PPT development, including technology selection, design and fabrication, acceptance testing, and initial spacecraft integration and test.

In-space technology development: Atomic oxygen and orbital debris effects

NASA Technical Reports Server (NTRS)

Visentine, James T.; Potter, Andrew E., Jr.

1989-01-01

Earlier Shuttle flight experiments have shown atomic oxygen within the orbital environment can interact with many materials to produce surface recession and mass loss and combine catalytically with other constituents to generate visible and infrared glows. In addition to these effects, examinations of returned satellite hardware have shown many spacecraft materials are also susceptible to damage from high velocity impacts with orbital space debris. These effects are of particular concern for large, multi-mission spacecraft, such as Space Station and SDI operational satellites, that will operate in low-Earth orbit (LEO) during the late 1990's. Not only must these spacecraft include materials and exterior coatings that are resistant to atomic oxygen surface interactions, but these materials must also provide adequate protection against erosion and pitting that could result from numerous impacts with small particles (less than 100 microns) of orbital space debris. An overview of these concerns is presented, and activities now underway to develop materials and coatings are outlined that will provide adequate atomic protection for future spacecraft. The report also discusses atomic oxygen and orbital debris flight experiments now under development to expand our limited data base, correlate ground-based measurments with flight results, and develop an orbital debris collision warning system for use by future spacecraft.

Bounding the Spacecraft Atmosphere Design Space for Future Exploration Missions

NASA Technical Reports Server (NTRS)

Lange, Kevin E.; Perka, Alan T.; Duffield, Bruce E.; Jeng, Frank F.

2005-01-01

The selection of spacecraft and space suit atmospheres for future human space exploration missions will play an important, if not critical, role in the ultimate safety, productivity, and cost of such missions. Internal atmosphere pressure and composition (particularly oxygen concentration) influence many aspects of spacecraft and space suit design, operation, and technology development. Optimal atmosphere solutions must be determined by iterative process involving research, design, development, testing, and systems analysis. A necessary first step in this process is the establishment of working bounds on the atmosphere design space.

Development of a Spacecraft Materials Selector Expert System

NASA Technical Reports Server (NTRS)

Pippin, G.; Kauffman, W. (Technical Monitor)

2002-01-01

This report contains a description of the knowledge base tool and examples of its use. A downloadable version of the Spacecraft Materials Selector (SMS) knowledge base is available through the NASA Space Environments and Effects Program. The "Spacecraft Materials Selector" knowledge base is part of an electronic expert system. The expert system consists of an inference engine that contains the "decision-making" code and the knowledge base that contains the selected body of information. The inference engine is a software package previously developed at Boeing, called the Boeing Expert System Tool (BEST) kit.

Telerobotic on-orbit remote fluid resupply system

NASA Technical Reports Server (NTRS)

1990-01-01

The development of a telerobotic on-orbit fluid resupply demonstration system is described. A fluid transfer demonstration system was developed which functionally simulates operations required to remotely transfer fluids (liquids or gases) from a servicing spacecraft to a receiving spacecraft through the use of telerobotic manipulations. The fluid system is representative of systems used by current or planned spacecraft and propulsion stages requiring on-orbit remote resupply. The system was integrated with an existing MSFC remotely controlled manipulator arm to mate/demate couplings for demonstration and evaluation of a complete remotely operated fluid transfer system.

Conceptual design of a two stage to orbit spacecraft

NASA Technical Reports Server (NTRS)

Armiger, Scott C.; Kwarta, Jennifer S.; Horsley, Kevin B.; Snow, Glenn A.; Koe, Eric C.; Single, Thomas G.

1993-01-01

This project, undertaken through the Advanced Space Design Program, developed a 'Conceptual Design of a Two Stage To Orbit Spacecraft (TSTO).' The design developed utilizes a combination of air breathing and rocket propulsion systems and is fully reusable, with horizontal takeoff and landing capability. The orbiter is carried in an aerodynamically designed bay in the aft section of the booster vehicle to the staging altitude. This TSTO Spacecraft design meets the requirements of replacing the aging Space Shuttle system with a more easily maintained vehicle with more flexible mission capability.

Radiation shielding estimates for manned Mars space flight.

PubMed

Dudkin, V E; Kovalev, E E; Kolomensky, A V; Sakovich, V A; Semenov, V F; Demin, V P; Benton, E V

1992-01-01

In the analysis of the required radiation shielding protection of spacecraft during a Mars flight, specific effects of solar activity (SA) on the intensity of galactic and solar cosmic rays were taken into consideration. Three spaceflight periods were considered: (1) maximum SA; (2) minimum SA; and (3) intermediate SA, when intensities of both galactic and solar cosmic rays are moderately high. Scenarios of spaceflights utilizing liquid-propellant rocket engines, low- and intermediate-thrust nuclear electrojet engines, and nuclear rocket engines, all of which have been designed in the Soviet Union, are reviewed. Calculations were performed on the basis of a set of standards for radiation protection approved by the U.S.S.R. State Committee for Standards. It was found that the lowest estimated mass of a Mars spacecraft, including the radiation shielding mass, obtained using a combination of a liquid propellant engine with low and intermediate thrust nuclear electrojet engines, would be 500-550 metric tons.

Entry trajectory, entry environment, and analysis of spacecraft motion for the RAM C-3 flight experiment

NASA Technical Reports Server (NTRS)

Weaver, W. L.; Bowen, J. T.

1972-01-01

The RAM C-3 flight experiment was launched to study the problem of radiofrequency blackout at an entry velocity of 24,300 ft/sec. The flight is described, and data for the entry trajectory and environment, which include the effects of actual temperature measured the day of launch, are presented. An analysis of entry spacecraft motions was performed. This analysis included the determination of wind angles from measured accelerations and estimates of wind angles at high altitudes from gyro-measured rotation rates. The maximum wind angles were found to be less than 5 deg to the point of pitch-roll resonance where the total wind angle increased to 8.5 deg and the roll rate started decreasing. A plausible cause for the decrease in roll rate was shown to be a combination of trim angle and an offset center of gravity.

Variations in plasma wave intensity with distance along the electron foreshock boundary at Venus

NASA Technical Reports Server (NTRS)

Crawford, G. K.; Strangeway, R. J.; Russell, C. T.

1991-01-01

Plasma waves are observed in the solar wind upstream of the Venus bow shock by the Pioneer Venus Orbiter. These wave signatures occur during periods when the interplanetary magnetic field through the spacecraft position intersects the bow shock, thereby placing the spacecraft in the foreshock region. Wave intensity is analyzed as a function of distance along the electron foreshock boundary. It is found that the peak wave intensity may increase along the foreshock boundary from the tangent point to a maximum value at several Venus radii, then decrease in intensity with subsequent increase in distance. These observations could be associated with the instability process: the instability of the distribution function increasing with distance from the tangent point to saturation at the peak. Thermalization of the beam for distances beyond this point could reduce the distribution function instability resulting in weaker wave signatures.

Fire-related medical science

NASA Technical Reports Server (NTRS)

Knight, Douglas R.

1987-01-01

Spacecraft fire safety may be improved by the use of a fire-retardant atmosphere in occupied spaces. Low concentrations of oxygen can protect humans from fire damage by reducing the rate and spread of combustion, but care must be taken to avoid the hypoxic effects of oxygen-lean atmospheres. Crews can live and work in 11 percent oxygen if barometric pressure were adjusted to maintain the partial pressure of oxygen above 16 kPa. Eleven percent oxygen should prevent most types of fires, since 15 percent oxygen retards the combustion of paper and 13 percent oxygen extinguishes pentane flames. Test results indicate that seated humans can perform mental tasks in atmospheres containing 11.5 percent oxygen. Although this strategy of fire safety is under consideration for submarines, it could be adapted to spacecraft once operational procedures define a maximum hyperbaric pressure and fire research defines the effects of reduced oxygen concentrations on combustion in low gravity environments.

SMM coronagraph observations of particulate contamination

NASA Technical Reports Server (NTRS)

St. Cyr, O. C.; Warner, T.

1991-01-01

Some recent images taken by the white light coronagraph telescope aboard the Solar Maximum Mission (SMM) observatory show bright streaks that are apparently caused by particles associated with the spacecraft. In this report we describe these observations, and we analyze the times of their occurrence. We demonstrate that the sightings occur most often near SMM's orbital dawn, and we speculate that thermal shock is the mechanism that produces the particles. Although these sightings have not seriously affected the coronagraph's scientific operations, the unexpected passage of bright material through the field of view of sensitive spaceborne telescopes can lead to data loss or, in some cases, serious detector damage. The topic of space debris has become a significant concern for designers of both manned and unmanned orbiting platforms. The returned samples from the SMM spacecraft and the observations reported here provide a baseline of experience for future orbital platforms that plan long duration missions.

NASA's Space Environments and Effects (SEE) Program: The Pursuit of Tomorrow's Space Technology

NASA Technical Reports Server (NTRS)

Pearson, Steven D.; Hardage, Donna M.

1998-01-01

A hazard to all spacecraft orbiting the earth and exploring the unknown in deep space is the existence of a harsh and ever changing environment with its subsequent effects. Some of these environmental hazards, such as plasma, extreme thermal excursions, meteoroids, and ionizing radiation result from natural sources, whereas others, such as orbital debris and neutral contamination are induced by the presence of spacecraft themselves. The subsequent effects can provide damaging or even disabling effects on spacecraft, its materials, and its instruments. In partnership with industry, academia, and other government agencies, National Aeronautics & Space Administration's (NASA's) Space Environments & Effects (SEE) Program defines the space environments and advocates technology development to accommodate or mitigate these harmful environments on the spacecraft. This program provides a very comprehensive and focused approach to understanding the space environment, to define the best techniques for both flight and ground-based experimentation, to update the models which predict both the environments and the environmental effects on spacecraft, and finally to ensure that this information is properly maintained and inserted into spacecraft design programs. This paper will provide an overview of the Program's purpose, goals, database management and technical activities. In particular, the SEE Program has been very active in developing improved ionizing radiation models and developing related flight experiments which should aid in determining the effect of the radiation environment on modern electronics.

Computer system evolution requirements for autonomous checkout of exploration vehicles

NASA Technical Reports Server (NTRS)

Davis, Tom; Sklar, Mike

1991-01-01

This study, now in its third year, has had the overall objective and challenge of determining the needed hooks and scars in the initial Space Station Freedom (SSF) system to assure that on-orbit assembly and refurbishment of lunar and Mars spacecraft can be accomplished with the maximum use of automation. In this study automation is all encompassing and includes physical tasks such as parts mating, tool operation, and human visual inspection, as well as non-physical tasks such as monitoring and diagnosis, planning and scheduling, and autonomous visual inspection. Potential tasks for automation include both extravehicular activity (EVA) and intravehicular activity (IVA) events. A number of specific techniques and tools have been developed to determine the ideal tasks to be automated, and the resulting timelines, changes in labor requirements and resources required. The Mars/Phobos exploratory mission developed in FY89, and the Lunar Assembly/Refurbishment mission developed in FY90 and depicted in the 90 Day Study as Option 5, have been analyzed in detailed in recent years. The complete methodology and results are presented in FY89 and FY90 final reports.

Scientific and technical services in the development of planetary quarantine measures for automated spacecraft

NASA Technical Reports Server (NTRS)

1973-01-01

Primary goals of the Planetary Quarantine Program are defined and used to provide a basis for planning and source allocation toward the development of planetary quarantine measures for the following automated spacecrafts: Viking 1975, Pioneer F and G, and Mariner Venus-Mercury 1973.

Performance Evaluation of the Gravity Probe B Design

NASA Technical Reports Server (NTRS)

Francis, Ronnie; Wells, Eugene M.

1996-01-01

This report documents the simulation of the Lockheed Martin designed Gravity Probe B (GPB) spacecraft developed tool by bd Systems Inc using the TREETOPS simulation. This study quantifies the effects of flexibility and liquid helium slosh on GPB spacecraft control performance. The TREETOPS simulation tool permits the simulation of flexible structures given that a flexible body model of the structure is available. For purposes of this study, a flexible model of the GPB spacecraft was obtained from Lockheed Martin. To model the liquid helium slosh effects, computational fluid dynamics (CFD) results' were obtained, and used to develop a dynamic model of the slosh effects. The flexible body and slosh effects were incorporated separately into the TREETOPS simulation, which places the vehicle in a 650 km circular polar orbit and subjects the spacecraft to realistic environmental disturbances and sensor error quantities. In all of the analysis conducted in this study the spacecraft is pointed at an inertially fixed guide star (GS) and is rotating at a constant rate about this line of sight.

Low-Cost Innovation in Spaceflight: The Near Earth Asteroid Rendezvous (NEAR) Shoemaker Mission

NASA Technical Reports Server (NTRS)

McCurdy, Howard E.

2005-01-01

On a spring day in 1996, at their research center in the Maryland countryside, representatives from the Johns Hopkins University Applied Physics Laboratory (APL) presented Administrator Daniel S. Goldin of the National Aeronautics and Space Administration (NASA) with a check for $3.6 million. 1 Two and a half years earlier, APL officials had agreed to develop a spacecraft capable of conducting an asteroid rendezvous and to do so for slightly more than $122 million. This was a remarkably low sum for a spacecraft due to conduct a planetaryclass mission. By contrast, the Mars Observer spacecraft launched in 1992 for an orbital rendezvous with the red planet had cost $479 million to develop, while the upcoming Cassini mission to Saturn required a spacecraft whose total cost was approaching $1.4 billion. In an Agency accustomed to cost overruns on major missions, the promise to build a planetary-class spacecraft for about $100 million seemed excessively optimistic.

NASA Handbook for Spacecraft Structural Dynamics Testing

NASA Technical Reports Server (NTRS)

Kern, Dennis L.; Scharton, Terry D.

2005-01-01

Recent advances in the area of structural dynamics and vibrations, in both methodology and capability, have the potential to make spacecraft system testing more effective from technical, cost, schedule, and hardware safety points of view. However, application of these advanced test methods varies widely among the NASA Centers and their contractors. Identification and refinement of the best of these test methodologies and implementation approaches has been an objective of efforts by the Jet Propulsion Laboratory on behalf of the NASA Office of the Chief Engineer. But to develop the most appropriate overall test program for a flight project from the selection of advanced methodologies, as well as conventional test methods, spacecraft project managers and their technical staffs will need overall guidance and technical rationale. Thus, the Chief Engineer's Office has recently tasked JPL to prepare a NASA Handbook for Spacecraft Structural Dynamics Testing. An outline of the proposed handbook, with a synopsis of each section, has been developed and is presented herein. Comments on the proposed handbook are solicited from the spacecraft structural dynamics testing community.

NASA Handbook for Spacecraft Structural Dynamics Testing

NASA Technical Reports Server (NTRS)

Kern, Dennis L.; Scharton, Terry D.

2004-01-01

Recent advances in the area of structural dynamics and vibrations, in both methodology and capability, have the potential to make spacecraft system testing more effective from technical, cost, schedule, and hardware safety points of view. However, application of these advanced test methods varies widely among the NASA Centers and their contractors. Identification and refinement of the best of these test methodologies and implementation approaches has been an objective of efforts by the Jet Propulsion Laboratory on behalf of the NASA Office of the Chief Engineer. But to develop the most appropriate overall test program for a flight project from the selection of advanced methodologies, as well as conventional test methods, spacecraft project managers and their technical staffs will need overall guidance and technical rationale. Thus, the Chief Engineer's Office has recently tasked JPL to prepare a NASA Handbook for Spacecraft Structural Dynamics Testing. An outline of the proposed handbook, with a synopsis of each section, has been developed and is presented herein. Comments on the proposed handbook is solicited from the spacecraft structural dynamics testing community.

AIAA spacecraft GN&C interface standards initiative: Overview

NASA Technical Reports Server (NTRS)

Challoner, A. Dorian

1995-01-01

The American Institute of Aeronautics and Astronautics (AIAA) has undertaken an important standards initiative in the area of spacecraft guidance, navigation, and control (GN&C) subsystem interfaces. The objective of this effort is to establish standards that will promote interchangeability of major GN&C components, thus enabling substantially lower spacecraft development costs. Although initiated by developers of conventional spacecraft GN&C, it is anticipated that interface standards will also be of value in reducing the development costs of micro-engineered spacecraft. The standardization targets are specifically limited to interfaces only, including information (i.e. data and signal), power, mechanical, thermal, and environmental interfaces between various GN&C components and between GN&C subsystems and other subsystems. The current emphasis is on information interfaces between various hardware elements (e.g., between star trackers and flight computers). The poster presentation will briefly describe the program, including the mechanics and schedule, and will publicize the technical products as they exist at the time of the conference. In particular, the rationale for the adoption of the AS1773 fiber-optic serial data bus and the status of data interface standards at the application layer will be presented.

Operational Analysis of Time-Optimal Maneuvering for Imaging Spacecraft

DTIC Science & Technology

2013-03-01

imaging spacecraft. The analysis is facilitated through the use of AGI’s Systems Tool Kit ( STK ) software. An Analytic Hierarchy Process (AHP)-based...the Singapore-developed X-SAT imaging spacecraft. The analysis is facilitated through the use of AGI’s Systems Tool Kit ( STK ) software. An Analytic...89  B.  FUTURE WORK................................................................................. 90  APPENDIX A. STK DATA AND BENEFIT

Evaluation of spacecraft technology programs (effects on communication satellite business ventures), volume 1

NASA Technical Reports Server (NTRS)

Greenburg, J. S.; Gaelick, C.; Kaplan, M.; Fishman, J.; Hopkins, C.

1985-01-01

Commercial organizations as well as government agencies invest in spacecraft (S/C) technology programs that are aimed at increasing the performance of communications satellites. The value of these programs must be measured in terms of their impacts on the financial performane of the business ventures that may ultimately utilize the communications satellites. An economic evaluation and planning capability was developed and used to assess the impact of NASA on-orbit propulsion and space power programs on typical fixed satellite service (FSS) and direct broadcast service (DBS) communications satellite business ventures. Typical FSS and DBS spin and three-axis stabilized spacecraft were configured in the absence of NASA technology programs. These spacecraft were reconfigured taking into account the anticipated results of NASA specified on-orbit propulsion and space power programs. In general, the NASA technology programs resulted in spacecraft with increased capability. The developed methodology for assessing the value of spacecraft technology programs in terms of their impact on the financial performance of communication satellite business ventures is described. Results of the assessment of NASA specified on-orbit propulsion and space power technology programs are presented for typical FSS and DBS business ventures.

Development of a 3D numerical code to calculate the trajectories of the blow off electrons emitted by a vacuum surface discharge: Application to the study of the electromagnetic interference induced on a spacecraft

NASA Astrophysics Data System (ADS)

Froger, Etienne

1993-05-01

A description of the electromagnetic behavior of a satellite subjected to an electric discharge is given using a specially developed numerical code. One of the particularities of vacuum discharges, obtained by irradiation of polymers, is the intense emission of electrons into the spacecraft environment. Electromagnetic radiation, associated with the trajectories of the particles around the spacecraft, is considered as the main source of the interference observed. In the absence of accurate orbital data and realistic ground tests, the assessment of these effects requires numerical simulation of the interaction between this electron source and the spacecraft. This is done by the GEODE particle code which is applied to characteristic configurations in order to estimate the spacecraft response to a discharge, which is simulated from a vacuum discharge model designed in laboratory. The spacecraft response to a current injection is simulated by the ALICE numerical three dimensional code. The comparison between discharge and injection effects, from the results given by the two codes, illustrates the representativity of electromagnetic susceptibility tests and the main parameters for their definition.

Evaluation of spacecraft technology programs (effects on communication satellite business ventures), volume 1

NASA Astrophysics Data System (ADS)

Greenburg, J. S.; Gaelick, C.; Kaplan, M.; Fishman, J.; Hopkins, C.

1985-09-01

Commercial organizations as well as government agencies invest in spacecraft (S/C) technology programs that are aimed at increasing the performance of communications satellites. The value of these programs must be measured in terms of their impacts on the financial performane of the business ventures that may ultimately utilize the communications satellites. An economic evaluation and planning capability was developed and used to assess the impact of NASA on-orbit propulsion and space power programs on typical fixed satellite service (FSS) and direct broadcast service (DBS) communications satellite business ventures. Typical FSS and DBS spin and three-axis stabilized spacecraft were configured in the absence of NASA technology programs. These spacecraft were reconfigured taking into account the anticipated results of NASA specified on-orbit propulsion and space power programs. In general, the NASA technology programs resulted in spacecraft with increased capability. The developed methodology for assessing the value of spacecraft technology programs in terms of their impact on the financial performance of communication satellite business ventures is described. Results of the assessment of NASA specified on-orbit propulsion and space power technology programs are presented for typical FSS and DBS business ventures.

A shuttle and space station manipulator system for assembly, docking, maintenance, cargo handling and spacecraft retrieval (preliminary design). Volume 4: Simulation studies

NASA Technical Reports Server (NTRS)

1972-01-01

Laboratory simulations of three concepts, based on maximum use of available off-the-shelf hardware elements, are described. The concepts are a stereo-foveal-peripheral TV system with symmetric steroscopic split-image registration and 90 deg counter rotation; a computer assisted model control system termed the trajectory following control system; and active manipulator damping. It is concluded that the feasibility of these concepts is established.

Optimum reentry trajectories of a lifting vehicle

NASA Technical Reports Server (NTRS)

Chern, J. S.; Vinh, N. X.

1980-01-01

Research results are presented of an investigation of the optimum maneuvers of advanced shuttle type spacecraft during reentry. The equations are formulated by means of modified Chapman variables resulting in a general set of equations for flight analysis which are exact for reentry and for flight in a vacuum. Four planar flight typical optimum manuevers are investigated. For three-dimensional flight the optimum trajectory for maximum cross range is discussed in detail. Techniques for calculating reentry footprints are presented.

Proximity Operations for Space Situational Awareness Spacecraft Rendezvous and Maneuvering using Numerical Simulations and Fuzzy Logic

NASA Astrophysics Data System (ADS)

Carrico, T.; Langster, T.; Carrico, J.; Alfano, S.; Loucks, M.; Vallado, D.

The authors present several spacecraft rendezvous and close proximity maneuvering techniques modeled with a high-precision numerical integrator using full force models and closed loop control with a Fuzzy Logic intelligent controller to command the engines. The authors document and compare the maneuvers, fuel use, and other parameters. This paper presents an innovative application of an existing capability to design, simulate and analyze proximity maneuvers; already in use for operational satellites performing other maneuvers. The system has been extended to demonstrate the capability to develop closed loop control laws to maneuver spacecraft in close proximity to another, including stand-off, docking, lunar landing and other operations applicable to space situational awareness, space based surveillance, and operational satellite modeling. The fully integrated end-to-end trajectory ephemerides are available from the authors in electronic ASCII text by request. The benefits of this system include: A realistic physics-based simulation for the development and validation of control laws A collaborative engineering environment for the design, development and tuning of spacecraft law parameters, sizing actuators (i.e., rocket engines), and sensor suite selection. An accurate simulation and visualization to communicate the complexity, criticality, and risk of spacecraft operations. A precise mathematical environment for research and development of future spacecraft maneuvering engineering tasks, operational planning and forensic analysis. A closed loop, knowledge-based control example for proximity operations. This proximity operations modeling and simulation environment will provide a valuable adjunct to programs in military space control, space situational awareness and civil space exploration engineering and decision making processes.

Development of hydrogen peroxide technique for bioburden reduction

NASA Astrophysics Data System (ADS)

Rohatgi, N.; Schwartz, L.; Stabekis, P.; Barengoltz, J.

In order to meet the National Aeronautics and Space Administration (NASA) Planetary Protection microbial reduction requirements for Mars in-situ life detection and sample return missions, entire planetary spacecraft (including planetary entry probes and planetary landing capsules) may have to be exposed to a qualified sterilization process. Presently, dry heat is the only NASA approved sterilization technique available for spacecraft application. However, with the increasing use of various man-made materials, highly sophisticated electronic circuit boards, and sensors in a modern spacecraft, compatibility issues may render this process unacceptable to design engineers and thus impractical to achieve terminal sterilization of the entire spacecraft. An alternative vapor phase hydrogen peroxide sterilization process, which is currently used in various industries, has been selected for further development. Strategic Technology Enterprises, Incorporated (STE), a subsidiary of STERIS Corporation, under a contract from the Jet Propulsion Laboratory (JPL) is developing systems and methodologies to decontaminate spacecraft using vaporized hydrogen peroxide (VHP) technology. The VHP technology provides an effective, rapid and low temperature means for inactivation of spores, mycobacteria, fungi, viruses and other microorganisms. The VHP application is a dry process affording excellent material compatibility with many of the components found in spacecraft such as polymers, paints and electronic systems. Furthermore, the VHP process has innocuous residuals as it decomposes to water vapor and oxygen. This paper will discuss the approach that is being used to develop this technique and will present lethality data that have been collected to establish deep vacuum VHP sterilization cycles. In addition, the application of this technique to meet planetary protection requirements will be addressed.

Landing Characteristics of the Apollo Spacecraft with Deployed Heat Shield Impact Attenuation Systems

NASA Technical Reports Server (NTRS)

Stubbs, Sandy M.

1965-01-01

An experimental investigation was made to determine the landing characteristics of a 1/4-scale dynamic model of the Apollo spacecraft command module using two different active (heat shield deployed prior to landing) landing systems for impact attenuation. One landing system (configuration 1) consisted of six hydraulic struts and eight crushable honeycomb struts. The other landing system (configuration 2), consisted of four hydraulic struts and six strain straps. Tests made on water and the hard clay-gravel composite landing surfaces simulated parachute letdown (vertical) velocities of 23 ft/sec (7.0 m/s) (full scale). Landings made on the sand landing surface simulated vertical velocities of 30 ft/sec (9.1 m/s). Horizontal velocities of from 0 to 50 ft/sec (15 m/s) were simulated. Landing attitudes ranged from -30'degrees to 20 degrees, and the roll attitudes were O degrees, 90 degrees, and 180 degrees. For configuration 1, maximum normal accelerations at the vehicle center of gravity for landings on water, sand, and the hard clay-gravel composite surface were 9g, 20g, and 18g, respectively. The maximum normal center-of-gravity acceleration for configuration 2 which was landed only on the hard clay-gravel landing surface was approximately 19g. Accelerations for configuration 2 were generally equal to or lower than accelerations for configuration 1 and normal.

The location and rate of occurrence of near-Earth magnetotail reconnection as observed by Cluster and Geotail

NASA Astrophysics Data System (ADS)

Genestreti, K. J.; Fuselier, S. A.; Goldstein, J.; Nagai, T.; Eastwood, J. P.

2014-12-01

A statistical characterization of the location and rate of occurrence of magnetic reconnection in the near-Earth magnetotail is performed by analyzing the set of ion diffusion region (DR) observations made by the Cluster and Geotail spacecraft during solar maximum and the declining phase. The occurrence rate is analyzed in terms of its dependence on both XGSM* and YGSM* (where coordinates are in the solar wind aberrated geocentric solar magnetospheric system). Within the limits of the statistics available to this study, we find the purely XGSM* -dependent occurrence rate to be roughly constant over a large portion of the near-Earth magnetotail. In contrast, we find the purely YGSM* -dependent occurrence rate to be biased towards dusk with a local maximum between 0RE ≤YGSM* ≤ 5RE. The YGSM* -dependent occurrence rate is then used to construct a quasi-2D formulation of the DR occurrence rate, which has explicit dependence on XGSM* and implicit dependence on YGSM*. The quasi-2D occurrence rate is then used to examine the predicted ephemeris of the Magnetospheric MultiScale (MMS) spacecraft. We estimate that, during its near-Earth magnetotail survey phase, MMS will likely observe 11±4 DR events. • The occurrence rate of events is calculated as a function of XGSM and YGSM. • The occurrence rate is used to estimate the number of events MMS will observe.

Taurus lightweight manned spacecraft Earth orbiting vehicle

NASA Technical Reports Server (NTRS)

Chase, Kevin A.; Vandersall, Eric J.; Plotkin, Jennifer; Travisano, Jeffrey J.; Loveless, Dennis; Kaczmarek, Michael; White, Anthony G.; Est, Andy; Bulla, Gregory; Henry, Chris

1991-01-01

The Taurus Lightweight Manned Spacecraft (LMS) was developed by students of the University of Maryland's Aerospace Engineering course in Space Vehicle Design. That course required students to design an Alternative Manned Spacecraft (AMS) to augment or replace the Space Transportation System and meet the following design requirements: (1) launch on the Taurus Booster being developed by Orbital Sciences Corporation; (2) 99.9 percent assured crew survival rate; (3) technology cutoff data of 1 Jan. 1991; (4) compatibility with current space administration infrastructure; and (5) first flight by May 1995. The Taurus LMS design meets the above requirements and represents an initial step towards larger and more complex spacecraft. The Taurus LMS has a very limited application when compared to the Space Shuttle, but it demonstrates that the U.S. can have a safe, reliable, and low cost space system. The Taurus LMS is a short mission duration spacecraft designed to place one man into low earth orbit (LEO). The driving factor for this design was the low payload carrying capabilities of the Taurus Booster--1300 kg to a 300 km orbit. The Taurus LMS design is divided into six major design sections. The human factors system deals with the problems of life support and spacecraft cooling. The propulsion section contains the abort system, the Orbital Maneuvering System (OMS), the Reaction Control System (RCS), and power generation. The thermal protection systems and spacecraft structure are contained in the structures section. The avionics section includes navigation, attitude determination, data processing, communication systems, and sensors. The mission analysis section was responsible for ground processing and spacecraft astrodynamics. The systems integration section pulled the above sections together into one spacecraft and addressed costing and reliability.

Taurus Lightweight Manned Spacecraft Earth orbiting vehicle

NASA Technical Reports Server (NTRS)

Bosset, M.

1991-01-01

The Taurus Lightweight Manned Spacecraft (LMS) was developed by students of the University of Maryland's Aerospace Engineering course in Space Vehicle Design. That course required students to design an Alternative Manned Spacecraft (AMS) to augment or replace the Space Transportation System and meet the following design requirements: (1) launch on the Taurus Booster being developed by Orbital Sciences Corporation; (2) 99.9 percent assured crew survival rate; (3) technology cutoff date of 1 Jan. 1991; (4) compatibility with current space administration infrastructure; and (5) first flight by May 1995. The Taurus LMS design meets the above requirements and represents an initial step toward larger and more complex spacecraft. The Taurus LMS has a very limited application when compared to the space shuttle, but it demonstrates that the U.S. can have a safe, reliable, and low-cost space system. The Taurus LMS is a short mission duration spacecraft designed to place one man into low Earth orbit (LEO). The driving factor for this design was the low payload carrying capabilities of the Taurus Booster - 1300 kg to a 300-km orbit. The Taurus LMS design is divided into six major design sections. The Human Factors section deals with the problems of life support and spacecraft cooling. The Propulsion section contains the Abort System, the Orbital Maneuvering System (OMS), the Reaction Control System (RCS), and Power Generation. The thermal protection systems and spacecraft structure are contained in the Structures section. The Avionics section includes Navigation, Attitude Determination, Data Processing, Communication systems, and Sensors. The Mission Analysis section was responsible for ground processing and spacecraft astrodynamics. The Systems Integration Section pulled the above sections together into one spacecraft, and addressed costing and reliability.

Cryogenic On-Orbit Liquid Depot-Storage, Acquisition and Transfer (COLD-SAT) Experiment Conceptual Design and Feasibility Study

NASA Technical Reports Server (NTRS)

Kramer, Edward (Editor)

1998-01-01

The cryogenic fluid management technologies required for the exploration of the solar system can only be fully developed via space-based experiments. A dedicated spacecraft is the most efficient way to perform these experiments. This report documents the extended conceptual design of the COLD-SAT spacecraft, capable of meeting these experimental requirements. All elements, including the spacecraft, ground segment, launch site modifications and launch vehicle operations, and flight operations are included. Greatly expanded coverage is provided for those areas unique to this cryogenic spacecraft, such as the experiment system, attitude control system, and spacecraft operations. Supporting analyses are included as are testing requirements, facilities surveys, and proposed project timelines.

Vehicle/Atmosphere Interaction Glows: Far Ultraviolet, Visible, and Infrared

NASA Technical Reports Server (NTRS)

Swenson, G.

1999-01-01

Spacecraft glow information has been gathered from a number of spacecraft including Atmospheric and Dynamic satellites, and Space Shuttles (numerous flights) with dedicated pallet flow observations on STS-39 (DOD) and STS-62 (NASA). In addition, a larger number of laboratory experiments with low energy oxygen beam studies have made important contributions to glow understanding. The following report provides information on three engineering models developed for spacecraft glow including the far ultraviolet to ultraviolet (1400-4000 A), and infrared (0.9-40 microns) spectral regions. The models include effects resulting from atmospheric density/altitude, spacecraft temperature, spacecraft material, and ram angle. Glow brightness would be predicted as a function of distance from surfaces for all wavelengths.

GSMS and space views: Advanced spacecraft monitoring tools

NASA Technical Reports Server (NTRS)

Carlton, Douglas; Vaules, David, Jr.; Mandl, Daniel

1993-01-01

The Graphical Spacecraft Monitoring System (GSMS) processes and translates real-time telemetry data from the Gamma Ray Observatory (GRO) spacecraft into high resolution 2-D and 3-D color displays showing the spacecraft's position relative to the Sun, Earth, Moon, and stars, its predicted orbit path, its attitude, instrument field of views, and other items of interest to the GRO Flight Operations Team (FOT). The GSMS development project is described and the approach being undertaken for implementing Space Views, the next version of GSMS, is presented. Space Views is an object-oriented graphical spacecraft monitoring system that will become a standard component of Goddard Space Flight Center's Transportable Payload Operations Control Center (TPOCC).

Autonomic Computing for Spacecraft Ground Systems

NASA Technical Reports Server (NTRS)

Li, Zhenping; Savkli, Cetin; Jones, Lori

2007-01-01

Autonomic computing for spacecraft ground systems increases the system reliability and reduces the cost of spacecraft operations and software maintenance. In this paper, we present an autonomic computing solution for spacecraft ground systems at NASA Goddard Space Flight Center (GSFC), which consists of an open standard for a message oriented architecture referred to as the GMSEC architecture (Goddard Mission Services Evolution Center), and an autonomic computing tool, the Criteria Action Table (CAT). This solution has been used in many upgraded ground systems for NASA 's missions, and provides a framework for developing solutions with higher autonomic maturity.

Spacecraft configuration study for second generation mobile satellite system

NASA Technical Reports Server (NTRS)

Louie, M.; Vonstentzsch, W.; Zanella, F.; Hayes, R.; Mcgovern, F.; Tyner, R.

1985-01-01

A high power, high performance communicatons satellite bus being developed is designed to satisfy a broad range of multimission payload requirements in a cost effective manner and is compatible with both STS and expendable launchers. Results are presented of tradeoff studies conducted to optimize the second generation mobile satellite system for its mass, power, and physical size. Investigations of the 20-meter antenna configuration, transponder linearization techniques, needed spacecraft modifications, and spacecraft power, dissipation, mass, and physical size indicate that the advanced spacecraft bus is capable of supporting the required payload for the satellite.

Modal analysis for Liapunov stability of rotating elastic bodies. Ph.D. Thesis. Final Report

NASA Technical Reports Server (NTRS)

Colin, A. D.

1973-01-01

This study consisted of four parallel efforts: (1) modal analyses of elastic continua for Liapunov stability analysis of flexible spacecraft; (2) development of general purpose simulation equations for arbitrary spacecraft; (3) evaluation of alternative mathematical models for elastic components of spacecraft; and (4) examination of the influence of vehicle flexibility on spacecraft attitude control system performance. A complete record is given of achievements under tasks (1) and (3), in the form of technical appendices, and a summary description of progress under tasks two and four.

Sensitivity Analysis and Mitigation with Applications to Ballistic and Low-thrust Trajectory Design

NASA Astrophysics Data System (ADS)

Alizadeh, Iman

The ever increasing desire to expand space mission capabilities within the limited budgets of space industries requires new approaches to the old problem of spacecraft trajectory design. For example, recent initiatives for space exploration involve developing new tools to design low-cost, fail-safe trajectories to visit several potential destinations beyond our celestial neighborhood such as Jupiter's moons, asteroids, etc. Designing and navigating spacecraft trajectories to reach these destinations safely are complex and challenging. In particular, fundamental questions of orbital stability imposed by planetary protection requirements are not easily taken into account by standard optimal control schemes. The event of temporary engine loss or an unexpected missed thrust can indeed quickly lead to impact with planetary bodies or other unrecoverable trajectories. While electric propulsion technology provides superior efficiency compared to chemical engines, the very low-control authority and engine performance degradation can impose higher risk to the mission in strongly perturbed orbital environments. The risk is due to the complex gravitational field and its associated chaotic dynamics which causes large navigation dispersions in a short time if left un-controlled. Moreover, in these situations it can be outside the low-thrust propulsion system capability to correct the spacecraft trajectory in a reasonable time frame. These concerns can lead to complete or partial mission failure or even an infeasible mission concept at the early design stage. The goal of this research is to assess and increase orbital stability of ballistic and low-thrust transfer trajectories in multi-body systems. In particular, novel techniques are presented to characterize sensitivity and improve recovery characteristics of ballistic and low-thrust trajectories in unstable orbital environments. The techniques developed are based on perturbation analysis around ballistic trajectories to determine analytically the maximum divergence directions and also optimal control theory with nonstandard cost functions along with inverse dynamics applied to low-thrust trajectories. Several mission scenarios are shown to demonstrate the applicability of the techniques in the Earth-Moon and the Jupiter-Europa system. In addition, the results provide fundamental insight into design, stability analysis and guidance, navigation and control of low-thrust trajectories to meet challenging mission requirements in support of NASA's vision for space exploration.

Practical Applications of Cosmic Ray Science: Spacecraft, Aircraft, Ground-Based Computation and Control Systems, and Human Health and Safety

NASA Technical Reports Server (NTRS)

Atwell, William; Koontz, Steve; Normand, Eugene

2012-01-01

Three twentieth century technological developments, 1) high altitude commercial and military aircraft; 2) manned and unmanned spacecraft; and 3) increasingly complex and sensitive solid state micro-electronics systems, have driven an ongoing evolution of basic cosmic ray science into a set of practical engineering tools needed to design, test, and verify the safety and reliability of modern complex technological systems. The effects of primary cosmic ray particles and secondary particle showers produced by nuclear reactions with the atmosphere, can determine the design and verification processes (as well as the total dollar cost) for manned and unmanned spacecraft avionics systems. Similar considerations apply to commercial and military aircraft operating at high latitudes and altitudes near the atmospheric Pfotzer maximum. Even ground based computational and controls systems can be negatively affected by secondary particle showers at the Earth s surface, especially if the net target area of the sensitive electronic system components is large. Finally, accumulation of both primary cosmic ray and secondary cosmic ray induced particle shower radiation dose is an important health and safety consideration for commercial or military air crews operating at high altitude/latitude and is also one of the most important factors presently limiting manned space flight operations beyond low-Earth orbit (LEO). In this paper we review the discovery of cosmic ray effects on the performance and reliability of microelectronic systems as well as human health and the development of the engineering and health science tools used to evaluate and mitigate cosmic ray effects in ground-based atmospheric flight, and space flight environments. Ground test methods applied to microelectronic components and systems are used in combinations with radiation transport and reaction codes to predict the performance of microelectronic systems in their operating environments. Similar radiation transport codes are used to evaluate possible human health effects of cosmic ray exposure, however, the health effects are based on worst-case analysis and extrapolation of a very limited human exposure data base combined with some limited experimental animal data. Finally, the limitations on human space operations beyond low-Earth orbit imposed by long term exposure to galactic cosmic rays are discussed.

Performance of active vibration control technology: the ACTEX flight experiments

NASA Astrophysics Data System (ADS)

Nye, T. W.; Manning, R. A.; Qassim, K.

1999-12-01

This paper discusses the development and results of two intelligent structures space-flight experiments, each of which could affect architecture designs of future spacecraft. The first, the advanced controls technology experiment I (ACTEX I), is a variable stiffness tripod structure riding as a secondary payload on a classified spacecraft. It has been operating well past its expected life since becoming operational in 1996. Over 60 on-orbit experiments have been run on the ACTEX I flight experiment. These experiments form the basis for in-space controller design problems and for concluding lifetime/reliability data on the active control components. Transfer functions taken during the life of ACTEX I have shown consistent predictability and stability in structural behavior, including consistency with those measurements taken on the ground prior to a three year storage period and the launch event. ACTEX I can change its modal characteristics by employing its dynamic change mechanism that varies preloads in portions of its structure. Active control experiments have demonstrated maximum vibration reductions of 29 dB and 16 dB in the first two variable modes of the system, while operating over a remarkable on-orbit temperature range of -80 °C to 129 °C. The second experiment, ACTEX II, was successfully designed, ground-tested, and integrated on an experimental Department of Defense satellite prior to its loss during a launch vehicle failure in 1995. ACTEX II also had variable modal behavior by virtue of a two-axis gimbal and added challenges of structural flexibility by being a large deployable appendage. Although the loss of ACTEX II did not provide space environment experience, ground testing resulted in space qualifying the hardware and demonstrated 21 dB, 14 dB, and 8 dB reductions in amplitude of the first three primary structural modes. ACTEX II could use either active and/or passive techniques to affect vibration suppression. Both experiments trailblazed spacecraft bus smart structures by developing over 20 new technologies. As pathfinders, experience was gained in the implications of space system analyses, verification tests, and for ways to leverage this technology to meet new satellite performance requirements.

Spacecraft telecommunications system mass estimates

NASA Technical Reports Server (NTRS)

Yuen, J. H.; Sakamoto, L. L.

1988-01-01

Mass is the most important limiting parameter for present-day planetary spacecraft design, In fact, the entire design can be characterized by mass. The more efficient the design of the spacecraft, the less mass will be required. The communications system is an essential and integral part of planetary spacecraft. A study is presented of the mass attributable to the communications system for spacecraft designs used in recent missions in an attempt to help guide future design considerations and research and development efforts. The basic approach is to examine the spacecraft by subsystem and allocate a portion of each subsystem to telecommunications. Conceptually, this is to divide the spacecraft into two parts, telecommunications and nontelecommunications. In this way, it is clear what the mass attributable to the communications system is. The percentage of mass is calculated using the actual masses of the spacecraft parts, except in the case of CRAF. In that case, estimated masses are used since the spacecraft was not yet built. The results show that the portion of the spacecraft attributable to telecommunications is substantial. The mass fraction for Voyager, Galileo, and CRAF (Mariner Mark 2) is 34, 19, and 18 percent, respectively. The large reduction of telecommunications mass from Voyager to Galileo is mainly due to the use of a deployable antenna instead of the solid antenna on Voyager.

Guidance and Navigation for Rendezvous and Proximity Operations with a Non-Cooperative Spacecraft at Geosynchronous Orbit

NASA Technical Reports Server (NTRS)

Barbee, Brent William; Carpenter, J. Russell; Heatwole, Scott; Markley, F. Landis; Moreau, Michael; Naasz, Bo J.; VanEepoel, John

2010-01-01

The feasibility and benefits of various spacecraft servicing concepts are currently being assessed, and all require that the servicer spacecraft perform rendezvous, proximity, and capture operations with the target spacecraft to be serviced. Many high-value spacecraft, which would be logical targets for servicing from an economic point of view, are located in geosynchronous orbit, a regime in which autonomous rendezvous and capture operations are not commonplace. Furthermore, existing GEO spacecraft were not designed to be serviced. Most do not have cooperative relative navigation sensors or docking features, and some servicing applications, such as de-orbiting of a non-functional spacecraft, entail rendezvous and capture with a spacecraft that may be non-functional or un-controlled. Several of these challenges have been explored via the design of a notional mission in which a nonfunctional satellite in geosynchronous orbit is captured by a servicer spacecraft and boosted into super-synchronous orbit for safe disposal. A strategy for autonomous rendezvous, proximity operations, and capture is developed, and the Orbit Determination Toolbox (ODTBX) is used to perform a relative navigation simulation to assess the feasibility of performing the rendezvous using a combination of angles-only and range measurements. Additionally, a method for designing efficient orbital rendezvous sequences for multiple target spacecraft is utilized to examine the capabilities of a servicer spacecraft to service multiple targets during the course of a single mission.

Spacecraft Autonomy and Automation: A Comparative Analysis of Strategies for Cost Effective Mission Operations

NASA Technical Reports Server (NTRS)

Wright, Nathaniel, Jr.

2000-01-01

The evolution of satellite operations over the last 40 years has drastically changed. October 4, 1957 (during the cold war) the Soviet Union launched the world's first spacecraft into orbit. The Sputnik satellite orbited Earth for three months and catapulted the United States into a race for dominance in space. A year after Sputnik, President Dwight Eisenhower formed the National Space and Aeronautics Administration (NASA). With a team of scientists and engineers, NASA successfully launched Explorer 1, the first US satellite to orbit Earth. During these early years, massive amounts of ground support equipment and operators were required to successfully operate spacecraft vehicles. Today, budget reductions and technological advances have forced new approaches to spacecraft operations. These approaches require increasingly complex, on board spacecraft systems, that enable autonomous operations, resulting in more cost-effective mission operations. NASA's Goddard Space Flight Center, considered world class in satellite development and operations, has developed and operated over 200 satellites during its 40 years of existence. NASA Goddard is adopting several new millennium initiatives that lower operational costs through the spacecraft autonomy and automation. This paper examines NASA's approach to spacecraft autonomy and ground system automation through a comparative analysis of satellite missions for Hubble Space Telescope-HST, Near Earth Asteroid Rendezvous-NEAR, and Solar Heliospheric Observatory-SoHO, with emphasis on cost reduction methods, risk analysis and anomalies and strategies employed for mitigating risk.

Characterizing Secondary Debris Impact Ejecta

NASA Technical Reports Server (NTRS)

Schonberg, W. P.

1999-01-01

All spacecraft in low-Earth orbit are subject to high-speed impacts by meteoroids and orbital debris particles. These impacts can damage flight-critical systems which can in turn lead to catastrophic failure of the spacecraft. Therefore, the design of a spacecraft for an Earth-orbiting mission must take into account the possibility of such impacts and their effects on the spacecraft structure and on all of its exposed subsystem components. In addition to threatening the operation of the spacecraft itself, on-orbit impacts also generate a significant amount of ricochet particles. These high-speed particles can destroy critical external spacecraft subsystem and also increase the contamination of the orbital environment. This report presents a summary of the work performed towards the development of an empirical model that characterizes the secondary ejecta created by a high-speed impacta on a typical aerospace structural surface.

Spacecraft design project: Low Earth orbit communications satellite

NASA Technical Reports Server (NTRS)

Moroney, Dave; Lashbrook, Dave; Mckibben, Barry; Gardener, Nigel; Rivers, Thane; Nottingham, Greg; Golden, Bill; Barfield, Bill; Bruening, Joe; Wood, Dave

1991-01-01

This is the final product of the spacecraft design project completed to fulfill the academic requirements of the Spacecraft Design and Integration 2 course (AE-4871) taught at the U.S. Naval Postgraduate School. The Spacecraft Design and Integration 2 course is intended to provide students detailed design experience in selection and design of both satellite system and subsystem components, and their location and integration into a final spacecraft configuration. The design team pursued a design to support a Low Earth Orbiting (LEO) communications system (GLOBALSTAR) currently under development by the Loral Cellular Systems Corporation. Each of the 14 team members was assigned both primary and secondary duties in program management or system design. Hardware selection, spacecraft component design, analysis, and integration were accomplished within the constraints imposed by the 11 week academic schedule and the available design facilities.

The Near Earth Object Scout Spacecraft: A Low Cost Approach to in-situ Characterization of the NEO Population

NASA Technical Reports Server (NTRS)

Koontz, Steven L.; Condon, Gerald; Graham, Lee; Bevilacqua, Ricardo

2014-01-01

In this paper we describe a micro/nano satellite spacecraft and a supporting mission profile and architecture designed to enable preliminary in-situ characterization of a significant number of Near Earth Objects (NEOs) at reasonable cost. The spacecraft will be referred to as the NEO Scout. NEO Scout spacecraft are to be placed in GTO, GEO, or cis-lunar space as secondary payloads on launch vehicles headed for GTO or beyond and will begin their mission after deployment from the launcher. A distinguishing key feature of the NEO scout system is to design the mission timeline and spacecraft to rendezvous with and land on the target NEOs during close approach to the Earth-Moon system using low-thrust/high- impulse propulsion systems. Mission feasibility and preliminary design analysis are presented along with detailed trajectory calculations. The use of micro/nano satellites in low-cost interplanetary exploration is attracting increasing attention and is the subject of several annual workshops and published design studies (1-4). The NEO population consists of those asteroids and short period comets orbiting the Sun with a perihelion of 1.3 astronomical units or less (5-8). As of July 30, 2013 10065 Near-Earth objects have been discovered. The spin rate, mass, density, surface physical (especially mechanical) properties, composition, and mineralogy of the vast majority of these objects are highly uncertain and the limited available telescopic remote sensing data imply a very diverse population (5-8). In-situ measurements by robotic spacecraft are urgently needed to provide the characterization data needed to support hardware and mission design for more ambitious human and robotic NEO operations. Large numbers of NEOs move into close proximity with the Earth-Moon system every year (9). The JPL Near-Earth Object Human Space Flight Accessible Targets Study (NHATS) (10) has produced detailed mission profile and delta V requirements for various NEO missions ranging from 30 to 420 days in duration and assuming chemical propulsion. Similar studies have been reported assuming high power electric propulsion for manned NEO rendezvous missions (11). The delta V requirement breakdown and mission profile data from references 10 and 11 are used as a basis for sizing the NEO Scout spacecraft and for conducting preliminary feasibility assessments using the Tsiokolvsky rocket equation, a (worst-case) delta V requirement of 10 km/sec, and a maximum spacecraft dry mass of 20 kg. Using chemical propellant for a 10 km/sec delta V drives spacecraft wet mass well above 300 kg so that chemical propulsion is a non-starter for the proposed mission profile and spacecraft wet mass limits. In contrast, a solar electric propulsion system needs only 8 kg of Xe propellant to accelerate the spacecraft to 10 km/sec in 163 days with 0.02 N of thrust and 500 W of power from1.6 sq m of 29% efficient solar panels. In a second example, accelerating a 4 kg payload to 7 km/sec over 180 days requires about 6.7 kg of propellant and 1.2 kg of solar panels (12 kg total spacecraft wet mass).

Analysis of Static Spacecraft Floating Potential at Low Earth Orbit (LEO)

NASA Technical Reports Server (NTRS)

Herr, Joel L.; Hwang, K. S.; Wu, S. T.

1995-01-01

Spacecraft floating potential is the charge on the external surfaces of orbiting spacecraft relative to the space. Charging is caused by unequal negative and positive currents to spacecraft surfaces. The charging process continues until the accelerated particles can be collected rapidly enough to balance the currents at which point the spacecraft has reached its equilibrium or floating potential. In low inclination. Low Earth Orbit (LEO), the collection of positive ion and negative electrons. in a particular direction. are typically not equal. The level of charging required for equilibrium to be established is influenced by the characteristics of the ambient plasma environment. by the spacecraft motion, and by the geometry of the spacecraft. Using the kinetic theory, a statistical approach for studying the interaction is developed. The approach used to study the spacecraft floating potential depends on which phenomena are being applied. and on the properties of the plasma. especially the density and temperature. The results from kinetic theory derivation are applied to determine the charging level and the electric potential distribution at an infinite flat plate perpendicular to a streaming plasma using finite-difference scheme.

Spacecraft Hybrid (Mixed-Actuator) Attitude Control Experiences on NASA Science Missions

NASA Technical Reports Server (NTRS)

Dennehy, Cornelius J.

2014-01-01

There is a heightened interest within NASA for the design, development, and flight implementation of mixed-actuator hybrid attitude control systems for science spacecraft that have less than three functional reaction wheel actuators. This interest is driven by a number of recent reaction wheel failures on aging, but what could be still scientifically productive, NASA spacecraft if a successful hybrid attitude control mode can be implemented. Over the years, hybrid (mixed-actuator) control has been employed for contingency attitude control purposes on several NASA science mission spacecraft. This paper provides a historical perspective of NASA's previous engineering work on spacecraft mixed-actuator hybrid control approaches. An update of the current situation will also be provided emphasizing why NASA is now so interested in hybrid control. The results of the NASA Spacecraft Hybrid Attitude Control Workshop, held in April of 2013, will be highlighted. In particular, the lessons learned captured from that workshop will be shared in this paper. An update on the most recent experiences with hybrid control on the Kepler spacecraft will also be provided. This paper will close with some future considerations for hybrid spacecraft control.

Got risk? risk-centric perspective for spacecraft technology decision-making

NASA Technical Reports Server (NTRS)

Feather, Martin S.; Cornford, Steven L.; Moran, Kelly

2004-01-01

A risk-based decision-making methodology conceived and developed at JPL and NASA has been used to aid in decision making for spacecraft technology assessment, adoption, development and operation. It takes a risk-centric perspective, through which risks are used as a reasoning step to interpose between mission objectives and risk mitigation measures.

The Recovery of TOMS-EP

NASA Technical Reports Server (NTRS)

Robertson, Brent; Sabelhaus, Phil; Mendenhall, Todd; Fesq, Lorraine

1998-01-01

On December 13th 1998, the Total Ozone Mapping Spectrometer - Earth Probe (TOMS-EP) spacecraft experienced a Single Event Upset which caused the system to reconfigure and enter a Safe Mode. This incident occurred two and a half years after the launch of the spacecraft which was designed for a two year life. A combination of factors, including changes in component behavior due to age and extended use, very unfortunate initial conditions and the safe mode processing logic prevented the spacecraft from entering its nominal long term storage mode. The spacecraft remained in a high fuel consumption mode designed for temporary use. By the time the onboard fuel was exhausted, the spacecraft was Sun pointing in a high rate flat spin. Although the uncontrolled spacecraft was initially in a power and thermal safe orientation, it would not stay in this state indefinitely due to a slow precession of its momentum vector. A recovery team was immediately assembled to determine if there was time to develop a method of despinning the vehicle and return it to normal science data collection. A three stage plan was developed that used the onboard magnetic torque rods as actuators. The first stage was designed to reduce the high spin rate to within the linear range of the gyros. The second stage transitioned the spacecraft from sun pointing to orbit reference pointing. The final stage returned the spacecraft to normal science operation. The entire recovery scenario was simulated with a wide range of initial conditions to establish the expected behavior. The recovery sequence was started on December 28th 1998 and completed by December 31st. TOMS-EP was successfully returned to science operations by the beginning of 1999. This paper describes the TOMS-EP Safe Mode design and the factors which led to the spacecraft anomaly and loss of fuel. The recovery and simulation efforts are described. Flight data are presented which show the performance of the spacecraft during its return to science. Finally, lessons learned are presented.

Low-Cost Approaches to Deep Space Missions

NASA Technical Reports Server (NTRS)

Squibb, G. F.; Edwards, C. D.; Schober, W. R.; Hooke, A. J.; Tai, W. S.; Pollmeier, V. M.

2000-01-01

The past decade has brought about a radical transformation in NASA's planetary exploration program. At the beginning of this decade, NASA was focused on the Cassini mission to Saturn. Following on the heels of the successful Voyager and Galileo missions, Cassini represents the culmination of an evolution towards successively larger, more complex, and more expensive spacecraft. The Cassini spacecraft weighs in at over 5 metric tons, and carries an entry probe and a sophisticated suite of sensors supporting 27 different science investigations enabling a comprehensive scientific investigation of Saturn with a single spacecraft. The cost of this spacecraft exceeded $2B, including the cost of the large Titan IV launch vehicle. During Cassini development, NASA realized that it could no longer afford these "flagship" missions, and the agency moved aggressively towards a "faster, better, cheaper" design philosophy of focused science goals and simpler, rapidly-developed spacecraft, allowing much more frequent launches of smaller, lower-cost missions. The Mars Global Surveyor, launched in November 1996, is an example of this new paradigm. Developed in less than 3-years, MGS is only one-fifth the mass of Cassini, and only cost on the order of $220M. The reduced spacecraft mass allows use of the smaller, lower cost Delta launch vehicle. Currently in orbit about Mars, MGS carries a focused suite of six science instruments that are currently returning high-resolution remote sensing of the Martian surface. The future calls for continued even more aggressive mass and cost targets. Examples of these next-generation goals are embodied in the Mars Micromission spacecraft concept, targeted for launch in 2003. With a mass of only 200kg, this lightweight bus can be tailored to carry a variety of payloads to Mars or other inner-planet destinations. The design of the Micromission spacecraft enable them to be launched at extremely low cost as a secondary "piggyback" payload.

Tracking and data systems support for the Helios project. Volume 1: Project development through end of mission, phase 2

NASA Technical Reports Server (NTRS)

Goodwin, P. S.; Traxler, M. R.; Meeks, W. G.; Flanagan, F. M.

1976-01-01

The overall evolution of the Helios Project is summarized from its conception through to the completion of the Helios-1 mission phase 2. Beginning with the project objectives and concluding with the Helios-1 spacecraft entering its first superior conjunction (end of mission phase 2), descriptions of the project, the mission and its phases, international management and interfaces, and Deep Space Network-spacecraft engineering development in telemetry, tracking, and command systems to ensure compatibility between the U.S. Deep Space Network and the German-built spacecraft are included.

Spacecraft-borne long life cryogenic refrigeration: Status and trends

NASA Technical Reports Server (NTRS)

Johnson, A. L.

1983-01-01

The status of cryogenic refrigerator development intended for, or possibly applicable to, long life spacecraft-borne application is reviewed. Based on these efforts, the general development trends are identified. Using currently projected technology needs, the various trends are compared and evaluated. The linear drive, non-contacting bearing Stirling cycle refrigerator concept appears to be the best current approach that will meet the technology projection requirements for spacecraft-borne cryogenic refrigerators. However, a multiply redundant set of lightweight, moderate life, moderate reliability Stirling cycle cryogenic refrigerators using high-speed linear drive and sliding contact bearings may possibly suffice.

Deployable Propulsion, Power and Communications Systems for Solar System Exploration

NASA Technical Reports Server (NTRS)

Johnson, L.; Carr, J.; Boyd, D.

2017-01-01

NASA is developing thin-film based, deployable propulsion, power, and communication systems for small spacecraft that could provide a revolutionary new capability allowing small spacecraft exploration of the solar system. By leveraging recent advancements in thin films, photovoltaics, and miniaturized electronics, new mission-level capabilities will be enabled aboard lower-cost small spacecraft instead of their more expensive, traditional counterparts, enabling a new generation of frequent, inexpensive deep space missions. Specifically, thin-film technologies are allowing the development and use of solar sails for propulsion, small, lightweight photovoltaics for power, and omnidirectional antennas for communication.

Enabling Higher Data Rates for Planetary Science Missions

NASA Astrophysics Data System (ADS)

Deutsch, L. J.; Townes, S. A.; Lazio, J.; Bell, D. J.; Chahat, N. E.; Kovalik, J. M.; Kuperman, I.; Sauder, J.; Liebrecht, P. E.

2017-12-01

The data rate from deep space spacecraft has increased by more than 10 orders of magnitude since the first lunar missions in the 1960s. The demand for increased data rates has stemmed from the increasing sophistication of the science questions being addressed and the concomitant increase in the complexity of the missions themselves (from fly-by to orbit to land and rove). Projections for the next few decades suggest the demand for data rates for deep space missions will continue to increase by approximately one order of magnitude every decade, driven by these same factors. Achieving higher data rates requires a partnership between the spacecraft and the ground system. We describe a series of technology developments for flight telecommunications systems, both at radio frequency (RF) and optical, to enable spacecraft to transmit and receive larger data volumes. These technology developments include deployable high gain antennas for small spacecraft, re-programmable software-defined radios, and optical communication packages designed for CubeSat form factors. The intent is that these developments would provide enhancements in capability for both spacecraft-Earth and spacecraft-spacecraft telecommunications. We also describe the future planning for NASA's Deep Space Network (DSN), which remains the prime conduit for data from all planetary science missions. Through a combination of new antennas and backends being installed over the next five years and incorporation of optical communications, the DSN aims to ensure that the historical improvements in data rates and volumes will continue for many decades. Part of this research was carried out at the Jet Propulsion Laboratory, California Institute of Technology, under a contract with the National Aeronautics and Space Administration.

Embedded Thermal Control for Subsystems for Next Generation Spacecraft Applications

NASA Technical Reports Server (NTRS)

Didion, Jeffrey R.

2015-01-01

Thermal Fluids and Analysis Workshop, Silver Spring MD NCTS 21070-15. NASA, the Defense Department and commercial interests are actively engaged in developing miniaturized spacecraft systems and scientific instruments to leverage smaller cheaper spacecraft form factors such as CubeSats. This paper outlines research and development efforts among Goddard Space Flight Center personnel and its several partners to develop innovative embedded thermal control subsystems. Embedded thermal control subsystems is a cross cutting enabling technology integrating advanced manufacturing techniques to develop multifunctional intelligent structures to reduce Size, Weight and Power (SWaP) consumption of both the thermal control subsystem and overall spacecraft. Embedded thermal control subsystems permit heat acquisition and rejection at higher temperatures than state of the art systems by employing both advanced heat transfer equipment (integrated heat exchangers) and high heat transfer phenomena. The Goddard Space Flight Center Thermal Engineering Branch has active investigations seeking to characterize advanced thermal control systems for near term spacecraft missions. The embedded thermal control subsystem development effort consists of fundamental research as well as development of breadboard and prototype hardware and spaceflight validation efforts. This paper will outline relevant fundamental investigations of micro-scale heat transfer and electrically driven liquid film boiling. The hardware development efforts focus upon silicon based high heat flux applications (electronic chips, power electronics etc.) and multifunctional structures. Flight validation efforts include variable gravity campaigns and a proposed CubeSat based flight demonstration of a breadboard embedded thermal control system. The CubeSat investigation is technology demonstration will characterize in long-term low earth orbit a breadboard embedded thermal subsystem and its individual components to develop optimized operational schema.

Viking lander spacecraft battery

NASA Technical Reports Server (NTRS)

Newell, D. R.

1976-01-01

The Viking Lander was the first spacecraft to fly a sterilized nickel-cadmium battery on a mission to explore the surface of a planet. The significant results of the battery development program from its inception through the design, manufacture, and test of the flight batteries which were flown on the two Lander spacecraft are documented. The flight performance during the early phase of the mission is also presented.

IUS solid rocket motor contamination prediction methods

NASA Technical Reports Server (NTRS)

Mullen, C. R.; Kearnes, J. H.

1980-01-01

A series of computer codes were developed to predict solid rocket motor produced contamination to spacecraft sensitive surfaces. Subscale and flight test data have confirmed some of the analytical results. Application of the analysis tools to a typical spacecraft has provided early identification of potential spacecraft contamination problems and provided insight into their solution; e.g., flight plan modifications, plume or outgassing shields and/or contamination covers.

Onboard Science Data Analysis: Opportunities, Benefits, and Effects on Mission Design

NASA Technical Reports Server (NTRS)

Stolorz, P.; Cheeseman, P.

1998-01-01

Much of the initial focus for spacecraft autonomy has been on developing new software and systems concepts to automate engineering functions of the spacecraft: guidance, navigation and control, fault protection, and resources management. However, the ultimate objectives of NASA missions are science objectives, which implies that we need a new framework for perfoming science data evaluation and observation planning autonomously onboard spacecraft.

Viking Lander spacecraft battery. Final report

DOE Office of Scientific and Technical Information (OSTI.GOV)

Newell, D.R.

1976-09-01

The Viking Lander was the first spacecraft to fly a sterilized nickel--cadmium battery on a mission to explore the surface of a planet. The significant results of the battery development program from its inception through the design, manufacture, and test of the flight batteries which were flown on the two Lander spacecraft are documented. The flight performance during the early phase of the mission is also presented.

The 1981 Goddard Space Flight Center Battery Workshop

NASA Technical Reports Server (NTRS)

Halpert, G. (Editor)

1982-01-01

Results of testing, analysis, and development of lithium, nickel-cadmium, and nickel-hydrogen batteries are reported. Focus is on the improvement of power systems in the areas of high capacity, high energy density, and long cycle and storage life. Applications of these batteries as spacecraft power supplies are discussed. Those spacecraft include deepspace probes, spacecraft in geostationary orbit, and large space systems in low-Earth orbit.

Development and fabrication of a chargeable magnet system for spacecraft control

NASA Technical Reports Server (NTRS)

1972-01-01

The design of variable permanent magnets for use in magnetic balancing and control of earth orbiting spacecraft is discussed. These magnets can be used instead of air coils or electromagnets in applications where the objective is to produce, or eliminate, torque on the spacecraft through interaction with the earth's magnetic field. The configuration of the magnet for minimum size and weight is described.

Application of partial differential equation modeling of the control/structural dynamics of flexible spacecraft

NASA Technical Reports Server (NTRS)

Taylor, Lawrence W., Jr.; Rajiyah, H.

1991-01-01

Partial differential equations for modeling the structural dynamics and control systems of flexible spacecraft are applied here in order to facilitate systems analysis and optimization of these spacecraft. Example applications are given, including the structural dynamics of SCOLE, the Solar Array Flight Experiment, the Mini-MAST truss, and the LACE satellite. The development of related software is briefly addressed.

Development of a Robust star identification technique for use in attitude determination of the ACE spacecraft

NASA Technical Reports Server (NTRS)

Woodard, Mark; Rohrbaugh, Dave

1995-01-01

The Advanced Composition Explorer (ACE) spacecraft is designed to fly in a spin-stabilized attitude. The spacecraft will carry two attitude sensors - a digital fine Sun sensor and a charge coupled device (CCD) star tracker - to allow ground-based determination of the spacecraft attitude and spin rate. Part of the processing that must be performed on the CCD star tracker data is the star identification. Star data received from the spacecraft must be matched with star information in the SKYMAP catalog to determine exactly which stars the sensor is tracking. This information, along with the Sun vector measured by the Sun sensor, is used to determine the spacecraft attitude. Several existing star identification (star ID) systems were examined to determine whether they could be modified for use on the ACE mission. Star ID systems which exist for three-axis stabilized spacecraft tend to be complex in nature and many require fairly good knowledge of the spacecraft attitude, making their use for ACE excessive. Star ID systems used for spinners carrying traditional slit star sensors would have to be modified to model the CCD star tracker. The ACE star ID algorithm must also be robust, in that it will be able to correctly identify stars even though the attitude is not known to a high degree of accuracy, and must be very efficient to allow real-time star identification. The paper presents the star ID algorithm that was developed for ACE. Results from prototype testing are also presented to demonstrate the efficiency, accuracy, and robustness of the algorithm.

Ground Data System Risk Mitigation Techniques for Faster, Better, Cheaper Missions

NASA Technical Reports Server (NTRS)

Catena, John J.; Saylor, Rick; Casasanta, Ralph; Weikel, Craig; Powers, Edward I. (Technical Monitor)

2000-01-01

With the advent of faster, cheaper, and better missions, NASA Projects acknowledged that a higher level of risk was inherent and accepted with this approach. It was incumbent however upon each component of the Project whether spacecraft, payload, launch vehicle, or ground data system to ensure that the mission would nevertheless be an unqualified success. The Small Explorer (SMEX) program's ground data system (GDS) team developed risk mitigation techniques to achieve these goals starting in 1989. These techniques have evolved through the SMEX series of missions and are practiced today under the Triana program. These techniques are: (1) Mission Team Organization--empowerment of a closeknit ground data system team comprising system engineering, software engineering, testing, and flight operations personnel; (2) Common Spacecraft Test and Operational Control System--utilization of the pre-launch spacecraft integration system as the post-launch ground data system on-orbit command and control system; (3) Utilization of operations personnel in pre-launch testing--making the flight operations team an integrated member of the spacecraft testing activities at the beginning of the spacecraft fabrication phase; (4) Consolidated Test Team--combined system, mission readiness and operations testing to optimize test opportunities with the ground system and spacecraft; and (5). Reuse of Spacecraft, Systems and People--reuse of people, software and on-orbit spacecraft throughout the SMEX mission series. The SMEX ground system development approach for faster, cheaper, better missions has been very successful. This paper will discuss these risk management techniques in the areas of ground data system design, implementation, test, and operational readiness.

Microwave reflectometer ionization sensor

NASA Technical Reports Server (NTRS)

Seals, Joseph; Fordham, Jeffrey A.; Pauley, Robert G.; Simonutti, Mario D.

1993-01-01

The development of the Microwave Reflectometer Ionization Sensor (MRIS) Instrument for use on the Aeroassist Flight Experiment (AFE) spacecraft is described. The instrument contract was terminated, due to cancellation of the AFE program, subsequent to testing of an engineering development model. The MRIS, a four-frequency reflectometer, was designed for the detection and location of critical electron density levels in spacecraft reentry plasmas. The instrument would sample the relative magnitude and phase of reflected signals at discrete frequency steps across 4 GHz bandwidths centered at four frequencies: 20, 44, 95, and 140 GHz. The sampled data would be stored for later processing to calculate the distance from the spacecraft surface to the critical electron densities versus time. Four stepped PM CW transmitter receivers were located behind the thermal protection system of the spacecraft with horn antennas radiating and receiving through an insulating tile. Techniques were developed to deal with interference, including multiple reflections and resonance effects, resulting from the antenna configuration and operating environment.