Gravity-gradient dynamics experiments performed in orbit utilizing the Radio Astronomy Explorer (RAE-1) spacecraft

Science.gov (United States)

Walden, H.

1973-01-01

Six dynamic experiments were performed in earth orbit utilizing the RAE spacecraft in order to test the accuracy of the mathematical model of RAE dynamics. The spacecraft consisted of four flexible antenna booms, mounted on a rigid cylindrical spacecraft hub at center, for measuring radio emissions from extraterrestrial sources. Attitude control of the gravity stabilized spacecraft was tested by using damper clamping, single lower leading boom operations, and double lower boom operations. Results and conclusions of the in-orbit dynamic experiments proved the accuracy of the analytic techniques used to model RAE dynamical behavior.

On-orbit assembly of a team of flexible spacecraft using potential field based method

Science.gov (United States)

Chen, Ti; Wen, Hao; Hu, Haiyan; Jin, Dongping

2017-04-01

In this paper, a novel control strategy is developed based on artificial potential field for the on-orbit autonomous assembly of four flexible spacecraft without inter-member collision. Each flexible spacecraft is simplified as a hub-beam model with truncated beam modes in the floating frame of reference and the communication graph among the four spacecraft is assumed to be a ring topology. The four spacecraft are driven to a pre-assembly configuration first and then to the assembly configuration. In order to design the artificial potential field for the first step, each spacecraft is outlined by an ellipse and a virtual leader of circle is introduced. The potential field mainly depends on the attitude error between the flexible spacecraft and its neighbor, the radial Euclidian distance between the ellipse and the circle and the classical Euclidian distance between the centers of the ellipse and the circle. It can be demonstrated that there are no local minima for the potential function and the global minimum is zero. If the function is equal to zero, the solution is not a certain state, but a set. All the states in the set are corresponding to the desired configurations. The Lyapunov analysis guarantees that the four spacecraft asymptotically converge to the target configuration. Moreover, the other potential field is also included to avoid the inter-member collision. In the control design of the second step, only small modification is made for the controller in the first step. Finally, the successful application of the proposed control law to the assembly mission is verified by two case studies.

Guidance and Navigation for Rendezvous and Proximity Operations with a Non-Cooperative Spacecraft at Geosynchronous Orbit

Science.gov (United States)

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.

LO2/LH2 propulsion for outer planet orbiter spacecraft

Science.gov (United States)

Garrison, P. W.; Sigurdson, K. B.

1983-01-01

Galileo class orbiter missions (750-1500 kg) to the outer planets require a large postinjection delta-V for improved propulsion performance. The present investigation shows that a pump-fed low thrust LO2/LH2 propulsion system can provide a significantly larger net on-orbit mass for a given delta-V than a state-of-the-art earth storable, N2O4/monomethylhydrazine pressure-fed propulsion system. A description is given of a conceptual design for a LO2/LH2 pump-fed propulsion system developed for a Galileo class mission to the outer planets. Attention is given to spacecraft configuration, details regarding the propulsion system, the thermal control of the cryogenic propellants, and aspects of mission performance.

Single-photon double ionization: renormalized-natural-orbital theory versus multi-configurational Hartree–Fock

International Nuclear Information System (INIS)

Brics, M; Rapp, J; Bauer, D

2017-01-01

The N -particle wavefunction has too many dimensions for a direct time propagation of a many-body system according to the time-dependent Schrödinger equation (TDSE). On the other hand, time-dependent density functional theory (TDDFT) tells us that the single-particle density is, in principle, sufficient. However, a practicable equation of motion for the accurate time evolution of the single-particle density is unknown. It is thus an obvious idea to propagate a quantity which is not as reduced as the single-particle density but less dimensional than the N -body wavefunction. Recently, we have introduced time-dependent renormalized-natural-orbital theory (TDRNOT). TDRNOT is based on the propagation of the eigenfunctions of the one-body reduced density matrix, the so-called natural orbitals. In this paper we demonstrate how TDRNOT is related to the multi-configurational time-dependent Hartree–Fock (MCTDHF) approach. We also compare the performance of MCTDHF and TDRNOT versus the TDSE for single-photon double ionization (SPDI) of a 1D helium model atom. SPDI is one of the effects where TDDFT does not work in practice, especially if one is interested in correlated photoelectron spectra, for which no explicit density functional is known. (paper)

A Concept of Two-Stage-To-Orbit Reusable Launch Vehicle

Science.gov (United States)

Yang, Yong; Wang, Xiaojun; Tang, Yihua

2002-01-01

Reusable Launch Vehicle (RLV) has a capability of delivering a wide rang of payload to earth orbit with greater reliability, lower cost, more flexibility and operability than any of today's launch vehicles. It is the goal of future space transportation systems. Past experience on single stage to orbit (SSTO) RLVs, such as NASA's NASP project, which aims at developing an rocket-based combined-cycle (RBCC) airplane and X-33, which aims at developing a rocket RLV, indicates that SSTO RLV can not be realized in the next few years based on the state-of-the-art technologies. This paper presents a concept of all rocket two-stage-to-orbit (TSTO) reusable launch vehicle. The TSTO RLV comprises an orbiter and a booster stage. The orbiter is mounted on the top of the booster stage. The TSTO RLV takes off vertically. At the altitude about 50km the booster stage is separated from the orbiter, returns and lands by parachutes and airbags, or lands horizontally by means of its own propulsion system. The orbiter continues its ascent flight and delivers the payload into LEO orbit. After completing orbit mission, the orbiter will reenter into the atmosphere, automatically fly to the ground base and finally horizontally land on the runway. TSTO RLV has less technology difficulties and risk than SSTO, and maybe the practical approach to the RLV in the near future.

A simple method to design non-collision relative orbits for close spacecraft formation flying

Science.gov (United States)

Jiang, Wei; Li, JunFeng; Jiang, FangHua; Bernelli-Zazzera, Franco

2018-05-01

A set of linearized relative motion equations of spacecraft flying on unperturbed elliptical orbits are specialized for particular cases, where the leader orbit is circular or equatorial. Based on these extended equations, we are able to analyze the relative motion regulation between a pair of spacecraft flying on arbitrary unperturbed orbits with the same semi-major axis in close formation. Given the initial orbital elements of the leader, this paper presents a simple way to design initial relative orbital elements of close spacecraft with the same semi-major axis, thus preventing collision under non-perturbed conditions. Considering the mean influence of J 2 perturbation, namely secular J 2 perturbation, we derive the mean derivatives of orbital element differences, and then expand them to first order. Thus the first order expansion of orbital element differences can be added to the relative motion equations for further analysis. For a pair of spacecraft that will never collide under non-perturbed situations, we present a simple method to determine whether a collision will occur when J 2 perturbation is considered. Examples are given to prove the validity of the extended relative motion equations and to illustrate how the methods presented can be used. The simple method for designing initial relative orbital elements proposed here could be helpful to the preliminary design of the relative orbital elements between spacecraft in a close formation, when collision avoidance is necessary.

On the spacecraft attitude stabilization in the orbital frame

Directory of Open Access Journals (Sweden)

Antipov Kirill A.

2012-01-01

Full Text Available The paper deals with spacecraft in the circular near-Earth orbit. The spacecraft interacts with geomagnetic field by the moments of Lorentz and magnetic forces. The octupole approximation of the Earth’s magnetic field is accepted. The spacecraft electromagnetic parameters, namely the electrostatic charge moment of the first order and the eigen magnetic moment are the controlled quasiperiodic functions. The control algorithms for the spacecraft electromagnetic parameters, which allows to stabilize the spacecraft attitude position in the orbital frame are obtained. The stability of the spacecraft stabilized orientation is proved both analytically and by PC computations.

Investigation of different ethylenediamine-N,N'-disuccinic acid-enhanced washing configurations for remediation of a Cu-contaminated soil: process kinetics and efficiency comparison between single-stage and multi-stage configurations.

Science.gov (United States)

Ferraro, Alberto; Fabbricino, Massimiliano; van Hullebusch, Eric D; Esposito, Giovanni

2017-09-01

A comparison of Cu extraction yields for three different ethylenediamine-N,N'-disuccinic acid (EDDS)-enhanced washing configurations was performed on a Cu-contaminated soil. Batch experiments were used to simulate a single-stage continuous stirred tank reactor (CSTR) and a multi-stage (side feeding and counter-current) reactor. Single-stage CSTR conditions were simulated for various EDDS:(Cu + Cd + Pb + Co + Ni + Zn) molar ratio (EDDS:M ratio) (from 1 to 30) and liquid to soil (L/S) ratio (from 15 to 45). The highest Cu extraction yield (≃56%) was achieved with EDDS:M = 30. In contrast, a Cu extraction yield decrease was observed with increasing L/S ratio with highest extracted Cu achievement (≃48%) for L/S = 15. Side feeding configuration was tested in four experimental conditions through different fractionation mode of EDDS dose and treatment time at each washing step. Results from the four tests showed all enhanced Cu extraction (maximum values from ≃43 to ≃51%) achieved at lower treatment time and lower EDDS:M molar ratio compared to CSTR configuration with L/S = 25 and EDDS:M = 10. The counter-current washing was carried out through two washing flows achieving a process performance enhancement with 27% increase of extracted Cu compared to single-stage CSTR configuration. Higher Cu extraction percentage (36.8%) was observed in the first washing phase than in the second one (24.7%).

Implementation of a Single-Stage-To-Orbit (SSTO) model for stability and control analysis

Science.gov (United States)

Ingalls, Stephen A.

1995-07-01

Three NASA centers: Marshall Space Flight Center (MSFC), Langley Research Center (LaRC), and Johnson Space Center (JSC) are currently involved in studying a family of single-stage- and two-stage-to-orbit (SSTO/TSTO) vehicles to serve as the next generation space transportation system (STS). A rocketed winged-body is the current focus. The configuration (WB001) is a vertically-launched, horizontally-landing system with circular cross-section. Preliminary aerodynamic data was generated by LaRC and is a combination of wind-tunnel data, empirical methods, and Aerodynamic Preliminary Analysis System-(APAS) generated values. JSC's efforts involve descent trajectory design, stability analysis, and flight control system synthesis. Analysis of WB001's static stability indicates instability in 'tuck' (C(sub mu) less than 0: Mach = 0.30, alpha greater than 3.25 deg; Mach = 0.60, alpha greater than 8.04), an unstable dihedral effects (C(sub l(beta)) greater than 0: Mach = 30,alpha less than 12 deg.; Mach = 0.60, alpha less than 10.00 deg.), and, most significantly, an unstable weathercock stability derivative, C(sub n(beta)), at all angles of attack and subsonic Mach numbers. Longitudinal trim solutions for Mach = 0.30 and 0.60 indicate flight path angle possibilities ranging from around 12 (M = 0.30) to slightly over 20 degrees at Mach = 0.60. Trim angles of attack increase from 6.24 at Mach 0.60 and 10,000 feet to 17.7 deg. at Mach 0.30, sea-level. Lateral trim was attempted for a design cross-wind of 25.0 knots. The current vehicle aerodynamic and geometric characteristics will only yield a lateral trim solution at impractical tip-fin deflections (approximately equal to 43 deg.) and bank angles (21 deg.). A study of the lateral control surfaces, tip-fin controllers for WB001, indicate increased surface area would help address these instabilities, particularly the deficiency in C(sub n(beta)), but obviously at the expense of increased vehicle weight. Growth factors of

Global optimum spacecraft orbit control subject to bounded thrust in presence of nonlinear and random disturbances in a low earth orbit

Directory of Open Access Journals (Sweden)

Tamer Mekky Ahmed Habib

2012-06-01

Full Text Available The primary objective of this work is to develop an effective spacecraft orbit control algorithm suitable for spacecraft orbital maneuver and/or rendezvous. The actual governing equation of a spacecraft orbiting the earth is merely nonlinear. Disturbance forces resulting from aerodynamic drag, oblateness of the earth till the fourth order (i.e. J4, and random disturbances are modeled for the initial and target orbits. These disturbances increase the complexity of nonlinear governing equations. Global optimum solutions of the control algorithm parameters are determined throughout real coded genetic algorithms such that the steady state difference between the actual and desired trajectories is minimized. The resulting solutions are constrained to avoid spacecraft collision with the surface of the earth taking into account limited thrust budget.

Orbit Determination of Spacecraft in Earth-Moon L1 and L2 Libration Point Orbits

Science.gov (United States)

Woodard, Mark; Cosgrove, Daniel; Morinelli, Patrick; Marchese, Jeff; Owens, Brandon; Folta, David

2011-01-01

The ARTEMIS mission, part of the THEMIS extended mission, is the first to fly spacecraft in the Earth-Moon Lissajous regions. In 2009, two of the five THEMIS spacecraft were redeployed from Earth-centered orbits to arrive in Earth-Moon Lissajous orbits in late 2010. Starting in August 2010, the ARTEMIS P1 spacecraft executed numerous stationkeeping maneuvers, initially maintaining a lunar L2 Lissajous orbit before transitioning into a lunar L1 orbit. The ARTEMIS P2 spacecraft entered a L1 Lissajous orbit in October 2010. In April 2011, both ARTEMIS spacecraft will suspend Lissajous stationkeeping and will be maneuvered into lunar orbits. The success of the ARTEMIS mission has allowed the science team to gather unprecedented magnetospheric measurements in the lunar Lissajous regions. In order to effectively perform lunar Lissajous stationkeeping maneuvers, the ARTEMIS operations team has provided orbit determination solutions with typical accuracies on the order of 0.1 km in position and 0.1 cm/s in velocity. The ARTEMIS team utilizes the Goddard Trajectory Determination System (GTDS), using a batch least squares method, to process range and Doppler tracking measurements from the NASA Deep Space Network (DSN), Berkeley Ground Station (BGS), Merritt Island (MILA) station, and United Space Network (USN). The team has also investigated processing of the same tracking data measurements using the Orbit Determination Tool Kit (ODTK) software, which uses an extended Kalman filter and recursive smoother to estimate the orbit. The orbit determination results from each of these methods will be presented and we will discuss the advantages and disadvantages associated with using each method in the lunar Lissajous regions. Orbit determination accuracy is dependent on both the quality and quantity of tracking measurements, fidelity of the orbit force models, and the estimation techniques used. Prior to Lissajous operations, the team determined the appropriate quantity of tracking

The SPQR experiment: detecting damage to orbiting spacecraft with ground-based telescopes

Science.gov (United States)

Paolozzi, Antonio; Porfilio, Manfredi; Currie, Douglas G.; Dantowitz, Ronald F.

2007-09-01

The objective of the Specular Point-like Quick Reference (SPQR) experiment was to evaluate the possibility of improving the resolution of ground-based telescopic imaging of manned spacecraft in orbit. The concept was to reduce image distortions due to atmospheric turbulence by evaluating the Point Spread Function (PSF) of a point-like light reference and processing the spacecraft image accordingly. The target spacecraft was the International Space Station (ISS) and the point-like reference was provided by a laser beam emitted by the ground station and reflected back to the telescope by a Cube Corner Reflector (CCR) mounted on an ISS window. The ultimate objective of the experiment was to demonstrate that it is possible to image spacecraft in Low Earth Orbit (LEO) with a resolution of 20 cm, which would have probably been sufficient to detect the damage which caused the Columbia disaster. The experiment was successfully performed from March to May 2005. The paper provides an overview of the SPQR experiment.

Effects of the Eccentricity of a Perturbing Third Body on the Orbital Correction Maneuvers of a Spacecraft

Directory of Open Access Journals (Sweden)

R. C. Domingos

2014-01-01

Full Text Available The fuel consumption required by the orbital maneuvers when correcting perturbations on the orbit of a spacecraft due to a perturbing body was estimated. The main goals are the measurement of the influence of the eccentricity of the perturbing body on the fuel consumption required by the station keeping maneuvers and the validation of the averaged methods when applied to the problem of predicting orbital maneuvers. To study the evolution of the orbits, the restricted elliptic three-body problem and the single- and double-averaged models are used. Maneuvers are made by using impulsive and low thrust maneuvers. The results indicated that the averaged models are good to make predictions for the orbital maneuvers when the spacecraft is in a high inclined orbit. The eccentricity of the perturbing body plays an important role in increasing the effects of the perturbation and the fuel consumption required for the station keeping maneuvers. It is shown that the use of more frequent maneuvers decreases the annual cost of the station keeping to correct the orbit of a spacecraft. An example of an eccentric planetary system of importance to apply the present study is the dwarf planet Haumea and its moons, one of them in an eccentric orbit.

Solar Array Disturbances to Spacecraft Pointing During the Lunar Reconnaissance Orbiter (LRO) Mission

Science.gov (United States)

Calhoun, Philip

2010-01-01

The Lunar Reconnaissance Orbiter (LRO), the first spacecraft to support NASA s return to the Moon, launched on June 18, 2009 from the Cape Canaveral Air Force Station aboard an Atlas V launch vehicle. It was initially inserted into a direct trans-lunar trajectory to the Moon. After a five day transit to the Moon, LRO was inserted into the Lunar orbit and successfully lowered to a low altitude elliptical polar orbit for spacecraft commissioning. Successful commissioning was completed in October 2009 when LRO was placed in its near circular mission orbit with an approximate altitude of 50km. LRO will spend at least one year orbiting the Moon, collecting lunar environment science and mapping data, utilizing a suite of seven instruments to enable future human exploration. The objective is to provide key science data necessary to facilitate human return to the Moon as well as identification of opportunities for future science missions. LRO's instrument suite will provide the high resolution imaging data with sub-meter accuracy, highly accurate lunar cartographic maps, mineralogy mapping, amongst other science data of interest. LRO employs a 3-axis stabilized attitude control system (ACS) whose primary control mode, the "Observing Mode", provides Lunar nadir, off-nadir, and inertial fine pointing for the science data collection and instrument calibration. This controller combines the capability of fine pointing with on-demand large angle full-sky attitude reorientation. It provides simplicity of spacecraft operation as well as additional flexibility for science data collection. A conventional suite of ACS components is employed in the Observing Mode to meet the pointing and control objectives. Actuation is provided by a set of four reaction wheels developed in-house at NASA Goddard Space Flight Center (GSFC). Attitude feedback is provided by a six state Kalman filter which utilizes two SELEX Galileo Star Trackers for attitude updates, and a single Honeywell Miniature

Feasibility of a responsive, hybrid propulsion augmented, Vertical-Takeoff-and-Landing, Single-Stage-to-Orbit launch system

Science.gov (United States)

Pelaccio, Dennis G.

1996-03-01

A novel, reusable, Vertical-Takeoff-and-Landing, Single-Stage-to-Orbit (VTOL/SSTO) launch system concept, named HYP-SSTO, is presented in this paper. This launch vehicle system concept uses a highly coupled, main high performance liquid oxygen/liquid hydrogen (LOX/LH2) propulsion system, that is used only for launch, with a hybrid auxiliary propulsion system which is used during final orbit insertion, major orbit maneuvering, and landing propulsive burn phases of flight. By using a hybrid propulsion system for major orbit maneuver burns and landing, this launch system concept has many advantages over conventional VTOL/SSTO concepts that use LOX/LH2 propulsion system(s) burns for all phases of flight. Because hybrid propulsion systems are relatively simple and inert by their nature, this concept has the potential to support short turnaround times between launches, be economical to develop, and be competitive in terms of overall system life-cycle cost. This paper provides a technical description of the novel, reusable HYP-SSTO launch system concept. Launch capability performance, as well as major design and operational system attributes, are identified and discussed.

Spacecraft on-orbit deployment anomalies - What can be done?

Science.gov (United States)

Freeman, Michael T.

1993-04-01

Modern communications satellites rely heavily upon deployable appendage (i.e. solar arrays, communications antennas, etc.) to perform vital functions that enable the spacecraft to effectively conduct mission objectives. Communications and telemetry antennas provide the radiofrequency link between the spacecraft and the earth ground station, permitting data to be transmitted and received from the satellite. Solar arrays serve as the principle source of electrical energy to the satellite, and recharge internal batteries during operation. However, since satellites cannot carry backup systems, if a solar array fails to deploy, the mission is lost. This article examines the subject of on-orbit anomalies related to the deployment of spacecraft appendage, and possible causes of such failures. Topics discussed shall include mechanical launch loading, on-orbit thermal and solar concerns, reliability of spacecraft pyrotechnics, and practical limitations of ground-based deployment testing. Of particular significance, the article will feature an in-depth look at the lessons learned from the successful recovery of the Telesat Canada Anik-E2 satellite in 1991.

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

Science.gov (United States)

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.

Generating Animated Displays of Spacecraft Orbits

Science.gov (United States)

Candey, Robert M.; Chimiak, Reine A.; Harris, Bernard T.

2005-01-01

Tool for Interactive Plotting, Sonification, and 3D Orbit Display (TIPSOD) is a computer program for generating interactive, animated, four-dimensional (space and time) displays of spacecraft orbits. TIPSOD utilizes the programming interface of the Satellite Situation Center Web (SSCWeb) services to communicate with the SSC logic and database by use of the open protocols of the Internet. TIPSOD is implemented in Java 3D and effects an extension of the preexisting SSCWeb two-dimensional static graphical displays of orbits. Orbits can be displayed in any or all of the following seven reference systems: true-of-date (an inertial system), J2000 (another inertial system), geographic, geomagnetic, geocentric solar ecliptic, geocentric solar magnetospheric, and solar magnetic. In addition to orbits, TIPSOD computes and displays Sibeck's magnetopause and Fairfield's bow-shock surfaces. TIPSOD can be used by the scientific community as a means of projection or interpretation. It also has potential as an educational tool.

A review of findings of a study of rocket based combined cycle engines applied to extensively axisymmetric single stage to orbit vehicles

Science.gov (United States)

Foster, Richard W.

1992-01-01

Extensively axisymmetric and non-axisymmetric Single Stage To Orbit (SSTO) vehicles are considered. The information is presented in viewgraph form and the following topics are presented: payload comparisons; payload as a percent of dry weight - a system hardware cost indicator; life cycle cost estimations; operations and support costs estimation; selected engine type; and rocket engine specific impulse calculation.

Comparison of technologies for deorbiting spacecraft from low-earth-orbit at end of mission

Science.gov (United States)

Sánchez-Arriaga, G.; Sanmartín, J. R.; Lorenzini, E. C.

2017-09-01

An analytical comparison of four technologies for deorbiting spacecraft from Low-Earth-Orbit at end of mission is presented. Basic formulas based on simple physical models of key figures of merit for each device are found. Active devices - rockets and electrical thrusters - and passive technologies - drag augmentation devices and electrodynamic tethers - are considered. A basic figure of merit is the deorbit device-to-spacecraft mass ratio, which is, in general, a function of environmental variables, technology development parameters and deorbit time. For typical state-of-the-art values, equal deorbit time, middle inclination and initial altitude of 850 km, the analysis indicates that tethers are about one and two orders of magnitude lighter than active technologies and drag augmentation devices, respectively; a tether needs a few percent mass-ratio for a deorbit time of a couple of weeks. For high inclination, the performance drop of the tether system is moderate: mass ratio and deorbit time increase by factors of 2 and 4, respectively. Besides collision risk with other spacecraft and system mass considerations, such as main driving factors for deorbit space technologies, the analysis addresses other important constraints, like deorbit time, system scalability, manoeuver capability, reliability, simplicity, attitude control requirement, and re-entry and multi-mission capability (deorbit and re-boost) issues. The requirements and constraints are used to make a critical assessment of the four technologies as functions of spacecraft mass and initial orbit (altitude and inclination). Emphasis is placed on electrodynamic tethers, including the latest advances attained in the FP7/Space project BETs. The superiority of tape tethers as compared to round and multi-line tethers in terms of deorbit mission performance is highlighted, as well as the importance of an optimal geometry selection, i.e. tape length, width, and thickness, as function of spacecraft mass and initial

SAMSI: An orbiting spatial interferometer for micro-arc second astronomical observations. [Spacecraft Array for Michelson Spatial Interferometry (SAMSI)

Science.gov (United States)

Stachnik, R. V.; Gezari, D. Y.

1985-01-01

The concept and performance of (SAMSI) Spacecraft Array for Michelson Spatial Interferometry, an orbiting spatial interferometer comprised of three free-flying spacecraft, two collector telescopes and a central mixing station are described. In the one-dimensional interferometry mode orbits exist which provide natural scanning of the baseline. These orbits place extremely small demands on thrusters and fuel consumption. Resolution of 0.00001 arcsecond and magnitude limits of mv = 15 to 20 are achievable in a single orbit. In the imaging mode, SAMSI could synthesize images equivalent to those produced by equal diameter filled apertures in space, making use of the fuel resupply capability of a space station. Simulations indicate that image reconstruction can be performed with milliarcsecond resolution to a visual magnitude 12 in 12 hr of spiral scanning integration time.

Design and analysis of a single stage to orbit nuclear thermal rocket reactor engine

Energy Technology Data Exchange (ETDEWEB)

Labib, Satira, E-mail: Satira.Labib@duke-energy.com; King, Jeffrey, E-mail: kingjc@mines.edu

2015-06-15

Graphical abstract: - Highlights: • Three NTR reactors are optimized for the single stage launch of 1–15 MT payloads. • The proposed rocket engines have specific impulses in excess of 700 s. • Reactivity and submersion criticality requirements are satisfied for each reactor. - Abstract: Recent advances in the development of high power density fuel materials have renewed interest in nuclear thermal rockets (NTRs) as a viable propulsion technology for future space exploration. This paper describes the design of three NTR reactor engines designed for the single stage to orbit launch of payloads from 1 to 15 metric tons. Thermal hydraulic and rocket engine analyses indicate that the proposed rocket engines are able to reach specific impulses in excess of 800 s. Neutronics analyses performed using MCNP5 demonstrate that the hot excess reactivity, shutdown margin, and submersion criticality requirements are satisfied for each NTR reactor. The reactors each consist of a 40 cm diameter core packed with hexagonal tungsten cermet fuel elements. The core is surrounded by radial and axial beryllium reflectors and eight boron carbide control drums. The 40 cm long reactor meets the submersion criticality requirements (a shutdown margin of at least $1 subcritical in all submersion scenarios) with no further modifications. The 80 and 120 cm long reactors include small amounts of gadolinium nitride as a spectral shift absorber to keep them subcritical upon submersion in seawater or wet sand following a launch abort.

Design and analysis of a single stage to orbit nuclear thermal rocket reactor engine

International Nuclear Information System (INIS)

Labib, Satira; King, Jeffrey

2015-01-01

Graphical abstract: - Highlights: • Three NTR reactors are optimized for the single stage launch of 1–15 MT payloads. • The proposed rocket engines have specific impulses in excess of 700 s. • Reactivity and submersion criticality requirements are satisfied for each reactor. - Abstract: Recent advances in the development of high power density fuel materials have renewed interest in nuclear thermal rockets (NTRs) as a viable propulsion technology for future space exploration. This paper describes the design of three NTR reactor engines designed for the single stage to orbit launch of payloads from 1 to 15 metric tons. Thermal hydraulic and rocket engine analyses indicate that the proposed rocket engines are able to reach specific impulses in excess of 800 s. Neutronics analyses performed using MCNP5 demonstrate that the hot excess reactivity, shutdown margin, and submersion criticality requirements are satisfied for each NTR reactor. The reactors each consist of a 40 cm diameter core packed with hexagonal tungsten cermet fuel elements. The core is surrounded by radial and axial beryllium reflectors and eight boron carbide control drums. The 40 cm long reactor meets the submersion criticality requirements (a shutdown margin of at least $1 subcritical in all submersion scenarios) with no further modifications. The 80 and 120 cm long reactors include small amounts of gadolinium nitride as a spectral shift absorber to keep them subcritical upon submersion in seawater or wet sand following a launch abort

Overview of SDCM - The Spacecraft Design and Cost Model

Science.gov (United States)

Ferebee, Melvin J.; Farmer, Jeffery T.; Andersen, Gregory C.; Flamm, Jeffery D.; Badi, Deborah M.

1988-01-01

The Spacecraft Design and Cost Model (SDCM) is a computer-aided design and analysis tool for synthesizing spacecraft configurations, integrating their subsystems, and generating information concerning on-orbit servicing and costs. SDCM uses a bottom-up method in which the cost and performance parameters for subsystem components are first calculated; the model then sums the contributions from individual components in order to obtain an estimate of sizes and costs for each candidate configuration within a selected spacecraft system. An optimum spacraft configuration can then be selected.

Orbital single particle tracking on a commercial confocal microscope using piezoelectric stage feedback

International Nuclear Information System (INIS)

Lanzanò, L; Gratton, E

2014-01-01

Single Particle Tracking (SPT) is a technique used to locate fluorescent particles with nanometer precision. In the orbital tracking method the position of a particle is obtained analyzing the distribution of intensity along a circular orbit scanned around the particle. In combination with an active feedback this method allows tracking of particles in 2D and 3D with millisecond temporal resolution. Here we describe a SPT setup based on a feedback approach implemented with minimal modification of a commercially available confocal laser scanning microscope, the Zeiss LSM 510, in combination with an external piezoelectric stage scanner. The commercial microscope offers the advantage of a user-friendly software interface and pre-calibrated hardware components. The use of an external piezo-scanner allows the addition of feedback into the system but also represents a limitation in terms of its mechanical response. We describe in detail this implementation of the orbital tracking method and discuss advantages and limitations. As an example of application to live cell experiments we perform the 3D tracking of acidic vesicles in live polarized epithelial cells. (paper)

A 4-m evolvable space telescope configured for NASA's HabEx Mission: the initial stage of LUVOIR

Science.gov (United States)

Lillie, Charles F.; MacEwen, Howard A.; Polidan, Ronald S.; Breckinridge, James B.

2017-09-01

Previous papers have described our concept for a large telescope that would be assembled in space in several stages (in different configurations) over a period of fifteen to 20 years. Spreading the telescope development, launch and operations cost over 20 years would minimize the impact on NASA's annual budget and drastically shorten the time between program start and "first light" for this space observatory. The first Stage of this Evolvable Space Telescope (EST) would consist of an instrument module located at the prime focus of three 4-meter hexagonal mirrors arranged in a semi-circle to form one-half of a 12-m segmented mirror. After several years three additional 4-m mirrors would be added to create a 12-m filled aperture. Later, twelve more 4-m mirrors will be added to this Stage 2 telescope to create a 20-m filled aperture space telescope. At each stage the telescope would have an unparalleled capability for UVOIR observations, and the results of these observations will guide the evolution of the telescope and its instruments. In this paper we describe our design concept for an initial configuration of our Evolvable Space Telescope that can meet the requirements of the 4-m version of the HabEx spacecraft currently under consideration by NASA's Habitable Exoplanet Science and Technology Definition Team. This "Stage Zero" configuration will have only one 4-m mirror segment with the same 30-m focal length and a prime focus coronagraph with normal incidence optics to minimize polarization effects. After assembly and checkout in cis-lunar space, the telescope would transfer to a Sun-Earth L2 halo orbit and obtain high sensitivity, high resolution, high contrast UVOIR observations that address the scientific objectives of the Habitable-Exoplanet Imaging Missions.

Absolute orbit determination using line-of-sight vector measurements between formation flying spacecraft

Science.gov (United States)

Ou, Yangwei; Zhang, Hongbo; Li, Bin

2018-04-01

The purpose of this paper is to show that absolute orbit determination can be achieved based on spacecraft formation. The relative position vectors expressed in the inertial frame are used as measurements. In this scheme, the optical camera is applied to measure the relative line-of-sight (LOS) angles, i.e., the azimuth and elevation. The LIDAR (Light radio Detecting And Ranging) or radar is used to measure the range and we assume that high-accuracy inertial attitude is available. When more deputies are included in the formation, the formation configuration is optimized from the perspective of the Fisher information theory. Considering the limitation on the field of view (FOV) of cameras, the visibility of spacecraft and the installation of cameras are investigated. In simulations, an extended Kalman filter (EKF) is used to estimate the position and velocity. The results show that the navigation accuracy can be enhanced by using more deputies and the installation of cameras significantly affects the navigation performance.

Searching sequences of resonant orbits between a spacecraft and Jupiter

International Nuclear Information System (INIS)

Formiga, J K S; Prado, A F B A

2013-01-01

This research shows a study of the dynamical behavior of a spacecraft that performs a series of close approaches with the planet Jupiter. The main idea is to find a sequence of resonant orbits that allows the spacecraft to stay in the region of the space near the orbit of Jupiter around the Sun gaining energy from each passage by the planet. The dynamical model considers the existence of only two massive bodies in the systems, which are the Sun and Jupiter. They are assumed to be in circular orbits around their center of mass. Analytical equations are used to obtain the values of the parameters required to get this sequence of close approaches. Those equations are useful, because they show which orbits are physically possible when taking into account that the periapsis distances have to be above the surface of the Sun and that the closest approach distances during the passage by Jupiter have to be above its surface

Positron-attachment to small molecules: Vibrational enhancement of positron affinities with configuration interaction level of multi-component molecular orbital approach

Energy Technology Data Exchange (ETDEWEB)

Tachikawa, Masanori [Quantum Chemistry Division, Graduate School of NanoBioScience, Yokohama City University, 22-2 Seto, Kanazawa, Yokohama 236-0027 (Japan)

2015-12-31

To theoretically demonstrate the binding of a positron to small polarized molecules, we have calculated the vibrational averaged positron affinity (PA) values along the local vibrational contribution with the configuration interaction level of multi-component molecular orbital method. This method can take the electron-positron correlation contribution into account through single electronic - single positronic excitation configurations. The PA values are enhanced by including the local vibrational contribution from vertical PA values due to the anharmonicity of the potential.

On-orbit supervisor for controlling spacecraft

Science.gov (United States)

Vandervoort, Richard J.

1992-07-01

Spacecraft systems of the 1990's and beyond will be substantially more complex than their predecessors. They will have demanding performance requirements and will be expected to operate more autonomously. This underscores the need for innovative approaches to Fault Detection, Isolation and Recovery (FDIR). A hierarchical expert system is presented that provides on-orbit supervision using intelligent FDIR techniques. Each expert system in the hierarchy supervises the operation of a local set of spacecraft functions. Spacecraft operational goals flow top down while responses flow bottom up. The expert system supervisors have a fairly high degree of autonomy. Bureaucratic responsibilities are minimized to conserve bandwidth and maximize response time. Data for FDIR can be acquired local to an expert and from other experts. By using a blackboard architecture for each supervisor, the system provides a great degree of flexibility in implementing the problem solvers for each problem domain. In addition, it provides for a clear separation between facts and knowledge, leading to an efficient system capable of real time response.

Momentum accumulation due to solar radiation torque, and reaction wheel sizing, with configuration optimization

Science.gov (United States)

Hablani, Hari B.

1993-01-01

This paper has a two-fold objective: determination of yearly momentum accumulation due to solar radiation pressure, and optimum reaction wheel sizing. The first objective is confronted while determining propellant consumption by the attitude control system over a spacecraft's lifetime. This, however, cannot be obtained from the daily momentum accumulation and treating that constant throughout the year, because the orientation of the solar arrays relative to the spacecraft changes over a wide range in a year, particularly if the spacecraft has two arrays, one normal and the other off-normal to different extent at different times to the sun rays. The paper first develops commands for the arrays for tracking the sun, the arrays articulated to earth-pointing spacecraft with two rotational degrees of freedom, and spacecraft in an arbitrary circular orbit. After developing expressions for solar radiation torque due to one or both arrays, arranged symmetrically or asymmetrically relative to the spacecraft bus, momentum accumulation over an orbit and then over a year are determined. The remainder of the paper is concerned with designing reaction wheel configurations. Four-, six-, and three-wheel configurations are considered, and for given torque and momentum requirements, their cant angles with the roll/yaw plane are optimized for minimum power consumption. Finally, their momentum and torque capacities are determined for one-wheel failure scenario, and six configurations are compared and contrasted.

A Collaborative Analysis Tool for Thermal Protection Systems for Single Stage to Orbit Launch Vehicles

Science.gov (United States)

Alexander, Reginald; Stanley, Thomas Troy

2001-01-01

Presented is a design tool and process that connects several disciplines which are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to all other systems, as is the case with SSTO vehicles with air breathing propulsion, which is currently being studied by the National Aeronautics and Space Administration (NASA). In particular, the thermal protection system (TPS) is linked directly to almost every major system. The propulsion system pushes the vehicle to velocities on the order of 15 times the speed of sound in the atmosphere before pulling up to go to orbit which results in high temperatures on the external surfaces of the vehicle. Thermal protection systems to maintain the structural integrity of the vehicle must be able to mitigate the heat transfer to the structure and be lightweight. Herein lies the interdependency, in that as the vehicle's speed increases, the TPS requirements are increased. And as TPS masses increase the effect on the propulsion system and all other systems is compounded. To adequately calculate the TPS mass of this type of vehicle several engineering disciplines and analytical tools must be used preferably in an environment that data is easily transferred and multiple iterations are easily facilitated.

A Rocket Powered Single-Stage-to-Orbit Launch Vehicle With U.S. and Soviet Engineers

Science.gov (United States)

MacConochie, Ian O.; Stnaley, Douglas O.

1991-01-01

A single-stage-to-orbit launch vehicle is used to assess the applicability of Soviet Energia high-pressure-hydrocarbon engine to advanced U.S. manned space transportation systems. Two of the Soviet engines are used with three Space Shuttle Main Engines. When applied to a baseline vehicle that utilized advanced hydrocarbon engines, the higher weight of the Soviet engines resulted in a 20 percent loss of payload capability and necessitated a change in the crew compartment size and location from mid-body to forebody in order to balance the vehicle. Various combinations of Soviet and Shuttle engines were evaluated for comparison purposes, including an all hydrogen system using all Space Shuttle Main Engines. Operational aspects of the baseline vehicle are also discussed. A new mass properties program entitles Weights and Moments of Inertia (WAMI) is used in the study.

Simulated Aging of Spacecraft External Materials on Orbit

Science.gov (United States)

Khatipov, S.

Moscow State Engineering Physics Institute (MIFI), in cooperation with Air Force Research Laboratory's Satellite Assessment Center (SatAC), the European Office of Aerospace Research and Development (EOARD), and the International Science and Technology Center (ISTC), has developed a database describing the changes in optical properties of materials used on the external surfaces of spacecraft due to space environmental factors. The database includes data acquired from tests completed under contract with the ISTC and EOARD, as well as from previous Russian materials studies conducted within the last 30 years. The space environmental factors studied are for those found in Low Earth Orbits (LEO) and Geosynchronous orbits (GEO), including electron irradiation at 50, 100, and 200 keV, proton irradiation at 50, 150, 300, and 500 keV, and ultraviolet irradiation equivalent to 1 sun-year. The material characteristics investigated were solar absorption (aS), spectral reflectance (rl), solar reflectance (rS), emissivity (e), spectral transmission coefficient (Tl), solar transmittance (TS), optical density (D), relative optical density (D/x), Bi-directional Reflectance Distribution Function (BRDF), and change of appearance and color in the visible wavelengths. The materials tested in the project were thermal control coatings (paints), multilayer insulation (films), and solar cells. The ability to predict changes in optical properties of spacecraft materials is important to increase the fidelity of space observation tools, better understand observation of space objects, and increase the longevity of spacecraft. The end goal of our project is to build semi-empirical mathematical models to predict the long-term effects of space aging as a function of time and orbit.

Proposed gravity-gradient dynamics experiments in lunar orbit using the RAE-B spacecraft

Science.gov (United States)

Blanchard, D. L.; Walden, H.

1973-01-01

A series of seven gravity-gradient dynamics experiments is proposed utilizing the Radio Astronomy Explorer (RAE-B) spacecraft in lunar orbit. It is believed that none of the experiments will impair the spacecraft structure or adversely affect the continuation of the scientific mission of the satellite. The first experiment is designed to investigate the spacecraft dynamical behavior in the absence of libration damper action and inertia. It requires stable gravity-gradient capture of the spacecraft in lunar orbit with small amplitude attitude librations as a prerequisite. Four subsequent experiments involve partial retraction, ultimately followed by full redeployment, of one or two of the 230-meter booms forming the lunar-directed Vee-antenna. These boom length change operations will induce moderate amplitude angular librations of the spacecraft.

On-Orbit 3-Dimensional Electrostatic Detumble for Generic Spacecraft Geometries

Science.gov (United States)

Bennett, Trevor J.

In recent years, there is a growing interest in active debris removal and on-orbit servicing of Earth orbiting assets. The growing need for such approaches is often exemplified by the Iridium-Kosmos collision in 2009 that generated thousands of debris fragments. There exists a variety of active debris removal and on-orbit servicing technologies in development. Conventional docking mechanisms and mechanical capture by actuated manipulators, exemplified by NASA's Restore-L mission, require slow target tumble rates or more aggressive circumnavigation rate matching. The tumble rate limitations can be overcome with flexible capture systems such nets, harpoons, or tethers yet these systems require complex deployment, towing, and/or interfacing strategies to avoid servicer and target damage. Alternatively, touchless methods overcome the tumble rate limitations by provide detumble control prior to a mechanical interface. This thesis explores electrostatic detumble technology to touchlessly reduce large target rotation rates of Geostationary satellites and debris. The technical challenges preceding flight implementation largely reside in the long-duration formation flying guidance, navigation, and control of a servicer spacecraft equipped with electrostatic charge transfer capability. Leveraging prior research into the electrostatic charging of spacecraft, electrostatic detumble control formulations are developed for both axisymmetric and generic target geometries. A novel relative position vector and associated relative orbit control approach is created to manage the long-duration proximity operations. Through detailed numerical simulations, the proposed detumble and relative motion control formulations demonstrate detumble of several thousand kilogram spacecraft tumbling at several degrees per second in only several days. The availability, either through modeling or sensing, of the relative attitude, relative position, and electrostatic potential are among key concerns

Suitable configurations for triangular formation flying about collinear libration points under the circular and elliptic restricted three-body problems

Science.gov (United States)

Ferrari, Fabio; Lavagna, Michèle

2018-06-01

The design of formations of spacecraft in a three-body environment represents one of the most promising challenges for future space missions. Two or more cooperating spacecraft can greatly answer some very complex mission goals, not achievable by a single spacecraft. The dynamical properties of a low acceleration environment such as the vicinity of libration points associated to a three-body system, can be effectively exploited to design spacecraft configurations able of satisfying tight relative position and velocity requirements. This work studies the evolution of an uncontrolled formation orbiting in the proximity of periodic orbits about collinear libration points under the Circular and Elliptic Restricted Three-Body Problems. A three spacecraft triangularly-shaped formation is assumed as a representative geometry to be investigated. The study identifies initial configurations that provide good performance in terms of formation keeping, and investigates key parameters that control the relative dynamics between the spacecraft within the three-body system. Formation keeping performance is quantified by monitoring shape and size changes of the triangular formation. The analysis has been performed under five degrees of freedom to define the geometry, the orientation and the location of the triangle in the synodic rotating frame.

Space Transportation System Cargo projects: inertial stage/spacecraft integration plan. Volume 1: Management plan

Science.gov (United States)

1981-01-01

The Kennedy Space Center (KSC) Management System for the Inertial Upper Stage (IUS) - spacecraft processing from KSC arrival through launch is described. The roles and responsibilities of the agencies and test team organizations involved in IUS-S/C processing at KSC for non-Department of Defense missions are described. Working relationships are defined with respect to documentation preparation, coordination and approval, schedule development and maintenance, test conduct and control, configuration management, quality control and safety. The policy regarding the use of spacecraft contractor test procedures, IUS contractor detailed operating procedures and KSC operations and maintenance instructions is defined. Review and approval requirements for each documentation system are described.

A Sampling Based Approach to Spacecraft Autonomous Maneuvering with Safety Specifications

Science.gov (United States)

Starek, Joseph A.; Barbee, Brent W.; Pavone, Marco

2015-01-01

This paper presents a methods for safe spacecraft autonomous maneuvering that leverages robotic motion-planning techniques to spacecraft control. Specifically the scenario we consider is an in-plan rendezvous of a chaser spacecraft in proximity to a target spacecraft at the origin of the Clohessy Wiltshire Hill frame. The trajectory for the chaser spacecraft is generated in a receding horizon fashion by executing a sampling based robotic motion planning algorithm name Fast Marching Trees (FMT) which efficiently grows a tree of trajectories over a set of probabillistically drawn samples in the state space. To enforce safety the tree is only grown over actively safe samples for which there exists a one-burn collision avoidance maneuver that circularizes the spacecraft orbit along a collision-free coasting arc and that can be executed under potential thrusters failures. The overall approach establishes a provably correct framework for the systematic encoding of safety specifications into the spacecraft trajectory generations process and appears amenable to real time implementation on orbit. Simulation results are presented for a two-fault tolerant spacecraft during autonomous approach to a single client in Low Earth Orbit.

SSTO RLVs: More Global Reach? A Study of the Use of Single Stage to Orbit Reusable Launch Vehicles as Airlift Platforms.

Science.gov (United States)

1996-11-01

Orbit ( SSTO ) Reusable Launch Vehicles (RLVs) are currently under cooperative development by NASA, the Air Force, and the aerospace industry in the pursuit...exploit these rapid transit technologies to advance ’Global Reach for America.’ The SSTO RLV is a single stage rocket that will be completely reusable...investigated to assess the projected capabilities and costs of the SSTO system. This paper reviews the proposed capabilities of the SSTO system, discusses

A Study of Single- and Double-Averaged Second-Order Models to Evaluate Third-Body Perturbation Considering Elliptic Orbits for the Perturbing Body

Directory of Open Access Journals (Sweden)

R. C. Domingos

2013-01-01

Full Text Available The equations for the variations of the Keplerian elements of the orbit of a spacecraft perturbed by a third body are developed using a single average over the motion of the spacecraft, considering an elliptic orbit for the disturbing body. A comparison is made between this approach and the more used double averaged technique, as well as with the full elliptic restricted three-body problem. The disturbing function is expanded in Legendre polynomials up to the second order in both cases. The equations of motion are obtained from the planetary equations, and several numerical simulations are made to show the evolution of the orbit of the spacecraft. Some characteristics known from the circular perturbing body are studied: circular, elliptic equatorial, and frozen orbits. Different initial eccentricities for the perturbed body are considered, since the effect of this variable is one of the goals of the present study. The results show the impact of this parameter as well as the differences between both models compared to the full elliptic restricted three-body problem. Regions below, near, and above the critical angle of the third-body perturbation are considered, as well as different altitudes for the orbit of the spacecraft.

BepiColombo — The Next Step of Mercury Exploration with Two Orbiting Spacecraft

Science.gov (United States)

Benkhoff, J.

2018-05-01

BepiColombo is a joint project between ESA and JAXA. The mission consists of two orbiters — the Mercury Planetary Orbiter and the Mercury Magnetospheric Orbiter. From dedicated orbits, the spacecraft will be studying the planet and its environment.

Simulated orbits of heavy planetary ions at Mars for different IMF configurations

Science.gov (United States)

Curry, Shannon; Luhmann, Janet; Livi, Roberto; Hara, Takuya; Dong, Chuanfei; Ma, Yingjuan; McFadden, James; Bougher, Stephen

2014-11-01

We present simulated detections of O+, O2+ and CO2+ ions at Mars along a virtual orbit in the Mars space environment. Planetary pick-up ions are formed through the direct interaction of the solar wind with the neutral upper atmosphere, causing the newly created ions to be picked up and accelerated by the background convective electric field. Because previous missions such as Mars Global Surveyor (MGS) and Mars Express (MEX) have not been able to measure the interplanetary magnetic field (IMF) components simultaneously with plasma measurements, the response of heavy planetary pick-up ions to changes in the IMF has not been well characterized. Using a steady-state multi-species MHD model to provide the background electric and magnetic fields, the Mars Test Particle (MTP) simulation can trace each of these particles along field lines in near-Mars space and construct virtual ion detections from a spacecraft orbit. Specifically, we will present energy-time spectrograms and velocity space distributions (VSDs) for a selection of orbits during different IMF configurations and solar cycle conditions. These simulated orbits have broader implications for how to measure ion escape. Using individual particle traces, the origin and trajectories of different ion populations can be analyzed in order to assess how and where they contribute to the total atmospheric escape rate, which is a major objective of the upcoming MAVEN mission.

Proto-jet configurations in RADs orbiting a Kerr SMBH: symmetries and limiting surfaces

Science.gov (United States)

Pugliese, D.; Stuchlík, Z.

2018-05-01

Ringed accretion disks (RADs) are agglomerations of perfect-fluid tori orbiting around a single central attractor that could arise during complex matter inflows in active galactic nuclei. We focus our analysis to axi-symmetric accretion tori orbiting in the equatorial plane of a supermassive Kerr black hole; equilibrium configurations, possible instabilities, and evolutionary sequences of RADs were discussed in our previous works. In the present work we discuss special instabilities related to open equipotential surfaces governing the material funnels emerging at various regions of the RADs, being located between two or more individual toroidal configurations of the agglomerate. These open structures could be associated to proto-jets. Boundary limiting surfaces are highlighted, connecting the emergency of the jet-like instabilities with the black hole dimensionless spin. These instabilities are observationally significant for active galactic nuclei, being related to outflows of matter in jets emerging from more than one torus of RADs orbiting around supermassive black holes.

Optimizing Orbit-Instrument Configuration for Global Precipitation Mission (GPM) Satellite Fleet

Science.gov (United States)

Smith, Eric A.; Adams, James; Baptista, Pedro; Haddad, Ziad; Iguchi, Toshio; Im, Eastwood; Kummerow, Christian; Einaudi, Franco (Technical Monitor)

2001-01-01

Following the scientific success of the Tropical Rainfall Measuring Mission (TRMM) spearheaded by a group of NASA and NASDA scientists, their external scientific collaborators, and additional investigators within the European Union's TRMM Research Program (EUROTRMM), there has been substantial progress towards the development of a new internationally organized, global scale, and satellite-based precipitation measuring mission. The highlights of this newly developing mission are a greatly expanded scope of measuring capability and a more diversified set of science objectives. The mission is called the Global Precipitation Mission (GPM). Notionally, GPM will be a constellation-type mission involving a fleet of nine satellites. In this fleet, one member is referred to as the "core" spacecraft flown in an approximately 70 degree inclined non-sun-synchronous orbit, somewhat similar to TRMM in that it carries both a multi-channel polarized passive microwave radiometer (PMW) and a radar system, but in this case it will be a dual frequency Ku-Ka band radar system enabling explicit measurements of microphysical DSD properties. The remainder of fleet members are eight orbit-synchronized, sun-synchronous "constellation" spacecraft each carrying some type of multi-channel PMW radiometer, enabling no worse than 3-hour diurnal sampling over the entire globe. In this configuration the "core" spacecraft serves as a high quality reference platform for training and calibrating the PMW rain retrieval algorithms used with the "constellation" radiometers. Within NASA, GPM has advanced to the pre-formulation phase which has enabled the initiation of a set of science and technology studies which will help lead to the final mission design some time in the 2003 period. This presentation first provides an overview of the notional GPM program and mission design, including its organizational and programmatic concepts, scientific agenda, expected instrument package, and basic flight

Particle-in-Cell Simulation Study on the Floating Potential of Spacecraft in the Low Earth Orbit

International Nuclear Information System (INIS)

Tang Daotan; Yang Shengsheng; Zheng Kuohai; Qin Xiaogang; Li Detian; Liu Qing; Zhao Chengxuan; Du Shanshan

2015-01-01

In order to further understand the characteristics of the floating potential of low earth orbit spacecraft, the effects of the electron current collection area, background electron temperature, photocurrent emission, spacecraft wake, and the shape of spacecraft on spacecraft floating potential were studied here by particle-in-cell simulation in the low earth orbit. The simulation results show that the electron current collection area and background electron temperature impact on the floating potential by changing the electron current collection of spacecraft. By increasing the electron current collection area or background electron temperature, the spacecraft will float at a lower electric potential with respect to the surrounding plasma. However, the spacecraft wake affects the floating potential by increasing the ion current collected by spacecraft. The emission of the photocurrent from the spacecraft surface, which compensates for the electrons collected from background plasma, causes the floating potential to increase. The shape of the spacecraft is also an important factor influencing the floating potential. (paper)

Combined spacecraft orbit and attitude control through extended Kalman filtering of magnetometer, gyro, and GPS measurements

Directory of Open Access Journals (Sweden)

Tamer Mekky Ahmed Habib

2014-06-01

Full Text Available The main goal of this research is to establish spacecraft orbit and attitude control algorithms based on extended Kalman filter which provides estimates of spacecraft orbital and attitude states. The control and estimation algorithms must be capable of dealing with the spacecraft conditions during the detumbling and attitude acquisition modes of operation. These conditions are characterized by nonlinearities represented by large initial attitude angles, large initial angular velocities, large initial attitude estimation error, and large initial position estimation error. All of the developed estimation and control algorithms are suitable for application to the next Egyptian scientific satellite, EGYPTSAT-2. The parameters of the case-study spacecraft are similar but not identical to the former Egyptian satellite EGYPTSAT-1. This is done because the parameters of EGYPTSAT-2 satellite have not been consolidated yet. The sensors utilized are gyro, magnetometer, and GPS. Gyro and magnetometer are utilized to provide measurements for the estimates of spacecraft attitude state vector where as magnetometer and GPS are utilized to provide measurements for the estimates of spacecraft orbital state vector.

An Artificial Gravity Spacecraft Approach which Minimizes Mass, Fuel and Orbital Assembly Reg

Science.gov (United States)

Bell, L.

2002-01-01

The Sasakawa International Center for Space Architecture (SICSA) is undertaking a multi-year research and design study that is exploring near and long-term commercial space development opportunities. Space tourism in low-Earth orbit (LEO), and possibly beyond LEO, comprises one business element of this plan. Supported by a financial gift from the owner of a national U.S. hotel chain, SICSA has examined opportunities, requirements and facility concepts to accommodate up to 100 private citizens and crewmembers in LEO, as well as on lunar/planetary rendezvous voyages. SICSA's artificial gravity Science Excursion Vehicle ("AGSEV") design which is featured in this presentation was conceived as an option for consideration to enable round-trip travel to Moon and Mars orbits and back from LEO. During the course of its development, the AGSEV would also serve other important purposes. An early assembly stage would provide an orbital science and technology testbed for artificial gravity demonstration experiments. An ultimate mature stage application would carry crews of up to 12 people on Mars rendezvous missions, consuming approximately the same propellant mass required for lunar excursions. Since artificial gravity spacecraft that rotate to create centripetal accelerations must have long spin radii to limit adverse effects of Coriolis forces upon inhabitants, SICSA's AGSEV design embodies a unique tethered body concept which is highly efficient in terms of structural mass and on-orbit assembly requirements. The design also incorporates "inflatable" as well as "hard" habitat modules to optimize internal volume/mass relationships. Other important considerations and features include: maximizing safety through element and system redundancy; means to avoid destabilizing mass imbalances throughout all construction and operational stages; optimizing ease of on-orbit servicing between missions; and maximizing comfort and performance through careful attention to human needs. A

Operational factors affecting microgravity levels in orbit

Science.gov (United States)

Olsen, R. E.; Mockovciak, J., Jr.

1980-01-01

Microgravity levels desired for proposed materials processing payloads are fundamental considerations in the design of future space platforms. Disturbance sources, such as aerodynamic drag, attitude control torques, crew motion and orbital dynamics, influence the microgravity levels attainable in orbit. The nature of these effects are assessed relative to platform design parameters such as orbital altitude and configuration geometry, and examples are presented for a representative spacecraft configuration. The possible applications of control techniques to provide extremely low acceleration levels are also discussed.

Contingency Trajectory Design for a Lunar Orbit Insertion Maneuver Failure by the LADEE Spacecraft

Science.gov (United States)

Genova, A. L.

2014-01-01

This paper presents results from a contingency trajectory analysis performed for the Lunar Atmosphere & Dust Environment Explorer (LADEE) mission in the event of a missed lunar-orbit insertion (LOI) maneuver by the LADEE spacecraft. The effects of varying solar perturbations in the vicinity of the weak stability boundary (WSB) in the Sun-Earth system on the trajectory design are analyzed and discussed. It is shown that geocentric recovery trajectory options existed for the LADEE spacecraft, depending on the spacecraft's recovery time to perform an Earth escape-prevention maneuver after the hypothetical LOI maneuver failure and subsequent path traveled through the Sun-Earth WSB. If Earth-escape occurred, a heliocentric recovery option existed, but with reduced science capacapability for the spacecraft in an eccentric, not circular near-equatorial retrograde lunar orbit.

ADRC for spacecraft attitude and position synchronization in libration point orbits

Science.gov (United States)

Gao, Chen; Yuan, Jianping; Zhao, Yakun

2018-04-01

This paper addresses the problem of spacecraft attitude and position synchronization in libration point orbits between a leader and a follower. Using dual quaternion, the dimensionless relative coupled dynamical model is derived considering computation efficiency and accuracy. Then a model-independent dimensionless cascade pose-feedback active disturbance rejection controller is designed to spacecraft attitude and position tracking control problems considering parameter uncertainties and external disturbances. Numerical simulations for the final approach phase in spacecraft rendezvous and docking and formation flying are done, and the results show high-precision tracking errors and satisfactory convergent rates under bounded control torque and force which validate the proposed approach.

Model predictive control for spacecraft rendezvous in elliptical orbit

Science.gov (United States)

Li, Peng; Zhu, Zheng H.

2018-05-01

This paper studies the control of spacecraft rendezvous with attitude stable or spinning targets in an elliptical orbit. The linearized Tschauner-Hempel equation is used to describe the motion of spacecraft and the problem is formulated by model predictive control. The control objective is to maximize control accuracy and smoothness simultaneously to avoid unexpected change or overshoot of trajectory for safe rendezvous. It is achieved by minimizing the weighted summations of control errors and increments. The effects of two sets of horizons (control and predictive horizons) in the model predictive control are examined in terms of fuel consumption, rendezvous time and computational effort. The numerical results show the proposed control strategy is effective.

Analysis of payload bay magnetic fields due to dc power multipoint and single point ground configurations

Science.gov (United States)

Lawton, R. M.

1976-01-01

An analysis of magnetic fields in the Orbiter Payload Bay resulting from the present grounding configuration (structure return) was presented and the amount of improvement that would result from installing wire returns for the three dc power buses was determined. Ac and dc magnetic fields at five points in a cross-section of the bay are calculated for both grounding configurations. Y and Z components of the field at each point are derived in terms of a constant coefficient and the current amplitude of each bus. The dc loads assumed are 100 Amperes for each bus. The ac noise current used is a spectrum 6 db higher than the Orbiter equipment limit for narrowband conducted emissions. It was concluded that installing return wiring to provide a single point ground for the dc Buses in the Payload Bay would reduce the ac and dc magnetic field intensity by approximately 30 db.

Four-spacecraft determination of magnetopause orientation, motion and thickness: comparison with results from single-spacecraft methods

Directory of Open Access Journals (Sweden)

S. E. Haaland

2004-04-01

Full Text Available In this paper, we use Cluster data from one magnetopause event on 5 July 2001 to compare predictions from various methods for determination of the velocity, orientation, and thickness of the magnetopause current layer. We employ established as well as new multi-spacecraft techniques, in which time differences between the crossings by the four spacecraft, along with the duration of each crossing, are used to calculate magnetopause speed, normal vector, and width. The timing is based on data from either the Cluster Magnetic Field Experiment (FGM or the Electric Field Experiment (EFW instruments. The multi-spacecraft results are compared with those derived from various single-spacecraft techniques, including minimum-variance analysis of the magnetic field and deHoffmann-Teller, as well as Minimum-Faraday-Residue analysis of plasma velocities and magnetic fields measured during the crossings. In order to improve the overall consistency between multi- and single-spacecraft results, we have also explored the use of hybrid techniques, in which timing information from the four spacecraft is combined with certain limited results from single-spacecraft methods, the remaining results being left for consistency checks. The results show good agreement between magnetopause orientations derived from appropriately chosen single-spacecraft techniques and those obtained from multi-spacecraft timing. The agreement between magnetopause speeds derived from single- and multi-spacecraft methods is quantitatively somewhat less good but it is evident that the speed can change substantially from one crossing to the next within an event. The magnetopause thickness varied substantially from one crossing to the next, within an event. It ranged from 5 to 10 ion gyroradii. The density profile was sharper than the magnetic profile: most of the density change occured in the earthward half of the magnetopause.

Key words. Magnetospheric physics (magnetopause, cusp and

Construction and experimental verification of a novel flexible thermal control system configuration for the autonomous on-orbit services of space missions

International Nuclear Information System (INIS)

Guo, Wei; Li, Yunhua; Li, Yun-Ze; Wang, Sheng-Nan; Zhong, Ming-Liang; Wang, Ji-Xiang; Zhang, Jia-Xun

2017-01-01

Highlights: • A novel flexible thermal control system (F-TCS) for spacecraft on-orbit services is proposed. • Systemic construction and reconfiguration operating mechanism are highlighted. • Thermal-hydraulic performances of the F-TCS are investigated both numerically and experimentally. • The F-TCS has the vast potential for spacecrafts autonomous thermal management. - Abstract: This paper proposed a novel flexible thermal control system (F-TCS) configuration for realizing thermal management for spacecrafts autonomous on-orbit service (A-OOS) demands. With a dual-ring topology which composes of a heat collecting bus, a heat dissipating bus, connection brunches and inter-platform service interfaces, the F-TCS may realize not only self-reconfiguration operations but also providing heat dissipation resources for other spacecrafts or cabins. The F-TCS hydraulic and thermal dynamics were modeled, a verification testbed was also established to validate the F-TCS thermal control performance. Focused on investigating the self-reconfiguration and thermal control cooperative operations, several typical A-OOS cases were imposed on the F-TCS, numerical simulations and experimental validations were respectively implemented. Both results demonstrated that the meticulously designed F-TCS is capable of offering self-topological reconfiguration with fast time response and robust temperature control performances, high systemic heat transfer efficiency is also recommended from the point of view of energy saving. The F-TCS is suggested as a promising solution for A-OOS owing to its higher reliability and promising autonomous maintenance potential which is suitable for future spacecrafts thermal management requirements.

In-orbit evaluation of the control system/structural mode interactions of the OSO-8 spacecraft

Science.gov (United States)

Slafer, L. I.

1979-01-01

The Orbiting Solar Observatory-8 experienced severe structural mode/control loop interaction problems during the spacecraft development. Extensive analytical studies, using the hybrid coordinate modeling approach, and comprehensive ground testing were carried out in order to achieve the system's precision pointing performance requirements. A recent series of flight tests were conducted with the spacecraft in which a wide bandwidth, high resolution telemetry system was utilized to evaluate the on-orbit flexible dynamics characteristics of the vehicle along with the control system performance. The paper describes the results of these tests, reviewing the basic design problem, analytical approach taken, ground test philosophy, and on-orbit testing. Data from the tests was used to determine the primary mode frequency, damping, and servo coupling dynamics for the on-orbit condition. Additionally, the test results have verified analytically predicted differences between the on-orbit and ground test environments, and have led to a validation of both the analytical modeling and servo design techniques used during the development of the control system.

Application of Space Environmental Observations to Spacecraft Pre-Launch Engineering and Spacecraft Operations

Science.gov (United States)

Barth, Janet L.; Xapsos, Michael

2008-01-01

This presentation focuses on the effects of the space environment on spacecraft systems and applying this knowledge to spacecraft pre-launch engineering and operations. Particle radiation, neutral gas particles, ultraviolet and x-rays, as well as micrometeoroids and orbital debris in the space environment have various effects on spacecraft systems, including degradation of microelectronic and optical components, physical damage, orbital decay, biasing of instrument readings, and system shutdowns. Space climate and weather must be considered during the mission life cycle (mission concept, mission planning, systems design, and launch and operations) to minimize and manage risk to both the spacecraft and its systems. A space environment model for use in the mission life cycle is presented.

Micrometeoroid and Orbital Debris (MMOD) Shield Ballistic Limit Analysis Program

Science.gov (United States)

Ryan, Shannon

2013-01-01

This software implements penetration limit equations for common micrometeoroid and orbital debris (MMOD) shield configurations, windows, and thermal protection systems. Allowable MMOD risk is formulated in terms of the probability of penetration (PNP) of the spacecraft pressure hull. For calculating the risk, spacecraft geometry models, mission profiles, debris environment models, and penetration limit equations for installed shielding configurations are required. Risk assessment software such as NASA's BUMPERII is used to calculate mission PNP; however, they are unsuitable for use in shield design and preliminary analysis studies. The software defines a single equation for the design and performance evaluation of common MMOD shielding configurations, windows, and thermal protection systems, along with a description of their validity range and guidelines for their application. Recommendations are based on preliminary reviews of fundamental assumptions, and accuracy in predicting experimental impact test results. The software is programmed in Visual Basic for Applications for installation as a simple add-in for Microsoft Excel. The user is directed to a graphical user interface (GUI) that requires user inputs and provides solutions directly in Microsoft Excel workbooks.

Particle orbits in W VII-X configurations

International Nuclear Information System (INIS)

Wobig, H.

1987-01-01

It is shown that magnetic coordinates are very convenient for describing particle orbits and neoclassical losses in stellarator studies. In the configurations considered (WVII-A and AS, Helias, Heliac, and BSX) plateau losses can be reduced by a factor of 3-4 compared with an equivalent tokamak. It is possible to reduce plateau losses and bootstrap current simultaneously. The bootstrap current can be made negligibly small. The Heliac configuration shows larger plateau losses than the equivalent tokamak. It is possible to localize trapped particles in a region of minimum radial drift velocity. The Er fields strongly reduce localized particle losses

Chaos and its control in the pitch motion of an asymmetric magnetic spacecraft in polar elliptic orbit

Energy Technology Data Exchange (ETDEWEB)

Inarrea, Manuel [Universidad de La Rioja, Area de Fisica Aplicada, 26006 Logrono (Spain)], E-mail: manuel.inarrea@unirioja.es

2009-05-30

We study the pitch attitude dynamics of an asymmetric magnetic spacecraft in a polar almost circular orbit under the influence of a gravity gradient torque. The spacecraft is perturbed by the small eccentricity of the elliptic orbit and by a small magnetic torque generated by the interaction between the Earth's magnetic field and the magnetic moment of the spacecraft. Under both perturbations, we show that the pitch motion exhibits heteroclinic chaotic behavior by means of the Melnikov method. Numerical methods applied to simulations of the pitch motion also confirm the chaotic character of the spacecraft attitude dynamics. Finally, a linear time-delay feedback method for controlling chaos is applied to the governing equations of the spacecraft pitch motion in order to remove the chaotic character of initially irregular attitude motions and transform them into periodic ones.

Chaos and its control in the pitch motion of an asymmetric magnetic spacecraft in polar elliptic orbit

International Nuclear Information System (INIS)

Inarrea, Manuel

2009-01-01

We study the pitch attitude dynamics of an asymmetric magnetic spacecraft in a polar almost circular orbit under the influence of a gravity gradient torque. The spacecraft is perturbed by the small eccentricity of the elliptic orbit and by a small magnetic torque generated by the interaction between the Earth's magnetic field and the magnetic moment of the spacecraft. Under both perturbations, we show that the pitch motion exhibits heteroclinic chaotic behavior by means of the Melnikov method. Numerical methods applied to simulations of the pitch motion also confirm the chaotic character of the spacecraft attitude dynamics. Finally, a linear time-delay feedback method for controlling chaos is applied to the governing equations of the spacecraft pitch motion in order to remove the chaotic character of initially irregular attitude motions and transform them into periodic ones.

Laboratory investigations: Low Earth orbit environment chemistry with spacecraft surfaces

Science.gov (United States)

Cross, Jon B.

1990-01-01

Long-term space operations that require exposure of material to the low earth orbit (LEO) environment must take into account the effects of this highly oxidative atmosphere on material properties and the possible contamination of the spacecraft surroundings. Ground-based laboratory experiments at Los Alamos using a newly developed hyperthermal atomic oxygen (AO) source have shown that not only are hydrocarbon based materials effected but that inorganic materials such as MoS2 are also oxidized and that thin protective coatings such as Al2O3 can be breached, producing oxidation of the underlying substrate material. Gas-phase reaction products, such as SO2 from oxidation of MoS2 and CO and CO2 from hydrocarbon materials, have been detected and have consequences in terms of spacecraft contamination. Energy loss through gas-surface collisions causing spacecraft drag has been measured for a few select surfaces and has been found to be highly dependent on the surface reactivity.

Searching Less Perturbed Circular Orbits for a Spacecraft Travelling around Europa

Directory of Open Access Journals (Sweden)

J. P. S. Carvalho

2014-01-01

Full Text Available Space missions to visit the natural satellite of Jupiter, Europa, constitute an important topic in space activities today, because missions to this moon are under study now. Several considerations have to be made for these missions. The present paper searches for less perturbed circular orbits around Europa. This search is made based on the total effects of the perturbing forces over the time, evaluated by the integral of those forces over the time. This value depends on the dynamical model and on the orbit of the spacecraft. The perturbing forces considered are the third-body perturbation that comes from Jupiter and the J2, J3, and C22 terms of the gravitational potential of Europa. Several numerical studies are performed and the results show the locations of the less perturbed orbits. Using those results, it is possible to find near-circular frozen orbits with smaller amplitudes of variations of the orbital elements.

On-Orbit Assembly of a Universally Interlocking Modular Spacecraft (7224-110), Phase I

Data.gov (United States)

National Aeronautics and Space Administration — Physical Sciences Inc. and Advanced Solutions, Inc. propose a novel approach for on-orbit assembly of a modular spacecraft using a unique universal, intelligent,...

Chaos and nonlinear dynamics of single-particle orbits in a magnetotaillike magnetic field

Science.gov (United States)

Chen, J.; Palmadesso, P. J.

1986-01-01

The properties of charged-particle motion in Hamiltonian dynamics are studied in a magnetotaillike magnetic field configuration. It is shown by numerical integration of the equation of motion that the system is generally nonintegrable and that the particle motion can be classified into three distinct types of orbits: bounded integrable orbits, unbounded stochastic orbits, and unbounded transient orbits. It is also shown that different regions of the phase space exhibit qualitatively different responses to external influences. The concept of 'differential memory' in single-particle distributions is proposed. Physical implications for the dynamical properties of the magnetotail plasmas and the possible generation of non-Maxwellian features in the distribution functions are discussed.

Spin-flip configuration interaction singles with exact spin-projection: Theory and applications to strongly correlated systems.

Science.gov (United States)

Tsuchimochi, Takashi

2015-10-14

Spin-flip approaches capture static correlation with the same computational scaling as the ordinary single reference methods. Here, we extend spin-flip configuration interaction singles (SFCIS) by projecting out intrinsic spin-contamination to make it spin-complete, rather than by explicitly complementing it with spin-coupled configurations. We give a general formalism of spin-projection for SFCIS, applicable to any spin states. The proposed method is viewed as a natural unification of SFCIS and spin-projected CIS to achieve a better qualitative accuracy at a low computational cost. While our wave function ansatz is more compact than previously proposed spin-complete SF approaches, it successfully offers more general static correlation beyond biradicals without sacrificing good quantum numbers. It is also shown that our method is invariant with respect to open-shell orbital rotations, due to the uniqueness of spin-projection. We will report benchmark calculations to demonstrate its qualitative performance on strongly correlated systems, including conical intersections that appear both in ground-excited and excited-excited degeneracies.

A study of Schwarz converters for nuclear powered spacecraft

Science.gov (United States)

Stuart, Thomas A.; Schwarze, Gene E.

1987-01-01

High power space systems which use low dc voltage, high current sources such as thermoelectric generators, will most likely require high voltage conversion for transmission purposes. This study considers the use of the Schwarz resonant converter for use as the basic building block to accomplish this low-to-high voltage conversion for either a dc or an ac spacecraft bus. The Schwarz converter has the important assets of both inherent fault tolerance and resonant operation; parallel operation in modular form is possible. A regulated dc spacecraft bus requires only a single stage converter while a constant frequency ac bus requires a cascaded Schwarz converter configuration. If the power system requires constant output power from the dc generator, then a second converter is required to route unneeded power to a ballast load.

A Small Spacecraft Swarm Deployment and Stationkeeping Strategy for Sun-Earth L1 Halo Orbits

Science.gov (United States)

Renea Conn, Tracie; Bookbinder, Jay

2018-01-01

Spacecraft orbits about the Sun-Earth librarian point L1 have been of interest since the 1950s. An L1 halo orbit was first achieved with the International Sun-Earth Explorer-3 (ISEE-3) mission, and similar orbits around Sun-Earth L1 were achieved in the Solar and Heliospheric Observatory (SOHO), Advanced Composition Explorer (ACE), Genesis, and Deep Space Climate Observatory (DSCOVR) missions. With recent advancements in CubeSat technology, we envision that it will soon be feasible to deploy CubeSats at L1. As opposed to these prior missions where one large satellite orbited alone, a swarm of CubeSats at L1 would enable novel science data return, providing a topology for intersatellite measurements of heliophysics phenomena both spatially and temporally, at varying spatial scales.The purpose of this iPoster is to present a flight dynamics strategy for a swarm of numerous CubeSats orbiting Sun-Earth L1. The presented method is a coupled, two-part solution. First, we present a deployment strategy for the CubeSats that is optimized to produce prescribed, time-varying intersatellite baselines for the purposes of collecting magnetometer data as well as radiometric measurements from cross-links. Second, we employ a loose control strategy that was successfully applied to SOHO and ACE for minimized stationkeeping propellant expenditure. We emphasize that the presented solution is practical within the current state-of-the-art and heritage CubeSat technology, citing capabilities of CubeSat designs that will launch on the upcoming Exploration Mission 1 (EM-1) to lunar orbits and beyond. Within this iPoster, we present animations of the simulated deployment strategy and resulting spacecraft trajectories. Mission design parameters such as total Δv required for long-term station keeping and minimum/maximum/mean spacecraft separation distances are also presented.

A Shaftless Magnetically Levitated Multifunctional Spacecraft Flywheel Storage System

Science.gov (United States)

Stevens, Ken; Thornton, Richard; Clark, Tracy; Beaman, Bob G.; Dennehy, Neil; Day, John H. (Technical Monitor)

2002-01-01

Presently many types of spacecraft use a Spacecraft Attitude Control System (ACS) with momentum wheels for steering and electrochemical batteries to provide electrical power for the eclipse period of the spacecraft orbit. Future spacecraft will use Flywheels for combined use in ACS and Energy Storage. This can be done by using multiple wheels and varying the differential speed for ACS and varying the average speed for energy storage and recovery. Technology in these areas has improved since the 1990s so it is now feasible for flywheel systems to emerge from the laboratory for spacecraft use. This paper describes a new flywheel system that can be used for both ACS and energy storage. Some of the possible advantages of a flywheel system are: lower total mass and volume, higher efficiency, less thermal impact, improved satellite integration schedule and complexity, simplified satellite orbital operations, longer life with lower risk, less pointing jitter, and greater capability for high-rate slews. In short, they have the potential to enable new types of missions and provide lower cost. Two basic types of flywheel configurations are the Flywheel Energy Storage System (FESS) and the Integrated Power and Attitude Control System (IPACS).

Towards a standardized grasping and refuelling on-orbit servicing for geo spacecraft

Science.gov (United States)

Medina, Alberto; Tomassini, Angelo; Suatoni, Matteo; Avilés, Marcos; Solway, Nick; Coxhill, Ian; Paraskevas, Iosif S.; Rekleitis, Georgios; Papadopoulos, Evangelos; Krenn, Rainer; Brito, André; Sabbatinelli, Beatrice; Wollenhaupt, Birk; Vidal, Christian; Aziz, Sarmad; Visentin, Gianfranco

2017-05-01

Exploitation of space must benefit from the latest advances in robotics. On-orbit servicing is a clear candidate for the application of autonomous rendezvous and docking mechanisms. However, during the last three decades most of the trials took place combining extravehicular activities (EVAs) with telemanipulated robotic arms. The European Space Agency (ESA) considers that grasping and refuelling are promising near-mid-term capabilities that could be performed by servicing spacecraft. Minimal add-ons on spacecraft to enhance their serviceability may protect them for a changing future in which satellite servicing may become mainstream. ESA aims to conceive and promote standard refuelling provisions that can be installed in present and future European commercial geostationary orbit (GEO) satellite platforms and scientific spacecraft. For this purpose ESA has started the ASSIST activity addressing the analysis, design and validation of internal provisions (such as modifications to fuel, gas, electrical and data architecture to allow servicing) and external provisions (such as integrated berthing fixtures with peripheral electrical, gas, liquid connectors, leak check systems and corresponding optical and radio markers for cooperative rendezvous and docking). This refuelling approach is being agreed with European industry (OHB, Thales Alenia Space) and expected to be consolidated with European commercial operators as a first step to become an international standard; this approach is also being considered for on-orbit servicing spacecraft, such as the SpaceTug, by Airbus DS. This paper describes in detail the operational means, structure, geometry and accommodation of the system. Internal and external provisions will be designed with the minimum possible impact on the current architecture of GEO satellites without introducing additional risks in the development and commissioning of the satellite. End-effector and berthing fixtures are being designed in the range of few

Space-to-Space Power Beaming Enabling High Performance Rapid Geocentric Orbit Transfer

Science.gov (United States)

Dankanich, John W.; Vassallo, Corinne; Tadge, Megan

2015-01-01

The use of electric propulsion is more prevalent than ever, with industry pursuing all electric orbit transfers. Electric propulsion provides high mass utilization through efficient propellant transfer. However, the transfer times become detrimental as the delta V transitions from near-impulsive to low-thrust. Increasing power and therefore thrust has diminishing returns as the increasing mass of the power system limits the potential acceleration of the spacecraft. By using space-to-space power beaming, the power system can be decoupled from the spacecraft and allow significantly higher spacecraft alpha (W/kg) and therefore enable significantly higher accelerations while maintaining high performance. This project assesses the efficacy of space-to-space power beaming to enable rapid orbit transfer while maintaining high mass utilization. Concept assessment requires integrated techniques for low-thrust orbit transfer steering laws, efficient large-scale rectenna systems, and satellite constellation configuration optimization. This project includes the development of an integrated tool with implementation of IPOPT, Q-Law, and power-beaming models. The results highlight the viability of the concept, limits and paths to infusion, and comparison to state-of-the-art capabilities. The results indicate the viability of power beaming for what may be the only approach for achieving the desired transit times with high specific impulse.

High-precision relative position and attitude measurement for on-orbit maintenance of spacecraft

Science.gov (United States)

Zhu, Bing; Chen, Feng; Li, Dongdong; Wang, Ying

2018-02-01

In order to realize long-term on-orbit running of satellites, space stations, etc spacecrafts, in addition to the long life design of devices, The life of the spacecraft can also be extended by the on-orbit servicing and maintenance. Therefore, it is necessary to keep precise and detailed maintenance of key components. In this paper, a high-precision relative position and attitude measurement method used in the maintenance of key components is given. This method mainly considers the design of the passive cooperative marker, light-emitting device and high resolution camera in the presence of spatial stray light and noise. By using a series of algorithms, such as background elimination, feature extraction, position and attitude calculation, and so on, the high precision relative pose parameters as the input to the control system between key operation parts and maintenance equipment are obtained. The simulation results show that the algorithm is accurate and effective, satisfying the requirements of the precision operation technique.

A computer graphics system for visualizing spacecraft in orbit

Science.gov (United States)

Eyles, Don E.

1989-01-01

To carry out unanticipated operations with resources already in space is part of the rationale for a permanently manned space station in Earth orbit. The astronauts aboard a space station will require an on-board, spatial display tool to assist the planning and rehearsal of upcoming operations. Such a tool can also help astronauts to monitor and control such operations as they occur, especially in cases where first-hand visibility is not possible. A computer graphics visualization system designed for such an application and currently implemented as part of a ground-based simulation is described. The visualization system presents to the user the spatial information available in the spacecraft's computers by drawing a dynamic picture containing the planet Earth, the Sun, a star field, and up to two spacecraft. The point of view within the picture can be controlled by the user to obtain a number of specific visualization functions. The elements of the display, the methods used to control the display's point of view, and some of the ways in which the system can be used are described.

Investigation of the charge-orbital ordering mechanism in single-layered Pr0.5Ca1.5MnO4

Science.gov (United States)

Rangkuti, C. N.; Majidi, M. A.

2018-04-01

Motivated by the experimental study of half-doped single-layered Pr0.5Ca1.5MnO4 showing charge, orbital, and spin orderings [1], we propose a model to theoretically study the system to explain such ordering phenomena. The ground state electron configuration reveals that the charges form a checkerboard pattern with alternating Mn3+/Mn4+ sites, while the orbitals are aligned in zigzag chains [1, 2]. We calculate the ground state energy of this system to find the most preferable configuration by comparing three types of configurations (charge-unordered, charge-ordered, and charge-orbital-ordered states). The calculations are based on a tight-binding model representing effective electron hoppings among Mn ions in MnO2-plane. We take into account the horizontally- and vertically-oriented orbital and spin degrees of freedom at Mn sites. We assume that the hopping integral values depend on the relative orientation between the corresponding orbitals of adjacent Mn ions. The interaction terms we incorporate into our effective Hamiltonian include inter-orbital, intra-orbital Hubbard repulsions, and Jahn-Teller distortion [2]. We absorb the exchange interaction between spins into local self-energy that we calculate within dynamical mean field algorithm [2]. Within our model we show a circumstance in which the charge-orbital ordered configuration has the lowest energy, consistent with the ground state ordering revealed by the experimental data.

Evaluation of spacecraft technology programs (effects on communication satellite business ventures), volume 1

Science.gov (United States)

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.

Science Planning and Orbit Classification for Solar Probe Plus

Science.gov (United States)

Kusterer, M. B.; Fox, N. J.; Rodgers, D. J.; Turner, F. S.

2016-12-01

There are a number of challenges for the Science Planning Team (SPT) of the Solar Probe Plus (SPP) Mission. Since SPP is using a decoupled payload operations approach, tight coordination between the mission operations and payload teams will be required. The payload teams must manage the volume of data that they write to the spacecraft solid-state recorders (SSR) for their individual instruments for downlink to the ground. Making this process more difficult, the geometry of the celestial bodies and the spacecraft during some of the SPP mission orbits cause limited uplink and downlink opportunities. The payload teams will also be required to coordinate power on opportunities, command uplink opportunities, and data transfers from instrument memory to the spacecraft SSR with the operation team. The SPT also intend to coordinate observations with other spacecraft and ground based systems. To solve these challenges, detailed orbit activity planning is required in advance for each orbit. An orbit planning process is being created to facilitate the coordination of spacecraft and payload activities for each orbit. An interactive Science Planning Tool is being designed to integrate the payload data volume and priority allocations, spacecraft ephemeris, attitude, downlink and uplink schedules, spacecraft and payload activities, and other spacecraft ephemeris. It will be used during science planning to select the instrument data priorities and data volumes that satisfy the orbit data volume constraints and power on, command uplink and data transfer time periods. To aid in the initial stages of science planning we have created an orbit classification scheme based on downlink availability and significant science events. Different types of challenges arise in the management of science data driven by orbital geometry and operational constraints, and this scheme attempts to identify the patterns that emerge.

Semi-Autonomous Telerobotic Manipulation for On-Orbit Spacecraft Servicing and Assembly over Time-Delayed Telemetry

Data.gov (United States)

National Aeronautics and Space Administration — Despite the prevalence of this telerobotic surgery of humans, we still do not have the capability to service (refuel and repair) or assemble spacecraft on-orbit with...

Jupiter's magnetosphere and aurorae observed by the Juno spacecraft during its first polar orbits

DEFF Research Database (Denmark)

Connerney, J. E. P.; Adriani, Alberto; Allegrini, F.

2017-01-01

The Juno spacecraft acquired direct observations of the jovian magnetosphere and auroral emissions from a vantage point above the poles. Juno's capture orbit spanned the jovian magnetosphere from bow shock to the planet, providing magnetic field, charged particle, and wave phenomena context...

A user's guide to the Flexible Spacecraft Dynamics and Control Program

Science.gov (United States)

Fedor, J. V.

1984-01-01

A guide to the use of the Flexible Spacecraft Dynamics Program (FSD) is presented covering input requirements, control words, orbit generation, spacecraft description and simulation options, and output definition. The program can be used in dynamics and control analysis as well as in orbit support of deployment and control of spacecraft. The program is applicable to inertially oriented spinning, Earth oriented or gravity gradient stabilized spacecraft. Internal and external environmental effects can be simulated.

Gauge-Invariant Formulation of Time-Dependent Configuration Interaction Singles Method

Directory of Open Access Journals (Sweden)

Takeshi Sato

2018-03-01

Full Text Available We propose a gauge-invariant formulation of the channel orbital-based time-dependent configuration interaction singles (TDCIS method [Phys. Rev. A, 74, 043420 (2006], one of the powerful ab initio methods to investigate electron dynamics in atoms and molecules subject to an external laser field. In the present formulation, we derive the equations of motion (EOMs in the velocity gauge using gauge-transformed time-dependent, not fixed, orbitals that are equivalent to the conventional EOMs in the length gauge using fixed orbitals. The new velocity-gauge EOMs avoid the use of the length-gauge dipole operator, which diverges at large distance, and allows us to exploit computational advantages of the velocity-gauge treatment over the length-gauge one, e.g., a faster convergence in simulations with intense and long-wavelength lasers, and the feasibility of exterior complex scaling as an absorbing boundary. The reformulated TDCIS method is applied to an exactly solvable model of one-dimensional helium atom in an intense laser field to numerically demonstrate the gauge invariance. We also discuss the consistent method for evaluating the time derivative of an observable, which is relevant, e.g., in simulating high-harmonic generation.

Habitable Mars Ascent Vehicle (MAV) Concept. [Mars Ascent Vehicle (MAV) Layout and Configuration: 6-Crew, Habitable, Nested Tank Concept

Science.gov (United States)

Dang, Victor; Rucker, Michelle

2013-01-01

NASA's ultimate goal is the human exploration of Mars. Among the many difficult aspects of a trip to Mars is the return mission that would transport the astronauts from the Martian surface back into Mars orbit. One possible conceptual design to accomplish this task is a two-stage Mars Ascent Vehicle (MAV). In order to assess this design, a general layout and configuration for the spacecraft must be developed. The objective of my internship was to model a conceptual MAV design to support NASA's latest human Mars mission architecture trade studies, technology prioritization decisions, and mass, cost, and schedule estimates.

Contingency Trajectory Design for a Lunar Orbit Insertion Maneuver Failure by the Lunar Atmosphere Dust Environment Explorer (LADEE) Spacecraft

Science.gov (United States)

Genova, Anthony L.; Loucks, Michael; Carrico, John

2014-01-01

The purpose of this extended abstract is to present results from a failed lunar-orbit insertion (LOI) maneuver contingency analysis for the Lunar Atmosphere Dust Environment Explorer (LADEE) mission, managed and operated by NASA Ames Research Center in Moffett Field, CA. The LADEE spacecrafts nominal trajectory implemented multiple sub-lunar phasing orbits centered at Earth before eventually reaching the Moon (Fig. 1) where a critical LOI maneuver was to be performed [1,2,3]. If this LOI was missed, the LADEE spacecraft would be on an Earth-escape trajectory, bound for heliocentric space. Although a partial mission recovery is possible from a heliocentric orbit (to be discussed in the full paper), it was found that an escape-prevention maneuver could be performed several days after a hypothetical LOI-miss, allowing a return to the desired science orbit around the Moon without leaving the Earths sphere-of-influence (SOI).

Nuclear electric propulsion /NEP/ spacecraft for the outer planet orbiter mission

International Nuclear Information System (INIS)

Garrison, P.W.; Nock, K.T.

1982-01-01

The design, operating features, and a possible Neptune orbit for the spacecraft powered by the SP-100 nuclear electric propulsion (NEP) system under study by NASA and the DOE are described. The system features a reactor and a payload situated on opposite ends of a 0.5 m diam, 11 m long astromast. Mercury-ion thrusters are located beneath the reactor for side thrusting, and no contamination of the payload or obstruction of the viewing angles for scientific objectives occurs with the system, which would not degrade in performance even under high insolation during near-sun maneuvers. Results of a theoretical study of earth escapes are presented to show that an NEP powered spiral trajectory out of a 700 km Shuttle orbit and using a Triton gravity assist would be superior to departing from a 300 km orbit with a Centaur boost. The mission profile includes a 1249 kg Galileo payload. The SP-100 has a 1.4 MWth reactor with UO2 fuel tiles and weighs 19,904 kg

ROGER a potential orbital space debris removal system

Science.gov (United States)

Starke, Juergen; Bischof, Bernd; Foth, W.-O.; -J., J.; Günther

The previous activities in the field of On Orbit Servicing studied in the 1990's included in partic-ular the capability of vehicles in GEO to capture and support satellites (mainly communication satellites) to enable repair and continuation of operations, and finally the controlled transfer the target into a permanent graveyard orbit. The specific capture tools for these applications were mostly based on robotic systems to capture and fix the target under specific dynamic constraints (e.g. slowly tumbling target) without damage, and to allow the stabilization, re-orientation and potential repair of the target and subsequent release or transport to the final disposal orbit. Due to the drastically increasing number of debris particularly in the Low Earth Orbits (SSO) the active debris removal is now necessary to counteract to the predicted debris production cascade (Kessler Syndrome), which means the pollution of the total sphere in low earth orbit and not only the SSO area. In most of the debris congresses it was recommended to start removal with the still integrated systems as soon as possible. In the case of large debris objects, the soft capture system can be replaced by a simpler and robust system able to operate from a safe distance to the target and flexible enough to capture and hold different types of targets such as deactivated and/or defective satellites, upper stages and big fragments. These nominally non -cooperative targets might be partially destroyed by the capture process, but the production of additional debris shall be avoided. A major argument for the commercial applications is a multi-target mission potential, which is possible at GEO because the transfer propellant requirement to the disposal orbit and the return to the orbit of the next potential target is relative low (orbits with similar inclination and altitude). The proposed ROGER system is designed as a spacecraft with rendezvous capabilities including inspection in the vicinity of the

Configuration of management accounting information system for multi-stage manufacturing

Science.gov (United States)

Mkrtychev, S. V.; Ochepovsky, A. V.; Enik, O. A.

2018-05-01

The article presents an approach to configuration of a management accounting information system (MAIS) that provides automated calculations and the registration of normative production losses in multi-stage manufacturing. The use of MAIS with the proposed configuration at the enterprises of textile and woodworking industries made it possible to increase the accuracy of calculations for normative production losses and to organize accounting thereof with the reference to individual stages of the technological process. Thus, high efficiency of multi-stage manufacturing control is achieved.

JSC Orbital Debris Website Description

Science.gov (United States)

Johnson, Nicholas L.

2006-01-01

required. These data also help in the analysis and interpretation of impact features on returned spacecraft surfaces. 4) Mitigation - Controlling the growth of the orbital debris population is a high priority for NASA, the United States, and the major space-faring nations of the world to preserve near-Earth space for future generations. Mitigation measures can take the form of curtailing or preventing the creation of new debris, designing satellites to withstand impacts by small debris, and implementing operational procedures ranging from utilizing orbital regimes with less debris, adopting specific spacecraft attitudes, and even maneuvering to avoid collisions with debris. Downloadable items include several documents in PDF format and executable software.and 5) Reentry - Because of the increasing number of objects in space, NASA has adopted guidelines and assessment procedures to reduce the number of non-operational spacecraft and spent rocket upper stages orbiting the Earth. One method of postmission disposal is to allow reentry of these spacecraft, either from orbital decay (uncontrolled entry) or with a controlled entry. Orbital decay may be achieved by firing engines to lower the perigee altitude so that atmospheric drag will eventually cause the spacecraft to enter. However, the surviving debris impact footprint cannot be guaranteed to avoid inhabited landmasses. Controlled entry normally occurs by using a larger amount of propellant with a larger propulsion system to drive the spacecraft to enter the atmosphere at a steeper flight path angle. It will then enter at a more precise latitude, longitude, and footprint in a nearly uninhabited impact region, generally located in the ocean.

Risk Assessment Challenges in the Ares I Upper Stage

Science.gov (United States)

Stott, James E.; Ring, Robert W.; Elrada, Hassan A.; Hark, Frank

2007-01-01

NASA Marshall Space Flight Center (MSFC) is currently at work developing hardware and systems for the Ares I rocket that will send future astronauts into orbit. Built on cutting-edge launch technologies, evolved powerful Apollo and Space Shuttle propulsion elements, and decades of NASA spaceflight experience, Ares I is the essential core of a safe, reliable, cost-effective space transportation system -- one that will carry crewed missions back to the moon, on to Mars and out into the solar system. Ares I is an in-line, two-stage rocket configuration topped by the Orion crew vehicle and its launch abort system. In addition to the vehicle's primary mission -carrying crews of four to six astronauts to Earth orbit --Ares I may also use its 25-ton payload capacity to deliver resources and supplies to the International Space Station, or to "park" payloads in orbit for retrieval by other spacecraft bound for the moon or other destinations. Crew transportation to the International Space Station is planned to begin no later than 2014. The first lunar excursion is scheduled for the 2020 timeframe. This paper presents the challenges in designing the Ares I upper stage for reliability and safety while minimizing weight and maximizing performance.

A Ballistic Limit Analysis Program for Shielding Against Micrometeoroids and Orbital Debris

Science.gov (United States)

Ryan, Shannon; Christiansen, Erie

2010-01-01

A software program has been developed that enables the user to quickly and simply perform ballistic limit calculations for common spacecraft structures that are subject to hypervelocity impact of micrometeoroid and orbital debris (MMOD) projectiles. This analysis program consists of two core modules: design, and; performance. The design module enables a user to calculate preliminary dimensions of a shield configuration (e.g., thicknesses/areal densities, spacing, etc.) for a ?design? particle (diameter, density, impact velocity, incidence). The performance module enables a more detailed shielding analysis, providing the performance of a user-defined shielding configuration over the range of relevant in-orbit impact conditions.

A Collaborative Analysis Tool for Integrating Hypersonic Aerodynamics, Thermal Protection Systems, and RBCC Engine Performance for Single Stage to Orbit Vehicles

Science.gov (United States)

Stanley, Thomas Troy; Alexander, Reginald

1999-01-01

Presented is a computer-based tool that connects several disciplines that are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to every other system, as is the case of SSTO vehicles with air breathing propulsion, which is currently being studied by NASA. The deficiencies in the scramjet powered concept led to a revival of interest in Rocket-Based Combined-Cycle (RBCC) propulsion systems. An RBCC propulsion system integrates airbreathing and rocket propulsion into a single engine assembly enclosed within a cowl or duct. A typical RBCC propulsion system operates as a ducted rocket up to approximately Mach 3. At this point the transitions to a ramjet mode for supersonic-to-hypersonic acceleration. Around Mach 8 the engine transitions to a scram4jet mode. During the ramjet and scramjet modes, the integral rockets operate as fuel injectors. Around Mach 10-12 (the actual value depends on vehicle and mission requirements), the inlet is physically closed and the engine transitions to an integral rocket mode for orbit insertion. A common feature of RBCC propelled vehicles is the high degree of integration between the propulsion system and airframe. At high speeds the vehicle forebody is fundamentally part of the engine inlet, providing a compression surface for air flowing into the engine. The compressed air is mixed with fuel and burned. The combusted mixture must be expanded to an area larger than the incoming stream to provide thrust. Since a conventional nozzle would be too large, the entire lower after body of the vehicle is used as an expansion surface. Because of the high external temperatures seen during atmospheric flight, the design of an airbreathing SSTO vehicle requires delicate tradeoffs between engine design, vehicle shape, and thermal protection system (TPS) sizing in order to produce an optimum system in terms of weight (and cost) and maximum performance.

Long-term orbit prediction for Tiangong-1 spacecraft using the mean atmosphere model

Science.gov (United States)

Tang, Jingshi; Liu, Lin; Cheng, Haowen; Hu, Songjie; Duan, Jianfeng

2015-03-01

China is planning to complete its first space station by 2020. For the long-term management and maintenance, the orbit of the space station needs to be predicted for a long period of time. Since the space station is expected to work in a low-Earth orbit, the error in the a priori atmosphere model contributes significantly to the rapid increase of the predicted orbit error. When the orbit is predicted for 20 days, the error in the a priori atmosphere model, if not properly corrected, could induce a semi-major axis error of up to a few kilometers and an overall position error of several thousand kilometers respectively. In this work, we use a mean atmosphere model averaged from NRLMSISE00. The a priori reference mean density can be corrected during the orbit determination. For the long-term orbit prediction, we use sufficiently long period of observations and obtain a series of the diurnal mean densities. This series contains the recent variation of the atmosphere density and can be analyzed for various periodic components. After being properly fitted, the mean density can be predicted and then applied in the orbit prediction. Here we carry out the test with China's Tiangong-1 spacecraft at the altitude of about 340 km and we show that this method is simple and flexible. The densities predicted with this approach can serve in the long-term orbit prediction. In several 20-day prediction tests, most predicted orbits show semi-major axis errors better than 700 m and overall position errors better than 400 km.

Collision risk investigation for an operational spacecraft caused by space debris

Science.gov (United States)

Zhang, Binbin; Wang, Zhaokui; Zhang, Yulin

2017-04-01

The collision probability between an operational spacecraft and a population of space debris is investigated. By dividing the 3-dimensional operational space of the spacecraft into several space volume cells (SVC) and proposing a boundary selection method to calculate the collision probability in each SVC, the distribution of the collision risk, as functions of the time, the orbital height, the declination, the impact elevation, the collision velocity, etc., can be obtained. Thus, the collision risk could be carefully evaluated over a time span for the general orbital configurations of the spacecraft and the space debris. As an application, the collision risk for the Tiangong-2 space laboratory caused by the cataloged space debris is discussed and evaluated. Results show that most of the collision threat comes from the front left and front right in Tiangong-2's local, quasi-horizontal plane. And the collision probability will also accumulate when Tiangong-2 moves to the largest declinations (about {±} 42°). As a result, the manned space activities should be avoided at those declinations.

Secure communications with low-orbit spacecraft using quantum cryptography

Science.gov (United States)

Hughes, Richard J.; Buttler, William T.; Kwiat, Paul G.; Luther, Gabriel G.; Morgan, George L; Nordholt, Jane E.; Peterson, Charles G.; Simmons, Charles M.

1999-01-01

Apparatus and method for secure communication between an earth station and spacecraft. A laser outputs single pulses that are split into preceding bright pulses and delayed attenuated pulses, and polarized. A Pockels cell changes the polarization of the polarized delayed attenuated pulses according to a string of random numbers, a first polarization representing a "1," and a second polarization representing a "0." At the receiving station, a beamsplitter randomly directs the preceding bright pulses and the polarized delayed attenuated pulses onto longer and shorter paths, both terminating in a beamsplitter which directs the preceding bright pulses and a first portion of the polarized delayed attenuated pulses to a first detector, and a second portion of the polarized delayed attenuated pulses to a second detector to generate a key for secure communication between the earth station and the spacecraft.

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

Science.gov (United States)

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.

Titan Orbiter Aerorover Mission

Science.gov (United States)

Sittler Jr., E. C.; Acuna, M.; Burchell, M. J.; Coates, A.; Farrell, W.; Flasar, M.; Goldstein, B. E.; Gorevan, S.; Hartle, R. E.; Johnson, W. T. K.

2001-01-01

We propose a combined Titan orbiter and Titan Aerorover mission with an emphasis on both in situ and remote sensing measurements of Titan's surface, atmosphere, ionosphere, and magnetospheric interaction. The biological aspect of the Titan environment will be emphasized by the mission (i.e., search for organic materials which may include simple organics to 'amono' analogues of amino acids and possibly more complex, lightening detection and infrared, ultraviolet, and charged particle interactions with Titan's surface and atmosphere). An international mission is assumed to control costs. NASA will provide the orbiter, launch vehicle, DSN coverage and operations, while international partners will provide the Aerorover and up to 30% of the cost for the scientific instruments through collaborative efforts. To further reduce costs we propose a single PI for orbiter science instruments and a single PI for Aerorover science instruments. This approach will provide single command/data and power interface between spacecraft and orbiter instruments that will have redundant central DPU and power converter for their instruments. A similar approach could be used for the Aerorover. The mission profile will be constructed to minimize conflicts between Aerorover science, orbiter radar science, orbiter radio science, orbiter imaging science, and orbiter fields and particles (FP) science. Additional information is contained in the original extended abstract.

High-order moments of spin-orbit energy in a multielectron configuration

Science.gov (United States)

Na, Xieyu; Poirier, M.

2016-07-01

In order to analyze the energy-level distribution in complex ions such as those found in warm dense plasmas, this paper provides values for high-order moments of the spin-orbit energy in a multielectron configuration. Using second-quantization results and standard angular algebra or fully analytical expressions, explicit values are given for moments up to 10th order for the spin-orbit energy. Two analytical methods are proposed, using the uncoupled or coupled orbital and spin angular momenta. The case of multiple open subshells is considered with the help of cumulants. The proposed expressions for spin-orbit energy moments are compared to numerical computations from Cowan's code and agree with them. The convergence of the Gram-Charlier expansion involving these spin-orbit moments is analyzed. While a spectrum with infinitely thin components cannot be adequately represented by such an expansion, a suitable convolution procedure ensures the convergence of the Gram-Charlier series provided high-order terms are accounted for. A corrected analytical formula for the third-order moment involving both spin-orbit and electron-electron interactions turns out to be in fair agreement with Cowan's numerical computations.

Mesh Network Architecture for Enabling Inter-Spacecraft Communication

Science.gov (United States)

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..

Simultaneous spacecraft orbit estimation and control based on GPS measurements via extended Kalman filter

Directory of Open Access Journals (Sweden)

Tamer Mekky Ahmed Habib

2013-06-01

Full Text Available The primary aim of this work is to provide simultaneous spacecraft orbit estimation and control based on the global positioning system (GPS measurements suitable for application to the next coming Egyptian remote sensing satellites. Disturbance resulting from earth’s oblateness till the fourth order (i.e., J4 is considered. In addition, aerodynamic drag and random disturbance effects are taken into consideration.

Advanced Solar-propelled Cargo Spacecraft for Mars Missions

Science.gov (United States)

Auziasdeturenne, Jacqueline; Beall, Mark; Burianek, Joseph; Cinniger, Anna; Dunmire, Barbrina; Haberman, Eric; Iwamoto, James; Johnson, Stephen; Mccracken, Shawn; Miller, Melanie

1989-01-01

Three concepts for an unmanned, solar powered, cargo spacecraft for Mars support missions were investigated. These spacecraft are designed to carry a 50,000 kg payload from a low Earth orbit to a low Mars orbit. Each design uses a distinctly different propulsion system: A Solar Radiation Absorption (SRA) system, a Solar-Pumped Laser (SPL) system and a solar powered magnetoplasmadynamic (MPD) arc system. The SRA directly converts solar energy to thermal energy in the propellant through a novel process. In the SPL system, a pair of solar-pumped, multi-megawatt, CO2 lasers in sunsynchronous Earth orbit converts solar energy to laser energy. The MPD system used indium phosphide solar cells to convert sunlight to electricity, which powers the propulsion system. Various orbital transfer options are examined for these concepts. In the SRA system, the mother ship transfers the payload into a very high Earth orbit and a small auxiliary propulsion system boosts the payload into a Hohmann transfer to Mars. The SPL spacecraft and the SPL powered spacecraft return to Earth for subsequent missions. The MPD propelled spacecraft, however, remains at Mars as an orbiting space station. A patched conic approximation was used to determine a heliocentric interplanetary transfer orbit for the MPD propelled spacecraft. All three solar-powered spacecraft use an aerobrake procedure to place the payload into a low Mars parking orbit. The payload delivery times range from 160 days to 873 days (2.39 years).

Guidance and control of swarms of spacecraft

Science.gov (United States)

Morgan, Daniel James

using computer simulations. The swarm-keeping problem can be solved by placing the spacecraft on J2-invariant relative orbits, which prevent collisions and minimize the drift of the swarm over hundreds of orbits using a single burn. These orbits are achieved by energy matching the spacecraft to the reference orbit. Additionally, these conditions can be repeatedly applied to minimize the drift of the swarm when atmospheric drag has a large effect (orbits with an altitude under 500 km). The swarm reconfiguration is achieved using two steps: trajectory optimization and assignment. The trajectory optimization problem can be written as a nonlinear, optimal control problem. This optimal control problem is discretized, decoupled, and convexified so that the individual femtosats can efficiently solve the optimization. Sequential convex programming is used to generate the control sequences and trajectories required to safely and efficiently transfer a spacecraft from one position to another. The sequence of trajectories is shown to converge to a Karush-Kuhn-Tucker point of the nonconvex problem. In the case where many of the spacecraft are interchangeable, a variable-swarm, distributed auction algorithm is used to determine the assignment of spacecraft to target positions. This auction algorithm requires only local communication and all of the bidding parameters are stored locally. The assignment generated using this auction algorithm is shown to be near optimal and to converge in a finite number of bids. Additionally, the bidding process is used to modify the number of targets used in the assignment so that the reconfiguration can be achieved even when there is a disconnected communication network or a significant loss of agents. Once the assignment is achieved, the trajectory optimization can be run using the terminal positions determined by the auction algorithm. To implement these algorithms in real time a model predictive control formulation is used. Model predictive

Orbiting Depot and Reusable Lander for Lunar Transportation

Science.gov (United States)

Petro, Andrew

2009-01-01

A document describes a conceptual transportation system that would support exploratory visits by humans to locations dispersed across the surface of the Moon and provide transport of humans and cargo to sustain one or more permanent Lunar outpost. The system architecture reflects requirements to (1) minimize the amount of vehicle hardware that must be expended while maintaining high performance margins and (2) take advantage of emerging capabilities to produce propellants on the Moon while also enabling efficient operation using propellants transported from Earth. The system would include reusable single- stage lander spacecraft and a depot in a low orbit around the Moon. Each lander would have descent, landing, and ascent capabilities. A crew-taxi version of the lander would carry a pressurized crew module; a cargo version could carry a variety of cargo containers. The depot would serve as a facility for storage and for refueling with propellants delivered from Earth or propellants produced on the Moon. The depot could receive propellants and cargo sent from Earth on a variety of spacecraft. The depot could provide power and orbit maintenance for crew vehicles from Earth and could serve as a safe haven for lunar crews pending transport back to Earth.

Earth to Orbit Beamed Energy Experiment

Science.gov (United States)

Johnson, Les; Montgomery, Edward E.

2017-01-01

As a means of primary propulsion, beamed energy propulsion offers the benefit of offloading much of the propulsion system mass from the vehicle, increasing its potential performance and freeing it from the constraints of the rocket equation. For interstellar missions, beamed energy propulsion is arguably the most viable in the near- to mid-term. A near-term demonstration showing the feasibility of beamed energy propulsion is necessary and, fortunately, feasible using existing technologies. Key enabling technologies are large area, low mass spacecraft and efficient and safe high power laser systems capable of long distance propagation. NASA is currently developing the spacecraft technology through the Near Earth Asteroid Scout solar sail mission and has signed agreements with the Planetary Society to study the feasibility of precursor laser propulsion experiments using their LightSail-2 solar sail spacecraft. The capabilities of Space Situational Awareness assets and the advanced analytical tools available for fine resolution orbit determination now make it possible to investigate the practicalities of an Earth-to-orbit Beamed Energy eXperiment (EBEX) - a demonstration at delivered power levels that only illuminate a spacecraft without causing damage to it. The degree to which this can be expected to produce a measurable change in the orbit of a low ballistic coefficient spacecraft is investigated. Key system characteristics and estimated performance are derived for a near term mission opportunity involving the LightSail-2 spacecraft and laser power levels modest in comparison to those proposed previously. While the technology demonstrated by such an experiment is not sufficient to enable an interstellar precursor mission, if approved, then it would be the next step toward that goal.

Deployable Brake for Spacecraft

Science.gov (United States)

Rausch, J. R.; Maloney, J. W.

1987-01-01

Aerodynamic shield that could be opened and closed proposed. Report presents concepts for deployable aerodynamic brake. Brake used by spacecraft returning from high orbit to low orbit around Earth. Spacecraft makes grazing passes through atmosphere to slow down by drag of brake. Brake flexible shield made of woven metal or ceramic withstanding high temperatures created by air friction. Stored until needed, then deployed by set of struts.

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

Science.gov (United States)

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.

Collective excitations in circular atomic configurations and single-photon traps

International Nuclear Information System (INIS)

Hammer, Hanno

2004-01-01

Correlated excitations in a plane circular configuration of identical atoms with parallel dipole moments are investigated. The collective energy eigenstates, which are formally identical to Frenkel excitons, can be computed together with their level shifts and decay rates by decomposing the atomic state space into carrier spaces for the irreducible representations of the symmetry group Z N of the circle. It is shown that the index p of these representations can be used as a quantum number analogously to the orbital angular momentum quantum number l in hydrogenlike systems. Just as the hydrogen s states are the only electronic wave functions which can occupy the central region of the Coulomb potential, the quasiparticle corresponding to a collective excitation of the atoms in the circle can occupy the central atom only for vanishing Z N quantum number p. If a central atom is present, the p=0 state splits into two and shows level crossing at certain radii; in the regions between these radii, damped quantum beats between two 'extreme' p=0 configurations occur. The physical mechanisms behind super- and subradiance at a given radius are discussed. It is shown that, beyond a certain critical number of atoms in the circle, the lifetime of the maximally subradiant state increases exponentially with the number of atoms in the configuration, making the system a natural candidate for a single-photon trap

Long-term orbit prediction for China's Tiangong-1 spacecraft based on mean atmosphere model

Science.gov (United States)

Tang, Jingshi; Liu, Lin; Miao, Manqian

Tiangong-1 is China's test module for future space station. It has gone through three successful rendezvous and dockings with Shenzhou spacecrafts from 2011 to 2013. For the long-term management and maintenance, the orbit sometimes needs to be predicted for a long period of time. As Tiangong-1 works in a low-Earth orbit with an altitude of about 300-400 km, the error in the a priori atmosphere model contributes significantly to the rapid increase of the predicted orbit error. When the orbit is predicted for 10-20 days, the error in the a priori atmosphere model, if not properly corrected, could induce the semi-major axis error and the overall position error up to a few kilometers and several thousand kilometers respectively. In this work, we use a mean atmosphere model averaged from NRLMSIS00. The a priori reference mean density can be corrected during precise orbit determination (POD). For applications in the long-term orbit prediction, the observations are first accumulated. With sufficiently long period of observations, we are able to obtain a series of the diurnal mean densities. This series bears the recent variation of the atmosphere density and can be analyzed for various periods. After being properly fitted, the mean density can be predicted and then applied in the orbit prediction. We show that the densities predicted with this approach can serve to increase the accuracy of the predicted orbit. In several 20-day prediction tests, most predicted orbits show semi-major axis errors better than 700m and overall position errors better than 600km.

Spacecraft navigation at Mars using earth-based and in situ radio tracking techniques

Science.gov (United States)

Thurman, S. W.; Edwards, C. D.; Kahn, R. D.; Vijayaraghavan, A.; Hastrup, R. C.; Cesarone, R. J.

1992-08-01

A survey of earth-based and in situ radiometric data types and results from a number of studies investigating potential radio navigation performance for spacecraft approaching/orbiting Mars and for landed spacecraft and rovers on the surface of Mars are presented. The performance of Doppler, ranging and interferometry earth-based data types involving single or multiple spacecraft is addressed. This evaluation is conducted with that of in situ data types, such as Doppler and ranging measurements between two spacecraft near Mars, or between a spacecraft and one or more surface radio beacons.

A Brief History of Meteoroid and Orbital Debris Shielding Technology for US Manned Spacecraft

Science.gov (United States)

Bjorkman, Michael D.; Hyde, James L.

2008-01-01

Meteoroid and orbital debris shielding has played an important role from the beginning of manned spaceflight. During the early 60 s, meteoroid protection drove requirements for new meteor and micrometeoroid impact science. Meteoroid protection also stimulated advances in the technology of hypervelocity impact launchers and impact damage assessment methodologies. The first phase of meteoroid shielding assessments closed in the early 70 s with the end of the Apollo program. The second phase of meteoroid protection technology began in the early 80 s when it was determined that there is a manmade Earth orbital debris belt that poses a significant risk to LEO manned spacecraft. The severity of the Earth orbital debris environment has dictated changes in Space Shuttle and ISS operations as well as driven advances in shielding technology and assessment methodologies. A timeline of shielding technology and assessment methodology advances is presented along with a summary of risk assessment results.

Comprehensive Evaluation of Attitude and Orbit Estimation Using Actual Earth Magnetic Field Data

Science.gov (United States)

Deutschmann, Julie K.; Bar-Itzhack, Itzhack Y.

2000-01-01

A single, augmented Extended Kalman Filter (EKF), which simultaneously and autonomously estimates spacecraft attitude and orbit has been developed and successfully tested with real magnetometer and gyro data only. Because the earth magnetic field is a function of time and position, and because time is known quite precisely, the differences between the computed and measured magnetic field components, as measured by the magnetometers throughout the entire spacecraft orbit, are a function of both orbit and attitude errors. Thus, conceivably these differences could be used to estimate both orbit and attitude; an observability study validated this assumption. The results of testing the EKF with actual magnetometer and gyro data, from four satellites supported by the NASA Goddard Space Flight Center (GSFC) Guidance, Navigation, and Control Center, are presented and evaluated. They confirm the assumption that a single EKF can estimate both attitude and orbit when using gyros and magnetometers only.

Formation of Field-reversed-Configuration Plasma with Punctuated-betatron-orbit Electrons

International Nuclear Information System (INIS)

Welch, D.R.; Cohen, S.A.; Genoni, T.C.; Glasser, A.H.

2010-01-01

We describe ab initio, self-consistent, 3D, fully electromagnetic numerical simulations of current drive and field-reversed-configuration plasma formation by odd-parity rotating magnetic fields (RMFo). Magnetic-separatrix formation and field reversal are attained from an initial mirror configuration. A population of punctuated-betatron-orbit electrons, generated by the RMFo, carries the majority of the field-normal azimuthal electrical current responsible for field reversal. Appreciable current and plasma pressure exist outside the magnetic separatrix whose shape is modulated by the RMFo phase. The predicted plasma density and electron energy distribution compare favorably with RMFo experiments.

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

Science.gov (United States)

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

Production of activation products in space-craft components by protons in low earth orbit

International Nuclear Information System (INIS)

Normand, E.; Johnson, M.L.

1986-01-01

A spacecraft orbiting the Earth through trapped radiation belts will be subject to an induced effect as well as to the direct irradiation by the protons and electrons of the trapped belts. This induced effect is activation of the spacecraft materials by the trapped belt protons. This activation will produce many radioisotopes having half-lives ranging from seconds to millions of years, and emitting various types of radiation. Of primary concern are radioisotopes that emit gamma rays and have half-lives of several years or less. Cross-section data sets are currently being compiled for proton-induced activation products by the Los Alamos National Laboratory. Despite uncertainties in cross-section data, it is instructive to illustrate the magnitude of activation levels and the resulting dose rates calculated in an approximate manner. A number of simplifying assumptions are made

Comparisons of single-stage and two-stage approaches to genomic selection.

Science.gov (United States)

Schulz-Streeck, Torben; Ogutu, Joseph O; Piepho, Hans-Peter

2013-01-01

Genomic selection (GS) is a method for predicting breeding values of plants or animals using many molecular markers that is commonly implemented in two stages. In plant breeding the first stage usually involves computation of adjusted means for genotypes which are then used to predict genomic breeding values in the second stage. We compared two classical stage-wise approaches, which either ignore or approximate correlations among the means by a diagonal matrix, and a new method, to a single-stage analysis for GS using ridge regression best linear unbiased prediction (RR-BLUP). The new stage-wise method rotates (orthogonalizes) the adjusted means from the first stage before submitting them to the second stage. This makes the errors approximately independently and identically normally distributed, which is a prerequisite for many procedures that are potentially useful for GS such as machine learning methods (e.g. boosting) and regularized regression methods (e.g. lasso). This is illustrated in this paper using componentwise boosting. The componentwise boosting method minimizes squared error loss using least squares and iteratively and automatically selects markers that are most predictive of genomic breeding values. Results are compared with those of RR-BLUP using fivefold cross-validation. The new stage-wise approach with rotated means was slightly more similar to the single-stage analysis than the classical two-stage approaches based on non-rotated means for two unbalanced datasets. This suggests that rotation is a worthwhile pre-processing step in GS for the two-stage approaches for unbalanced datasets. Moreover, the predictive accuracy of stage-wise RR-BLUP was higher (5.0-6.1%) than that of componentwise boosting.

Fundamentals of the orbit and response for TianQin

Science.gov (United States)

Hu, Xin-Chun; Li, Xiao-Hong; Wang, Yan; Feng, Wen-Fan; Zhou, Ming-Yue; Hu, Yi-Ming; Hu, Shou-Cun; Mei, Jian-Wei; Shao, Cheng-Gang

2018-05-01

TianQin is a space-based laser interferometric gravitational wave detector aimed at detecting gravitational waves at low frequencies (0.1 mHz–1 Hz). It is formed by three identical drag-free spacecrafts in an equilateral triangular constellation orbiting around the Earth. The distance between each pair of spacecrafts is approximately 1.7 × 105 ~km . The spacecrafts are interconnected by infrared laser beams forming up to three Michelson-type interferometers. The detailed mission design and the study of science objectives for the TianQin project depend crucially on the orbit and the response of the detector. In this paper, we provide the analytic expressions for the coordinates of the orbit for each spacecraft in the heliocentric-ecliptic coordinate system to the leading orders. This enables a sufficiently accurate study of science objectives and data analysis, and serves as a first step to further orbit design and optimization. We calculate the response of a single Michelson detector to plane gravitational waves in arbitrary waveform which is valid in the full range of the sensitive frequencies. It is then used to generate the more realistic sensitivity curve of TianQin. We apply this model on a reference white-dwarf binary as a proof of principle.

Attitude and Configuration Control of Flexible Multi-Body Spacecraft

Science.gov (United States)

Cho, Sung-Ki; Cochran, John E., Jr.

2002-06-01

Multi-body spacecraft attitude and configuration control formulations based on the use of collaborative control theory are considered. The control formulations are based on two-player, nonzero-sum, differential game theory applied using a Nash strategy. It is desired that the control laws allow different components of the multi-body system to perform different tasks. For example, it may be desired that one body points toward a fixed star while another body in the system slews to track another satellite. Although similar to the linear quadratic regulator formulation, the collaborative control formulation contains a number of additional design parameters because the problem is formulated as two control problems coupled together. The use of the freedom of the partitioning of the total problem into two coupled control problems and the selection of the elements of the cross-coupling matrices are specific problems addressed in this paper. Examples are used to show that significant improvement in performance, as measured by realistic criteria, of collaborative control over conventional linear quadratic regulator control can be achieved by using proposed design guidelines.

Attitude and Configuration Control of Flexible Multi-Body Spacecraft

Directory of Open Access Journals (Sweden)

Sungki Cho

2002-06-01

Full Text Available Multi-body spacecraft attitude and configuration control formulations based on the use of collaborative control theory are considered. The control formulations are based on two-player, nonzero-sum, differential game theory applied using a Nash strategy. It is desired that the control laws allow different components of the multi-body system to perform different tasks. For example, it may be desired that one body points toward a fixed star while another body in the system slews to track another satellite. Although similar to the linear quadratic regulator formulation, the collaborative control formulation contains a number of additional design parameters because the problem is formulated as two control problems coupled together. The use of the freedom of the partitioning of the total problem into two coupled control problems and the selection of the elements of the cross-coupling matrices are specific problems addressed in this paper. Examples are used to show that significant improvement in performance, as measured by realistic criteria, of collaborative control over conventional linear quadratic regulator control can be achieved by using proposed design guidelines.

Automating Trend Analysis for Spacecraft Constellations

Science.gov (United States)

Davis, George; Cooter, Miranda; Updike, Clark; Carey, Everett; Mackey, Jennifer; Rykowski, Timothy; Powers, Edward I. (Technical Monitor)

2001-01-01

Spacecraft trend analysis is a vital mission operations function performed by satellite controllers and engineers, who perform detailed analyses of engineering telemetry data to diagnose subsystem faults and to detect trends that may potentially lead to degraded subsystem performance or failure in the future. It is this latter function that is of greatest importance, for careful trending can often predict or detect events that may lead to a spacecraft's entry into safe-hold. Early prediction and detection of such events could result in the avoidance of, or rapid return to service from, spacecraft safing, which not only results in reduced recovery costs but also in a higher overall level of service for the satellite system. Contemporary spacecraft trending activities are manually intensive and are primarily performed diagnostically after a fault occurs, rather than proactively to predict its occurrence. They also tend to rely on information systems and software that are oudated when compared to current technologies. When coupled with the fact that flight operations teams often have limited resources, proactive trending opportunities are limited, and detailed trend analysis is often reserved for critical responses to safe holds or other on-orbit events such as maneuvers. While the contemporary trend analysis approach has sufficed for current single-spacecraft operations, it will be unfeasible for NASA's planned and proposed space science constellations. Missions such as the Dynamics, Reconnection and Configuration Observatory (DRACO), for example, are planning to launch as many as 100 'nanospacecraft' to form a homogenous constellation. A simple extrapolation of resources and manpower based on single-spacecraft operations suggests that trending for such a large spacecraft fleet will be unmanageable, unwieldy, and cost-prohibitive. It is therefore imperative that an approach to automating the spacecraft trend analysis function be studied, developed, and applied to

Comprehensive evaluation of attitude and orbit estimation using real earth magnetic field data

Science.gov (United States)

Deutschmann, Julie; Bar-Itzhack, Itzhack

1997-01-01

A single, augmented extended Kalman filter (EKF) which simultaneously and autonomously estimates spacecraft attitude and orbit was developed and tested with simulated and real magnetometer and rate data. Since the earth's magnetic field is a function of time and position, and since time is accurately known, the differences between the computed and measured magnetic field components, as measured by the magnetometers throughout the entire spacecraft's orbit, are a function of orbit and attitude errors. These differences can be used to estimate the orbit and attitude. The test results of the EKF with magnetometer and gyro data from three NASA satellites are presented and evaluated.

Return to Venus of AKATSUKI, the Japanese Venus Orbiter

Science.gov (United States)

Nakamura, M.; Iwagami, N.; Satoh, T.; Taguchi, M.; Watanabe, S.; Takahashi, Y.; Imamura, T.; Suzuki, M.; Ueno, M.; Yamazaki, A.; Fukuhara, T.; Yamada, M.; Ishii, N.; Ogohara, K.

2011-12-01

Japanese Venus Climate Orbiter 'AKATSUKI' (PLANET-C) was proposed in 2001 with strong support by international Venus science community and approved as an ISAS mission soon after the proposal. AKATSUKI and ESA's Venus Express complement each other in Venus climate study. Various coordinated observations using the two spacecraft have been planned. Also participating scientists from US have been selected. Its science target is to understand the climate of Venus. The mission life we expected was more than 2 Earth years in Venus orbit. AKATSUKI was successfully launched at 06:58:22JST on May 21, by H-IIA F17. After the separation from H-IIA, the telemetry from AKATSUKI was normally detected by DSN Goldstone station (10:00JST) and the solar cell paddles' expansion was confirmed. AKATSUKI was put into the 3-axis stabilized mode in the initial operation from Uchinoura station and the critical operation was finished at 20:00JST on the same day. The malfunction, which happened during the Venus Orbit Insertion (VOI) on7 Dec, 2010 is as follows. We set all commands on Dec. 5. Attitude control for Venus orbit insertion (VOI) was automatically done on Dec. 6. Orbital maneuver engine (OME) was fired 08:49 JST on Dec. 7. 1min. after firing the spacecraft went into the occultation region and we had no telemetry, but we expected to continuous firing for 12min. Recording on the spacecraft told us later that, unfortunately the firing continued just 152sec. and stopped. The reason of the malfunction of the OME was the blocking of check valve of the gas pressure line to push the fuel to the engine. We failed to make the spacecraft the Venus orbiter, and it is rotating the sun with the orbital period of 203 days. As the Venus orbit the sun with the period of 225 days, AKATSUKI has a chance to meet Venus again in 5 or 6 years depending on the orbit correction plan. Let us summarize the present situation of AKATSUKI. Most of the fuel still remains. But the condition of the propulsion

Development of an advanced spacecraft tandem mass spectrometer

Science.gov (United States)

Drew, Russell C.

1992-03-01

The purpose of this research was to apply current advanced technology in electronics and materials to the development of a miniaturized Tandem Mass Spectrometer that would have the potential for future development into a package suitable for spacecraft use. The mass spectrometer to be used as a basis for the tandem instrument would be a magnetic sector instrument, of Nier-Johnson configuration, as used on the Viking Mars Lander mission. This instrument configuration would then be matched with a suitable second stage MS to provide the benefits of tandem MS operation for rapid identification of unknown organic compounds. This tandem instrument is configured with a newly designed GC system to aid in separation of complex mixtures prior to MS analysis. A number of important results were achieved in the course of this project. Among them were the development of a miniaturized GC subsystem, with a unique desorber-injector, fully temperature feedback controlled oven with powered cooling for rapid reset to ambient conditions, a unique combination inlet system to the MS that provides for both membrane sampling and direct capillary column sample transfer, a compact and ruggedized alignment configuration for the MS, an improved ion source design for increased sensitivity, and a simple, rugged tandem MS configuration that is particularly adaptable to spacecraft use because of its low power and low vacuum pumping requirements. The potential applications of this research include use in manned spacecraft like the space station as a real-time detection and warning device for the presence of potentially harmful trace contaminants of the spacecraft atmosphere, use as an analytical device for evaluating samples collected on the Moon or a planetary surface, or even use in connection with monitoring potentially hazardous conditions that may exist in terrestrial locations such as launch pads, environmental test chambers or other sensitive areas. Commercial development of the technology

Integrated orbital servicing study follow-on. Volume 2: Technical analysis and system design

Science.gov (United States)

1978-01-01

In-orbit service functional and physical requirements to support both low and high Earth orbit servicing/maintenance operations were defined, an optimum servicing system configuration was developed and mockups and early prototype hardware were fabricated to demonstrate and validate the concepts selected. Significant issues addressed include criteria for concept selection; representative mission equipment and approaches to their design for serviceability; significant serviceable spacecraft design aspects; servicer mechanism operation in one-g; approaches for the demonstration/simulation; and service mechanism structure design approach.

Observations on the reliability of COTS-device-based solid state data recorders operating in low-earth orbit

International Nuclear Information System (INIS)

Underwood, C.I.

1999-01-01

This paper presents the results of Surrey Space Centre's experience in using different coding schemes and hardware configurations to protect data and protect data and software stored in COTS-device (Commercial-Off-The-Shelf) based memories on-board operational spacecraft in low Earth orbit. (author)

Space Launch System Spacecraft and Payload Elements: Progress Toward Crewed Launch and Beyond

Science.gov (United States)

Schorr, Andrew A.; Smith, David Alan; Holcomb, Shawn; Hitt, David

2017-01-01

While significant and substantial progress continues to be accomplished toward readying the Space Launch System (SLS) rocket for its first test flight, work is already underway on preparations for the second flight - using an upgraded version of the vehicle - and beyond. Designed to support human missions into deep space, SLS is the most powerful human-rated launch vehicle the United States has ever undertaken, and is one of three programs being managed by the National Aeronautics and Space Administration's (NASA's) Exploration Systems Development division. The Orion spacecraft program is developing a new crew vehicle that will support human missions beyond low Earth orbit (LEO), and the Ground Systems Development and Operations (GSDO) program is transforming Kennedy Space Center (KSC) into a next-generation spaceport capable of supporting not only SLS but also multiple commercial users. Together, these systems will support human exploration missions into the proving ground of cislunar space and ultimately to Mars. For its first flight, SLS will deliver a near-term heavy-lift capability for the nation with its 70-metric-ton (t) Block 1 configuration. Each element of the vehicle now has flight hardware in production in support of the initial flight of the SLS, which will propel Orion around the moon and back. Encompassing hardware qualification, structural testing to validate hardware compliance and analytical modeling, progress is on track to meet the initial targeted launch date. In Utah and Mississippi, booster and engine testing are verifying upgrades made to proven shuttle hardware. At Michoud Assembly Facility (MAF) in Louisiana, the world's largest spacecraft welding tool is producing tanks for the SLS core stage. Providing the Orion crew capsule/launch vehicle interface and in-space propulsion via a cryogenic upper stage, the Spacecraft/Payload Integration and Evolution (SPIE) element serves a key role in achieving SLS goals and objectives. The SPIE element

Spacecraft Charging and the Microwave Anisotropy Probe Spacecraft

Science.gov (United States)

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.

Benefits of a Single-Person Spacecraft for Weightless Operations

Science.gov (United States)

Griffin, Brand Norman

2012-01-01

Historically, less than 20 percent of crew time related to extravehicular activity (EVA) is spent on productive external work. For planetary operations space suits are still the logical choice; however for safe and rapid access to the weightless environment, spacecraft offer compelling advantages. FlexCraft, a concept for a single-person spacecraft, enables any-time access to space for short or long excursions by different astronauts. For the International Space Station (ISS), going outside is time-consuming, requiring pre-breathing, donning a fitted space suit, and pumping down an airlock. For each ISS EVA this is between 12.5 and 16 hours. FlexCraft provides immediate access to space because it operates with the same cabin atmosphere as its host. Furthermore, compared to the space suit pure oxygen environment, a mixed gas atmosphere lowers the fire risk and allows use of conventional materials and systems. For getting to the worksite, integral propulsion replaces hand-over-hand translation or having another crew member operate the robotic arm. This means less physical exertion and more time at the work site. Possibly more important, in case of an emergency, FlexCraft can return from the most distant point on ISS in less than a minute. The one-size-fits-all FlexCraft means no on-orbit inventory of parts or crew time required to fit all astronauts. With a shirtsleeve cockpit, conventional displays and controls are used, there is no suit trauma and because the work is not strenuous, no rest days are required. Furthermore, there is no need to collect hand tools because manipulators are equipped with force multiplying end-effectors that can deliver the precise torque for the job. FlexCraft is an efficient solution for asteroid exploration allowing all crew to use one vehicle with no risk of contamination. And, because FlexCraft is a vehicle, its design offers better radiation and micro-meteoroid protection than space suits.

Performance and genome-centric metagenomics of thermophilic single and two-stage anaerobic digesters treating cheese wastes

DEFF Research Database (Denmark)

Fontana, Alessandra; Campanaro, Stefano; Treu, Laura

2018-01-01

-depth characterization of the microbial community structure using genome-centric metagenomics. Both reactor configurations showed acidification problems under the tested organic loading rates (OLRs) of 3.6 and 2.4 g COD/L-reactor day and the hydraulic retention time (HRT) of 15 days. However, the two-stage design...... of the main population genomes highlighted specific metabolic pathways responsible for the AD process and the mechanisms of main intermediates production. Particularly, the acetate accumulation experienced by the single stage configuration was mainly correlated to the low abundant syntrophic acetate oxidizer...

Spacecraft Charging: Hazard Causes, Hazard Effects, Hazard Controls

Science.gov (United States)

Koontz, Steve.

2018-01-01

Spacecraft flight environments are characterized both by a wide range of space plasma conditions and by ionizing radiation (IR), solar ultraviolet and X-rays, magnetic fields, micrometeoroids, orbital debris, and other environmental factors, all of which can affect spacecraft performance. Dr. Steven Koontz's lecture will provide a solid foundation in the basic engineering physics of spacecraft charging and charging effects that can be applied to solving practical spacecraft and spacesuit engineering design, verification, and operations problems, with an emphasis on spacecraft operations in low-Earth orbit, Earth's magnetosphere, and cis-Lunar space.

Preliminary Sizing Completed for Single- Stage-To-Orbit Launch Vehicles Powered By Rocket-Based Combined Cycle Technology

Science.gov (United States)

Roche, Joseph M.

2002-01-01

Single-stage-to-orbit (SSTO) propulsion remains an elusive goal for launch vehicles. The physics of the problem is leading developers to a search for higher propulsion performance than is available with all-rocket power. Rocket-based combined cycle (RBCC) technology provides additional propulsion performance that may enable SSTO flight. Structural efficiency is also a major driving force in enabling SSTO flight. Increases in performance with RBCC propulsion are offset with the added size of the propulsion system. Geometrical considerations must be exploited to minimize the weight. Integration of the propulsion system with the vehicle must be carefully planned such that aeroperformance is not degraded and the air-breathing performance is enhanced. Consequently, the vehicle's structural architecture becomes one with the propulsion system architecture. Geometrical considerations applied to the integrated vehicle lead to low drag and high structural and volumetric efficiency. Sizing of the SSTO launch vehicle (GTX) is itself an elusive task. The weight of the vehicle depends strongly on the propellant required to meet the mission requirements. Changes in propellant requirements result in changes in the size of the vehicle, which in turn, affect the weight of the vehicle and change the propellant requirements. An iterative approach is necessary to size the vehicle to meet the flight requirements. GTX Sizer was developed to do exactly this. The governing geometry was built into a spreadsheet model along with scaling relationships. The scaling laws attempt to maintain structural integrity as the vehicle size is changed. Key aerodynamic relationships are maintained as the vehicle size is changed. The closed weight and center of gravity are displayed graphically on a plot of the synthesized vehicle. In addition, comprehensive tabular data of the subsystem weights and centers of gravity are generated. The model has been verified for accuracy with finite element analysis. The

Particle orbit analysis for LHD helical axis configurations

International Nuclear Information System (INIS)

Guasp, J.; Yamazaki, K.; Motojima, O.

1993-04-01

Fast ion orbits for helical magnetic axis configurations in LHD (Large Helical Device) are analyzed and compared with the standard circular axis case. Boundaries between passing and helically trapped particle regions show clear differences: in the non-planar axis case the helically trapped region spreads, near the magnetic axis, over a much wider band across the 90deg pitch angle value and shows a very marked asymmetry. The locally trapped particle region is also wider than in the standard case. The differences in the loss cone boundaries of the two cases are rather small, however, the effects of re-entering criteria are very important in both cases. On the contrary, effects of finite coil size are not significant. (author)

Removal of pharmaceuticals in pre-denitrifying MBBR – Influence of organic substrate availability in single- and three-stage configurations

DEFF Research Database (Denmark)

Polesel, Fabio; Torresi, Elena; Loreggian, L.

2017-01-01

kinetics were observed, with highest and lowest rates/rate constants in the first (S1) and the last (S3) staged sub-reactors, respectively. These observations were confirmed by removal efficiency data obtained during continuous-flow operation, with limited removal (... and highest removal in S1 within the three-stage MBBR. Notably, biotransformation rate constants obtained for non-recalcitrant pharmaceuticals correlated with mean specific denitrification rates, maximum specific growth rates and observed growth yield values. Overall, these findings suggest that: (i) the long......-term exposure to tiered substrate accessibility in the three-stage configuration shaped the denitrification and biotransformation capacity of biofilms, with significant reduction under substrate limitation; (ii) biotransformation of pharmaceuticals may have occurred as a result of cometabolism by heterotrophic...

How Spacecraft Fly Spaceflight Without Formulae

CERN Document Server

Swinerd, Graham

2009-01-01

About half a century ago a small satellite, Sputnik 1, was launched. The satellite did very little other than to transmit a radio signal to announce its presence in orbit. However, this humble beginning heralded the dawn of the Space Age. Today literally thousands of robotic spacecraft have been launched, many of which have flown to far-flung regions of the Solar System carrying with them the human spirit of scientific discovery and exploration. Numerous other satellites have been launched in orbit around the Earth providing services that support our technological society on the ground. How Spacecraft Fly: Spaceflight Without Formulae by Graham Swinerd focuses on how these spacecraft work. The book opens with a historical perspective of how we have come to understand our Solar System and the Universe. It then progresses through orbital flight, rocket science, the hostile environment within which spacecraft operate, and how they are designed. The concluding chapters give a glimpse of what the 21st century may ...

Spacecraft Software Maintenance: An Effective Approach to Reducing Costs and Increasing Science Return

Science.gov (United States)

Shell, Elaine M.; Lue, Yvonne; Chu, Martha I.

1999-01-01

Flight software is a mission critical element of spacecraft functionality and performance. When ground operations personnel interface to a spacecraft, they are typically dealing almost entirely with the capabilities of onboard software. This software, even more than critical ground/flight communications systems, is expected to perform perfectly during all phases of spacecraft life. Due to the fact that it can be reprogrammed on-orbit to accommodate degradations or failures in flight hardware, new insights into spacecraft characteristics, new control options which permit enhanced science options, etc., the on- orbit flight software maintenance team is usually significantly responsible for the long term success of a science mission. Failure of flight software to perform as needed can result in very expensive operations work-around costs and lost science opportunities. There are three basic approaches to maintaining spacecraft software--namely using the original developers, using the mission operations personnel, or assembling a center of excellence for multi-spacecraft software maintenance. Not planning properly for flight software maintenance can lead to unnecessarily high on-orbit costs and/or unacceptably long delays, or errors, in patch installations. A common approach for flight software maintenance is to access the original development staff. The argument for utilizing the development staff is that the people who developed the software will be the best people to modify the software on-orbit. However, it can quickly becomes a challenge to obtain the services of these key people. They may no longer be available to the organization. They may have a more urgent job to perform, quite likely on another project under different project management. If they havn't worked on the software for a long time, they may need precious time for refamiliarization to the software, testbeds and tools. Further, a lack of insight into issues related to flight software in its on-orbit

Single-stage versus two-stage anaerobic fluidized bed bioreactors in treating municipal wastewater: Performance, foulant characteristics, and microbial community.

Science.gov (United States)

Wu, Bing; Li, Yifei; Lim, Weikang; Lee, Shi Lin; Guo, Qiming; Fane, Anthony G; Liu, Yu

2017-03-01

This study examined the receptive performance, membrane foulant characteristics, and microbial community in the single-stage and two-stage anaerobic fluidized membrane bioreactor (AFMBR) treating settled raw municipal wastewater with the aims to explore fouling mechanisms and microbial community structure in both systems. Both AFMBRs exhibited comparable organic removal efficiency and membrane performances. In the single-stage AFMBR, less soluble organic substances were removed through biosorption by GAC and biodegradation than those in the two-stage AFMBR. Compared to the two-stage AFMBR, the formation of cake layer was the main cause of the observed membrane fouling in the single-stage AFMBR at the same employed flux. The accumulation rate of the biopolymers was linearly correlated with the membrane fouling rate. In the chemical-cleaned foulants, humic acid-like substances and silicon were identified as the predominant organic and inorganic fouants respectively. As such, the fluidized GAC particles might not be effective in removing these substances from the membrane surfaces. High-throughout pyrosequencing analysis further revealed that beta-Proteobacteria were predominant members in both AFMBRs, which contributed to the development of biofilms on the fluidized GAC and membrane surfaces. However, it was also noted that the abundance of the identified dominant in the membrane surface-associated biofilm seemed to be related to the permeate flux and reactor configuration. Copyright © 2016 Elsevier Ltd. All rights reserved.

GOES-R active vibration damping controller design, implementation, and on-orbit performance

Science.gov (United States)

Clapp, Brian R.; Weigl, Harald J.; Goodzeit, Neil E.; Carter, Delano R.; Rood, Timothy J.

2018-01-01

GOES-R series spacecraft feature a number of flexible appendages with modal frequencies below 3.0 Hz which, if excited by spacecraft disturbances, can be sources of undesirable jitter perturbing spacecraft pointing. To meet GOES-R pointing stability requirements, the spacecraft flight software implements an Active Vibration Damping (AVD) rate control law which acts in parallel with the nadir point attitude control law. The AVD controller commands spacecraft reaction wheel actuators based upon Inertial Measurement Unit (IMU) inputs to provide additional damping for spacecraft structural modes below 3.0 Hz which vary with solar wing angle. A GOES-R spacecraft dynamics and attitude control system identified model is constructed from pseudo-random reaction wheel torque commands and IMU angular rate response measurements occurring over a single orbit during spacecraft post-deployment activities. The identified Fourier model is computed on the ground, uplinked to the spacecraft flight computer, and the AVD controller filter coefficients are periodically computed on-board from the Fourier model. Consequently, the AVD controller formulation is based not upon pre-launch simulation model estimates but upon on-orbit nadir point attitude control and time-varying spacecraft dynamics. GOES-R high-fidelity time domain simulation results herein demonstrate the accuracy of the AVD identified Fourier model relative to the pre-launch spacecraft dynamics and control truth model. The AVD controller on-board the GOES-16 spacecraft achieves more than a ten-fold increase in structural mode damping for the fundamental solar wing mode while maintaining controller stability margins and ensuring that the nadir point attitude control bandwidth does not fall below 0.02 Hz. On-orbit GOES-16 spacecraft appendage modal frequencies and damping ratios are quantified based upon the AVD system identification, and the increase in modal damping provided by the AVD controller for each structural mode is

Operational Philosophy Concerning Manned Spacecraft Cabin Leaks

Science.gov (United States)

DeSimpelaere, Edward

2011-01-01

The last thirty years have seen the Space Shuttle as the prime United States spacecraft for manned spaceflight missions. Many lessons have been learned about spacecraft design and operation throughout these years. Over the next few decades, a large increase of manned spaceflight in the commercial sector is expected. This will result in the exposure of commercial crews and passengers to many of the same risks crews of the Space Shuttle have encountered. One of the more dire situations that can be encountered is the loss of pressure in the habitable volume of the spacecraft during on orbit operations. This is referred to as a cabin leak. This paper seeks to establish a general cabin leak response philosophy with the intent of educating future spacecraft designers and operators. After establishing a relative definition for a cabin leak, the paper covers general descriptions of detection equipment, detection methods, and general operational methods for management of a cabin leak. Subsequently, all these items are addressed from the perspective of the Space Shuttle Program, as this will be of the most value to future spacecraft due to similar operating profiles. Emphasis here is placed upon why and how these methods and philosophies have evolved to meet the Space Shuttle s needs. This includes the core ideas of: considerations of maintaining higher cabin pressures vs. lower cabin pressures, the pros and cons of a system designed to feed the leak with gas from pressurized tanks vs. using pressure suits to protect against lower cabin pressures, timeline and consumables constraints, re-entry considerations with leaks of unknown origin, and the impact the International Space Station (ISS) has had to the standard Space Shuttle cabin leak response philosophy. This last item in itself includes: procedural management differences, hardware considerations, additional capabilities due to the presence of the ISS and its resource, and ISS docking/undocking considerations with a

Effects of DeOrbitSail as applied to Lifetime predictions of Low Earth Orbit Satellites

Science.gov (United States)

Afful, Andoh; Opperman, Ben; Steyn, Herman

2016-07-01

Orbit lifetime prediction is an important component of satellite mission design and post-launch space operations. Throughout its lifetime in space, a spacecraft is exposed to risk of collision with orbital debris or operational satellites. This risk is especially high within the Low Earth Orbit (LEO) region where the highest density of space debris is accumulated. This paper investigates orbital decay of some LEO micro-satellites and accelerating orbit decay by using a deorbitsail. The Semi-Analytical Liu Theory (SALT) and the Satellite Toolkit was employed to determine the mean elements and expressions for the time rates of change. Test cases of observed decayed satellites (Iridium-85 and Starshine-1) are used to evaluate the predicted theory. Results for the test cases indicated that the theory fitted observational data well within acceptable limits. Orbit decay progress of the SUNSAT micro-satellite was analysed using relevant orbital parameters derived from historic Two Line Element (TLE) sets and comparing with decay and lifetime prediction models. This paper also explored the deorbit date and time for a 1U CubeSat (ZACUBE-01). The use of solar sails as devices to speed up the deorbiting of LEO satellites is considered. In a drag sail mode, the deorbitsail technique significantly increases the effective cross-sectional area of a satellite, subsequently increasing atmospheric drag and accelerating orbit decay. The concept proposed in this study introduced a very useful technique of orbit decay as well as deorbiting of spacecraft.

Second Stage (S-II) Plays Key Role in Apollo missions

Science.gov (United States)

1970-01-01

This photograph of the Saturn V Second Stage (S-II) clearly shows the cluster of five powerful J-2 engines needed to boost the Apollo spacecraft into earth orbit following first stage separation. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

Orbital Maneuvers for Spacecrafts Travelling to/from the Lagrangian Points

Science.gov (United States)

Bertachini, A.

The well-known Lagrangian points that appear in the planar restricted three-body problem (Szebehely, 1967) are very important for astronautical applications. They are five points of equilibrium in the equations of motion, what means that a particle located at one of those points with zero velocity will remain there indefinitely. The collinear points (L1, L2 and L3) are always unstable and the triangular points (L4 and L5) are stable in the present case studied (Sun-Earth system). They are all very good points to locate a space-station, since they require a small amount of V (and fuel), the control to be used for station-keeping. The triangular points are specially good for this purpose, since they are stable equilibrium points. In this paper, the planar restricted three-body problem is regularized (using Lemaître regularization) and combined with numerical integration and gradient methods to solve the two point boundary value problem (the Lambert's three-body problem). This combination is applied to the search of families of transfer orbits between the Lagrangian points and the Earth, in the Sun-Earth system, with the minimum possible cost of the control used. So, the final goal of this paper is to find the magnitude and direction of the two impulses to be applied in the spacecraft to complete the transfer: the first one when leaving/arriving at the Lagrangian point and the second one when arriving/living at the Earth. This paper is a continuation of two previous papers that studied transfers in the Earth-Moon system: Broucke (1979), that studied transfer orbits between the Lagrangian points and the Moon and Prado (1996), that studied transfer orbits between the Lagrangian points and the Earth. So, the equations of motion are: whereis the pseudo-potential given by: To solve the TPBVP in the regularized variables the following steps are used: i) Guess a initial velocity Vi, so together with the initial prescribed position ri the complete initial state is known; ii

Wuhan University Deep-space Orbit Determination and Gravity Recovery System(WUDOGS and Its Application Analysis

Directory of Open Access Journals (Sweden)

YE Mao

2017-03-01

Full Text Available WUDOGS(Wuhan University deep-space orbit determination and gravity recovery system is a software system designed for deep spacecraft precise orbit determination and planetary gravity recovery, developed independently at Wuhan University. WUDOGS now has the function for Lunar and Mars spacecraft precision orbit determination. Its design pattern and main function are briefly introduced. The cross verification test(CVT between WUDOGS and state of the art planetary precise orbit determination software GEODYN-Ⅱ are elaborated. The results show that:①for orbit propagation, with all the same forces and other configuration, the predicted orbit difference in R,T,N directions are less than 0.3 mm for one month arc, 5×10-3 mm for 2 days arc, compared with GEODYN-Ⅱ;②the difference RMS of computed values of observables for two-way range and two-way range rate is at levels of 0.06 mm and 0.002 mm/s respectively;③for Chinese Chang'E-1 POD, the reconstructed orbit difference between WUDOGS and GEODYN-Ⅱ is at 2 cm level, for ESA MEX POD, the reconstructed orbit difference between WUDOGS and ESA is at 25 m level. Current developing situation of WUDOGS and comparison with international research level show that WUDOGS has a good application prospect, which will be important for meeting the demand of Chinese future planetary exploration and the development of deep space spacecraft POD software.

Spacecraft Charge Monitor

Science.gov (United States)

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.

Spacecraft formation control using analytical finite-duration approaches

Science.gov (United States)

Ben Larbi, Mohamed Khalil; Stoll, Enrico

2018-03-01

This paper derives a control concept for formation flight (FF) applications assuming circular reference orbits. The paper focuses on a general impulsive control concept for FF which is then extended to the more realistic case of non-impulsive thrust maneuvers. The control concept uses a description of the FF in relative orbital elements (ROE) instead of the classical Cartesian description since the ROE provide a direct insight into key aspects of the relative motion and are particularly suitable for relative orbit control purposes and collision avoidance analysis. Although Gauss' variational equations have been first derived to offer a mathematical tool for processing orbit perturbations, they are suitable for several different applications. If the perturbation acceleration is due to a control thrust, Gauss' variational equations show the effect of such a control thrust on the Keplerian orbital elements. Integrating the Gauss' variational equations offers a direct relation between velocity increments in the local vertical local horizontal frame and the subsequent change of Keplerian orbital elements. For proximity operations, these equations can be generalized from describing the motion of single spacecraft to the description of the relative motion of two spacecraft. This will be shown for impulsive and finite-duration maneuvers. Based on that, an analytical tool to estimate the error induced through impulsive maneuver planning is presented. The resulting control schemes are simple and effective and thus also suitable for on-board implementation. Simulations show that the proposed concept improves the timing of the thrust maneuver executions and thus reduces the residual error of the formation control.

Compendium of Single Event Effects, Total Ionizing Dose, and Displacement Damage for Candidate Spacecraft Electronics for NASA

Science.gov (United States)

LaBel, Kenneth A.; OBryan, Martha V.; Chen, Dakai; Campola, Michael J.; Casey, Megan C.; Pellish, Jonathan A.; Lauenstein, Jean-Marie; Wilcox, Edward P.; Topper, Alyson D.; Ladbury, Raymond L.;

Internet Technology on Spacecraft

Science.gov (United States)

Rash, James; Parise, Ron; Hogie, Keith; Criscuolo, Ed; Langston, Jim; Powers, Edward I. (Technical Monitor)

2000-01-01

The Operating Missions as Nodes on the Internet (OMNI) project has shown that Internet technology works in space missions through a demonstration using the UoSAT-12 spacecraft. An Internet Protocol (IP) stack was installed on the orbiting UoSAT-12 spacecraft and tests were run to demonstrate Internet connectivity and measure performance. This also forms the basis for demonstrating subsequent scenarios. This approach provides capabilities heretofore either too expensive or simply not feasible such as reconfiguration on orbit. The OMNI project recognized the need to reduce the risk perceived by mission managers and did this with a multi-phase strategy. In the initial phase, the concepts were implemented in a prototype system that includes space similar components communicating over the TDRS (space network) and the terrestrial Internet. The demonstration system includes a simulated spacecraft with sample instruments. Over 25 demonstrations have been given to mission and project managers, National Aeronautics and Space Administration (NASA), Department of Defense (DoD), contractor technologists and other decisions makers, This initial phase reached a high point with an OMNI demonstration given from a booth at the Johnson Space Center (JSC) Inspection Day 99 exhibition. The proof to mission managers is provided during this second phase with year 2000 accomplishments: testing the use of Internet technologies onboard an actual spacecraft. This was done with a series of tests performed using the UoSAT-12 spacecraft. This spacecraft was reconfigured on orbit at very low cost. The total period between concept and the first tests was only 6 months! On board software was modified to add an IP stack to support basic IP communications. Also added was support for ping, traceroute and network timing protocol (NTP) tests. These tests show that basic Internet functionality can be used onboard spacecraft. The performance of data was measured to show no degradation from current

Current Single Event Effects Results for Candidate Spacecraft Electronics for NASA

Science.gov (United States)

OBryan, Martha V.; Seidleck, Christina M.; Carts, Martin A.; LaBel, Kenneth A.; Marshall, Cheryl J.; Reed, Robert A.; Sanders, Anthony B.; Hawkins, Donald K.; Cox, Stephen R.; Kniffin, Scott D.

2004-01-01

We present data on the vulnerability of a variety of candidate spacecraft electronics to proton and heavy ion induced single event effects. Devices tested include digital, analog, linear bipolar, and hybrid devices, among others.

Investigation of the McDonnell-Douglas orbiter and booster shuttle models in proximity at Mach numbers 2.0 to 6.0. Volume 7: Proximity data at Mach 4 and 6, interference free and launch vehicle data

Science.gov (United States)

Trimmer, L. L.; Love, D. A.; Decker, J. P.; Blackwell, K. L.; Strike, W. T.; Rampy, J. M.

1972-01-01

Aerodynamic data obtained from a space shuttle abort stage separation wind tunnel test are presented. The .00556 scale models of the orbiter and booster configuration were tested in close proximity using dual balances during the time period of April 21 to April 27 1971. Data were obtained for both booster and orbiter over an angle of attack range from -10 to 10 deg for zero degree sideslip angle. The models were tested at several relative incidence angles and separation distances and power conditions. Plug nozzles utilizing air were used to simulate booster and orbiter plumes at various altitudes along a nominal ascent trajectory. Powered conditions were 100, 50, 25 and 0 percent of full power for the orbiter and 100, 50 and 0 percent of full power for the booster. Pitch control effectiveness data were obtained for both booster and orbiter with power on and off. In addition, launch vehicle data with and without booster power were obtained utilizing a single balance in the booster model. Data were also obtained with the booster canard off in close proximity and for the launch configuration.

Retrieval of RTG'S in earth orbit

International Nuclear Information System (INIS)

Raab, B.; Frieder, M.A.; Skrabek, A.

1982-01-01

Since 1961, some ten Radioisotope Thermoelectric Generators (RTG's) have been placed into a variety of spacecraft which are now in earth orbit. All of these spacecraft are in orbits with lifetimes in excess of 100 years and pose no risk. However, since most of these spacecraft are no longer being actively used, these may be subject to an active removal program to reduce the population of objects in space. Therefore, a study was undertaken to evaluate the feasibility of retrieving or disposing of spacecraft with RTGs on board. Intervention scenarios are developed and an orbital rendezvous vehicle is conceptualized. The costs of RTG retrieval are derived and compared to the costs of RTG disposal, i.e., boost to a higher, multi-millenium-lifetime orbit, and are found to be not significantly different

Single stage buck-boost DC-AC neutral point clamped inverter

DEFF Research Database (Denmark)

Mo, Wei; Loh, Poh Chiang; Andrew, A.

2012-01-01

This paper proposes a new single stage buck-boost DC-AC neutral point clamped inverter topology which integrates the cascaded configurations of recently introduced inductor-capacitor-capacitor-transformer impedance source network (by Adamowicz) and classic NPC configuration. As a consequence......, it has enhanced buck-boost functionality and low output voltage distortions compared to the traditional Z-source inverter; it has continuous input current which reduces the source stress and inverter noise; it also contains two built-in capacitors which can block the DC current in the transformer...... windings thus preventing the core from saturation; lowers the voltage stresses and power losses of inverter switches and reduces the sizes of filtering devices and as well as obtains better output performance compared to the original two-level Z-source inverters. A phase disposition pulse width modulation...

SpaceX's Dragon America's next generation spacecraft

CERN Document Server

Seedhouse, Erik

2016-01-01

This book describes Dragon V2, a futuristic vehicle that not only provides a means for NASA to transport its astronauts to the orbiting outpost but also advances SpaceX’s core objective of reusability. A direct descendant of Dragon, Dragon V2 can be retrieved, refurbished and re-launched. It is a spacecraft with the potential to completely revolutionize the economics of an industry where equipment costing hundreds of millions of dollars is routinely discarded after a single use. It was presented by SpaceX CEO Elon Musk in May 2014 as the spaceship that will carry NASA astronauts to the International Space Station as soon as 2016 SpaceX’s Dragon – America’s Next Generation Spacecraft describes the extraordinary feats of engineering and human achievement that have placed this revolutionary spacecraft at the forefront of the launch industry and positioned it as the precursor for ultimately transporting humans to Mars. It describes the design and development of Dragon, provides mission highlights of the f...

Adaptive optics correction into single mode fiber for a low Earth orbiting space to ground optical communication link using the OPALS downlink.

Science.gov (United States)

Wright, Malcolm W; Morris, Jeffery F; Kovalik, Joseph M; Andrews, Kenneth S; Abrahamson, Matthew J; Biswas, Abhijit

2015-12-28

An adaptive optics (AO) testbed was integrated to the Optical PAyload for Lasercomm Science (OPALS) ground station telescope at the Optical Communications Telescope Laboratory (OCTL) as part of the free space laser communications experiment with the flight system on board the International Space Station (ISS). Atmospheric turbulence induced aberrations on the optical downlink were adaptively corrected during an overflight of the ISS so that the transmitted laser signal could be efficiently coupled into a single mode fiber continuously. A stable output Strehl ratio of around 0.6 was demonstrated along with the recovery of a 50 Mbps encoded high definition (HD) video transmission from the ISS at the output of the single mode fiber. This proof of concept demonstration validates multi-Gbps optical downlinks from fast slewing low-Earth orbiting (LEO) spacecraft to ground assets in a manner that potentially allows seamless space to ground connectivity for future high data-rates network.

Spatial analysis of galactic cosmic ray particles in low earth orbit/near equator orbit using SPENVIS

International Nuclear Information System (INIS)

Suparta, W; Zulkeple, S K

2014-01-01

The space environment has grown intensively harmful to spacecraft and astronauts. Galactic cosmic rays (GCRs) are one of the radiation sources that composed of high energetic particles originated from space and capable of damaging electronic systems through single event upset (SEU) process. In this paper, we analyzed GCR fluxes at different altitudes by using Space Environment Information System (SPENVIS) software and the results are compared to determine their intensities with respect to distance in the Earth's orbit. The altitudes are set at low earth orbit (400 km and 685 km), medium earth orbit (19,100 km and 20,200 km) and high earth orbit (35,793 km and 1,000,000 km). Then, within Low Earth Orbit (LEO) near the equator (NEqO), we used altitude of 685 km to compare GCRs with the intensities of solar particles and trapped particles in the radiation belt to determine the significance of GCRs in the orbit itself.

A combined DFT and restricted open-shell configuration interaction method including spin-orbit coupling: Application to transition metal L-edge X-ray absorption spectroscopy

Energy Technology Data Exchange (ETDEWEB)

Roemelt, Michael; Maganas, Dimitrios; Neese, Frank [Max-Planck Institute for Chemical Energy Conversion, Stiftstrasse 34-36, D-45470 Muelheim an der Ruhr (Germany); DeBeer, Serena [Max-Planck Institute for Chemical Energy Conversion, Stiftstrasse 34-36, D-45470 Muelheim an der Ruhr (Germany); Department of Chemistry and Chemical Biology, Cornell University, Ithaca, New York 14853 (United States)

2013-05-28

A novel restricted-open-shell configuration interaction with singles (ROCIS) approach for the calculation of transition metal L-edge X-ray absorption spectra is introduced. In this method, one first calculates the ground state and a number of excited states of the non-relativistic Hamiltonian. By construction, the total spin is a good quantum number in each of these states. For a ground state with total spin S excited states with spin S Prime = S, S - 1, and S + 1 are constructed. Using Wigner-Eckart algebra, all magnetic sublevels with M{sub S}= S, Horizontal-Ellipsis , -S for each multiplet of spin S are obtained. The spin-orbit operator is represented by a mean-field approximation to the full Breit-Pauli spin-orbit operator and is diagonalized over this N-particle basis. This is equivalent to a quasi-degenerate treatment of the spin-orbit interaction to all orders. Importantly, the excitation space spans all of the molecular multiplets that arise from the atomic Russell-Saunders terms. Hence, the method represents a rigorous first-principles approach to the complicated low-symmetry molecular multiplet problem met in L-edge X-ray absorption spectroscopy. In order to gain computational efficiency, as well as additional accuracy, the excitation space is restricted to single excitations and the configuration interaction matrix is slightly parameterized in order to account for dynamic correlation effects in an average way. To this end, it is advantageous to employ Kohn-Sham rather than Hartree-Fock orbitals thus defining the density functional theory/ROCIS method. However, the method can also be used in an entirely non-empirical fashion. Only three global empirical parameters are introduced and have been determined here for future application of the method to any system containing any transition metal. The three parameters were carefully calibrated using the L-edge X-ray absorption spectroscopy spectra of a test set of coordination complexes containing first row

One-impulse targeting strategy for longitudinal drift control of geosynchronous spacecraft subject to tesseral harmonics and luni-solar gravity perturbations

Science.gov (United States)

Kechichian, J. A.

1984-01-01

Kamel's (1973) East-West Stationkeeping Analysis is extended and an algorithm is presented that targets the geosynchronous spacecraft to the ideal initial conditions starting from any given relative longitude deviation within a given tolerance deadband in order to repeat the ideal longitudinal drift cycle that results in the longest possible period of time between maneuvers. The motion description takes into account the perturbations introduced by earth's tesseral harmonics and by the luni-solar gravity, assuming a near-circular orbit that requires only the control of orbital energy to repeat the ideal drift cycle via a single impulsive velocity change. The location of the maneuver along the orbit is such that the post-Delta-V eccentricity is always minimized.

Two-craft Coulomb formation study about circular orbits and libration points

Science.gov (United States)

Inampudi, Ravi Kishore

This dissertation investigates the dynamics and control of a two-craft Coulomb formation in circular orbits and at libration points; it addresses relative equilibria, stability and optimal reconfigurations of such formations. The relative equilibria of a two-craft tether formation connected by line-of-sight elastic forces moving in circular orbits and at libration points are investigated. In circular Earth orbits and Earth-Moon libration points, the radial, along-track, and orbit normal great circle equilibria conditions are found. An example of modeling the tether force using Coulomb force is discussed. Furthermore, the non-great-circle equilibria conditions for a two-spacecraft tether structure in circular Earth orbit and at collinear libration points are developed. Then the linearized dynamics and stability analysis of a 2-craft Coulomb formation at Earth-Moon libration points are studied. For orbit-radial equilibrium, Coulomb forces control the relative distance between the two satellites. The gravity gradient torques on the formation due to the two planets help stabilize the formation. Similar analysis is performed for along-track and orbit-normal relative equilibrium configurations. Where necessary, the craft use a hybrid thrusting-electrostatic actuation system. The two-craft dynamics at the libration points provide a general framework with circular Earth orbit dynamics forming a special case. In the presence of differential solar drag perturbations, a Lyapunov feedback controller is designed to stabilize a radial equilibrium, two-craft Coulomb formation at collinear libration points. The second part of the thesis investigates optimal reconfigurations of two-craft Coulomb formations in circular Earth orbits by applying nonlinear optimal control techniques. The objective of these reconfigurations is to maneuver the two-craft formation between two charged equilibria configurations. The reconfiguration of spacecraft is posed as an optimization problem using the

Validation of double Langmuir probe in-orbit performance onboard a nano-satellite

Science.gov (United States)

Tejumola, Taiwo Raphael; Zarate Segura, Guillermo Wenceslao; Kim, Sangkyun; Khan, Arifur; Cho, Mengu

2018-03-01

Many plasma measurement systems have been proposed and used onboard different satellites to characterize space plasma. Most of these systems employed the technique of Langmuir probes either using the single or double probes methods. Recent growth of lean satellites has positioned it on advantage to be used for space science missions using Langmuir probes because of its simplicity and convenience. However, single Langmuir probes are not appropriate to be used on lean satellites because of their limited conducting area which leads to spacecraft charging and drift of the instrument's electrical ground during measurement. Double Langmuir probes technique can overcome this limitation, as a measurement reference in relation to the spacecraft is not required. A double Langmuir probe measurement system was designed and developed at Kyushu Institute of Technology for HORYU-IV satellite, which is a 10 kg, 30 cm cubic class lean satellite launched into Low Earth Orbit on 17th February 2016. This paper presents the on-orbit performance and validation of the double Langmuir probe measurement using actual on-orbit measured data and computer simulations.

Cassini at Saturn Proximal Orbits - Attitude Control Challenges

Science.gov (United States)

Burk, Thomas A.

2013-01-01

The Cassini mission at Saturn will come to an end in the spring and summer of 2017 with a series of 22 orbits that will dip inside the rings of Saturn. These are called proximal orbits and will conclude with spacecraft disposal into the atmosphere of the ringed world on September 15, 2017. These unique orbits that cross the ring plane only a few thousand kilometers above the cloud tops of the planet present new attitude control challenges for the Cassini operations team. Crossing the ring plane so close to the inner edge of the rings means that the Cassini orientation during the crossing will be tailored to protect the sensitive electronics bus of the spacecraft. This orientation will put the sun sensors at some extra risk so this paper discusses how the team prepares for dust hazards. Periapsis is so close to the planet that spacecraft controllability with RCS thrusters needs to be evaluated because of the predicted atmospheric torque near closest approach to Saturn. Radiation during the ring plane crossings will likely trigger single event transients in some attitude control sensors. This paper discusses how the attitude control team deals with radiation hazards. The angular size and unique geometry of the rings and Saturn near periapsis means that star identification will be interrupted and this paper discusses how the safe mode attitude is selected to best deal with these large bright bodies during the proximal orbits.

Evaluation of planar 3D electrical capacitance tomography: from single-plane to dual-plane configuration

International Nuclear Information System (INIS)

Wei, Hsin-Yu; Qiu, Chang-Hua; Soleimani, Manuchehr

2015-01-01

Electrical capacitance tomography (ECT) is a non-invasive imaging technique that is sensitive to the dielectric permittivity property of an object. Conventional ECT systems have a circular/cylindrical or rectangular geometry, in which the electrode plates are usually spaced equally around the tank. It is the most common configuration as it can be easily applied to industrial pipelines. However, under some circumstances, the full access to the imaging geometry may not be applicable due to the limitation of the process area. In those cases, and with limited access, planar ECT sensors can fit the process structure if access to only one side is possible. A single-plane ECT configuration has been proposed for such applications. However, the planar array often suffers from a lack of sensitivity and difficulty with depth detection. To better understand these limitations we investigate the imaging performance from the single-plane ECT to dual-plane ECT structure. The limitations and constraints of the planar configuration will also be discussed. Several experiments were conducted using both single-plane and dual-plane configurations to evaluate the potential applications. The initial results are promising, and the quality of the reconstructed images are compared with the real condition for process validation. (paper)

Multiple spacecraft Michelson stellar interferometer

Science.gov (United States)

Stachnik, R. V.; Arnold, D.; Melroy, P.; Mccormack, E. F.; Gezari, D. Y.

1984-01-01

Results of an orbital analysis and performance assessment of SAMSI (Spacecraft Array for Michelson Spatial Interferometry) are presented. The device considered includes two one-meter telescopes in orbits which are identical except for slightly different inclinations; the telescopes achieve separations as large as 10 km and relay starlight to a central station which has a one-meter optical delay line in one interferometer arm. It is shown that a 1000-km altitude, zero mean inclination orbit affords natural scanning of the 10-km baseline with departures from optical pathlength equality which are well within the corrective capacity of the optical delay line. Electric propulsion is completely adequate to provide the required spacecraft motions, principally those needed for repointing. Resolution of 0.00001 arcsec and magnitude limits of 15 to 20 are achievable.

Navigating the MESSENGER Spacecraft through End of Mission

Science.gov (United States)

Bryan, C. G.; Williams, B. G.; Williams, K. E.; Taylor, A. H.; Carranza, E.; Page, B. R.; Stanbridge, D. R.; Mazarico, E.; Neumann, G. A.; O'Shaughnessy, D. J.; McAdams, J. V.; Calloway, A. B.

2015-12-01

The MErcury Surface, Space ENvironment, GEochemistry, and Ranging (MESSENGER) spacecraft orbited the planet Mercury from March 2011 until the end of April 2015, when it impacted the planetary surface after propellant reserves used to maintain the orbit were depleted. This highly successful mission was led by the principal investigator, Sean C. Solomon, of Columbia University. The Johns Hopkins University Applied Physics Laboratory (JHU/APL) designed and assembled the spacecraft and served as the home for spacecraft operations. Spacecraft navigation for the entirety of the mission was provided by the Space Navigation and Flight Dynamics Practice (SNAFD) of KinetX Aerospace. Orbit determination (OD) solutions were generated through processing of radiometric tracking data provided by NASA's Deep Space Network (DSN) using the MIRAGE suite of orbital analysis tools. The MESSENGER orbit was highly eccentric, with periapsis at a high northern latitude and periapsis altitude in the range 200-500 km for most of the orbital mission phase. In a low-altitude "hover campaign" during the final two months of the mission, periapsis altitudes were maintained within a narrow range between about 35 km and 5 km. Navigating a spacecraft so near a planetary surface presented special challenges. Tasks required to meet those challenges included the modeling and estimation of Mercury's gravity field and of solar and planetary radiation pressure, and the design of frequent orbit-correction maneuvers. Superior solar conjunction also presented observational modeling issues. One key to the overall success of the low-altitude hover campaign was a strategy to utilize data from an onboard laser altimeter as a cross-check on the navigation team's reconstructed and predicted estimates of periapsis altitude. Data obtained from the Mercury Laser Altimeter (MLA) on a daily basis provided near-real-time feedback that proved invaluable in evaluating alternative orbit estimation strategies, and

Combined effects of channel curvature and rotor configuration on the performance of two-stage viscous micropumps

Energy Technology Data Exchange (ETDEWEB)

Kang, Dong Jin [Yeungnam University, Gyeongsan (Korea, Republic of)

2017-06-15

The combined effects of channel curvature and rotor configuration on the performance of two-stage viscous micropumps were studied numerically. The Navier-Stokes equations were simulated to investigate the performance of two-stage micropumps. The performance of two-stage micropumps was studied in terms of the dimensionless mass flow rate and dimensionless driving power. Four different rotor configurations were designed by changing placement of two rotors inside a microchannel: Two aligned and two staggered configurations. The aligned rotor configuration of type 1 is to place the two rotors along the convex wall, while type 2 is to place them along the concave wall. Numerical results show that the rotor configuration plays a significant role in the performance of two-stage micropumps. The chan-nel curvature acts in a different way according to the rotor configuration. The mass flow rate of aligned rotor configuration of type 1 is greatly improved by the channel curvature, while it diminishes the mass flow rate of type 2. The maximum mass flow rate for the aligned rotor configuration of type 1 is obtained when the two rotors are placed at the junction of the circular and straight sections of the channel. The performance of staggered configurations is negligibly affected by the channel curvature. This characteristics is found due to rotation direction of the rotors. As the two rotors rotate in the opposite direction for the staggered configurations, the flow characteristics in the circular section is little affected by the channel curvature. The circumferential distance between the two rotors can be optimized in terms of the mass flow rate. The optimal value of the circumferential distance is about L = 1.4 for the staggered rotor configurations, and it is almost independent of the channel curvature. As the channel height increases, the circumferential distance becomes less significant for the staggered rotor configurations while it becomes significant for the aligned

Spin--orbit configuration-interaction study of valence and Rydberg states of LiBe

International Nuclear Information System (INIS)

Marino, M.M.; Ermler, W.C.; Kern, C.W.; Bondybey, V.E.

1992-01-01

Ab initio spin--orbit full configuration-interaction calculations in the context of relativistic effective core potentials are reported for the weakly bound metal dimer LiBe, a three-valence-electron system. The effects of basis set on the energies of valence and Rydberg states of the cluster are discussed, as are the effects of configuration space selection on the energy of the latter states. Results at the dissociative limit are compared to the experimental atomic spectra. Potential-energy curves and spectroscopic constants are presented for the ground state and fourteen excited states, which includes the Li and Be 2p valence states, the Li 3s, 3p, 3d, and 4s Rydberg states, as well as three low-lying states of the molecular cation

Manipulating localized molecular orbitals by single-atom contacts.

Science.gov (United States)

Wang, Weihua; Shi, Xingqiang; Lin, Chensheng; Zhang, Rui Qin; Minot, Christian; Van Hove, Michel A; Hong, Yuning; Tang, Ben Zhong; Lin, Nian

2010-09-17

We have fabricated atom-molecule contacts by attachment of single Cu atoms to terpyridine side groups of bis-terpyridine tetra-phenyl ethylene molecules on a Cu(111) surface. By means of scanning tunneling microscopy, spectroscopy, and density functional calculations, we have found that, due to the localization characteristics of molecular orbitals, the Cu-atom contact modifies the state localized at the terpyridine side group which is in contact with the Cu atom but does not affect the states localized at other parts of the molecule. These results illustrate the contact effects at individual orbitals and offer possibilities to manipulate orbital alignments within molecules.

Future earth orbit transportation systems/technology implications

Science.gov (United States)

Henry, B. Z.; Decker, J. P.

1976-01-01

Assuming Space Shuttle technology to be state-of-the-art, projected technological advances to improve the capabilities of single-stage-to-orbit (SSTO) derivatives are examined. An increase of about 30% in payload performance can be expected from upgrading the present Shuttle system through weight and drag reductions and improvements in the propellants and engines. The ODINEX (Optimal Design Integration Executive Computer Program) program has been used to explore design options. An advanced technology SSTO baseline system derived from ODINEX analysis has a conventional wing-body configuration using LOX/LH engines, three with two-position nozzles with expansion ratios of 40 and 200 and four with fixed nozzles with an expansion ratio of 40. Two assisted-takeoff approaches are under consideration in addition to a concept in which the orbital vehicle takes off empty using airbreathing propulsion and carries out a rendezvous with two large cryogenic tankers carrying propellant at an altitude of 6100 m. Further approaches under examination for propulsion, aerothermodynamic design, and design integration are described.

Attitude control for on-orbit servicing spacecraft using hybrid actuator

Science.gov (United States)

Wu, Yunhua; Han, Feng; Zheng, Mohong; He, Mengjie; Chen, Zhiming; Hua, Bing; Wang, Feng

2018-03-01

On-orbit servicing is one of the research hotspots of space missions. A small satellite equipped with multiple robotic manipulators is expected to carry out device replacement task for target large spacecraft. Attitude hyperstable control of a small satellite platform under rotations of the manipulators is a challenging problem. A hybrid momentum exchanging actuator consists of Control Moment Gyro (CMG) and Reaction Wheel (RW) is proposed to tackle the above issue, due to its huge amount of momentum storage capacity of the CMG and high control accuracy of the RW, in which the CMG produces large command torque while the RW offers additional control degrees. The constructed dynamic model of the servicing satellite advises that it's feasible for attitude hyperstable control of the platform with arbitrary manipulators through compensating the disturbance generated by rapid rotation of the manipulators. Then, null motion between the CMG and RW is exploited to drive the system to the expected target with favorable performance, and to overcome the CMG inherent geometric singularity and RW saturation. Simulations with different initial situations, including CMG hyperbolic and elliptic singularities and RW saturation, are executed. Compared to the scenarios where the CMG or RW fails stabilizing the platform, large control torque, precise control effect and escape of singularity are guaranteed by the introduced hybrid actuator, CMGRW (CMGRW refers to the hybrid momentum exchanging devices in this paper, consisting of 4 CMGs in classical pyramid cluster and 3 RWs in an orthogonal group (specific description can been found in Section 4)). The feasible performance of the satellite, CMG and RW under large disturbance demonstrates that the control architecture proposed is capable of attitude control for on-orbit servicing satellite with multiple robotic manipulators.

Artist concept of Galileo spacecraft

Science.gov (United States)

1988-01-01

Galileo spacecraft is illustrated in artist concept. Gallileo, named for the Italian astronomer, physicist and mathematician who is credited with construction of the first complete, practical telescope in 1620, will make detailed studies of Jupiter. A cooperative program with the Federal Republic of Germany the Galileo mission will amplify information acquired by two Voyager spacecraft in their brief flybys. Galileo is a two-element system that includes a Jupiter-orbiting observatory and an entry probe. Jet Propulsion Laboratory (JPL) is Galileo project manager and builder of the main spacecraft. Ames Research Center (ARC) has responsibility for the entry probe, which was built by Hughes Aircraft Company and General Electric. Galileo will be deployed from the payload bay (PLB) of Atlantis, Orbiter Vehicle (OV) 104, during mission STS-34.

Thermal elastic shock and its effect on TOPEX spacecraft attitude control

Science.gov (United States)

Zimbelman, Darrell F.

1991-01-01

Thermal elastic shock (TES) is a twice per orbit impulsive disturbance torque experienced by low-Earth orbiting spacecraft. The fundamental equations used to model the TES disturbance torque for typical spacecraft appendages (e.g., solar arrays and antenna booms) are derived in detail. In particular, the attitude-pointing performance of the TOPEX spacecraft, when subjected to the TES disturbance, is analyzed using a three-axis nonlinear time-domain simulation. Results indicate that the TOPEX spacecraft could exceed its roll-axis attitude-control requirement during penumbral transitions, and remain in violation for approximately 150 sec each orbit until the umbra collapses. A localized active-control system is proposed as a solution to minimize and/or eliminate the degrading effects of the TES disturbance.

Options for Staging Orbits in Cis-Lunar Space

Science.gov (United States)

Martinez, Roland; Whitley, Ryan

2016-01-01

NASA has been studying options to conduct missions beyond Low Earth Orbit, but within the Earth-Moon system, in preparation for deep space exploration including human missions to Mars. Referred to as the Proving Ground, this arena of exploration activities will enable the development of human spaceflight systems and operations to satisfy future exploration objectives beyond the cis-lunar environment. One option being considered includes the deployment of a habitable element or elements, which could be used as a central location for aggregation of supplies and resources for human missions in cis-lunar space and beyond. Characterizing candidate orbit locations for this asset and the impacts on system design and mission operations is important in the overall assessment of the options being considered. The orbits described in this paper were initially selected by taking advantage of previous studies conducted by NASA and the work of other authors. In this paper orbits are assessed for their relative attractiveness based on various factors. A set of constraints related to the capability of the combined Orion and SLS system to deliver humans and cargo to and from the orbit are evaluated. Deployed assets intended to spend multiple years in the Proving Ground would ideally require minimal station keeping costs to reduce the mass budget allocated to this function. Additional mission design drivers include eclipse frequency, potential for uninterrupted communication with deployed assets, thermal, attitude control, communications, and other operational implications. Also the ability to support potential lunar surface activities and excursion missions beyond Earth-Moon space is considered. The results of the characterization and evaluation of the selected orbits indicate a Near Rectilinear Orbit (NRO) is an attractive candidate as an aggregation point or staging location for operations. In this paper, the NRO is further described in terms which balance a number of key

Analysis of the half-projected Hartree--Fock function: density matrix, natural orbitals, and configuration interaction equivalence

International Nuclear Information System (INIS)

Smeyers, Y.G.; Delgado-Barrio, G.

1976-01-01

The half-projected Hartree--Fock function for singlet states (HPHF) is analyzed in terms of natural electronic configurations. For this purpose the HPHF spinless density matrix and its natural orbitals are first deduced. It is found that the HPHF function does not contain any contribution from odd-times excited configurations. It is seen in addition, in the case of the singlet ground states, this function is approximately equivalent to two closed-shell configurations, although the nature of the excited one depends on the nuclear geometry. An example is given in the case of the LiH ground state. Finally, the application of this model for studying systems of more than two atoms is criticized

NASA Earth-to-Orbit Engineering Design Challenges: Thermal Protection Systems

Science.gov (United States)

National Aeronautics and Space Administration (NASA), 2010

2010-01-01

National Aeronautics and Space Administration (NASA) Engineers at Marshall Space Flight Center, Dryden Flight Research Center, and their partners at other NASA centers and in private industry are currently developing X-33, a prototype to test technologies for the next generation of space transportation. This single-stage-to-orbit reusable launch…

Distant retrograde orbits and the asteroid hazard

Science.gov (United States)

Perozzi, Ettore; Ceccaroni, Marta; Valsecchi, Giovanni B.; Rossi, Alessandro

2017-08-01

Distant Retrograde Orbits (DROs) gained a novel wave of fame in space mission design because of their numerous advantages within the framework of the US plans for bringing a large asteroid sample in the vicinity of the Earth as the next target for human exploration. DROs are stable solutions of the three-body problem that can be used whenever an object, whether of natural or artificial nature, is required to remain in the neighborhood of a celestial body without being gravitationally captured by it. As such, they represent an alternative option to Halo orbits around the collinear Lagrangian points L1 and L2. Also known under other names ( e.g., quasi-satellite orbits, cis-lunar orbits, family- f orbits) these orbital configurations found interesting applications in several mission profiles, like that of a spacecraft orbiting around the small irregularly shaped satellite of Mars Phobos or the large Jovian moon Europa. In this paper a basic explanation of the DRO dynamics is presented in order to clarify some geometrical properties that characterize them. Their accessibility is then discussed from the point of view of mission analysis under different assumptions. Finally, their relevance within the framework of the present asteroid hazard protection programs is shown, stressing the significant increase in warning time they would provide in the prediction of impactors coming from the direction of the Sun.

Galileo spacecraft power management and distribution system

International Nuclear Information System (INIS)

Detwiler, R.C.; Smith, R.L.

1990-01-01

It has been twelve years since two Voyager spacecraft began the direct route to the outer planets. In October 1989 a single Galileo spacecraft started the return to Jupiter. Conceived as a simple Voyager look-alike, the Galileo power management and distribution (PMAD) system has undergone many iterations in configuration. Major changes to the PMAD resulted from dual spun slip ring limitations, variations in launch vehicle thrust capabilities, and launch delays. Lack of an adequate launch vehicle for an interplanetary mission of Galileo's size has resulted in an extremely long flight duration. A Venus-Earth-Earth Gravity Assist (VEEGA) tour, vital to attain the required energy, results in a 6 year trip to Jupiter and its moons. This paper provides a description of the Galileo PMAD and documents the design drivers that established the final as-built hardware

Investigating fundamental physics and space environment with a dedicated Earth-orbiting spacecraft

Science.gov (United States)

Peron, Roberto

-year requirement and thus they need specific arrangements for deorbiting at the end of life or they can simply rely on mother nature for reentry. The goal of this proposed approach is to utilize existing technology developed for acceleration measurement in space and state-of-the-art satellite tracking to precisely determine the orbit of a satellite with well-defined geometrical and mass characteristics (i.e., (A/m) ratio), at the same time accurately measuring over a long period of time the drag deceleration (as well as others non-gravitational effects) acting on the satellite. This will result in a virtually drag-free object that can be exploited to: 1. perform fundamental physics tests by verifying the equation of motion of a test mass in the general relativistic context and placing limits to alternative theories of gravitation; 2. improve the knowledge of selected tidal terms; 3. map, through acceleration measurements, the atmospheric density in the orbital region of interest. In its preliminary incarnation, the satellite would be cylindrical in shape and spinning about its cylinder axis that would be also orthogonal to the orbital plane. The satellite should be placed on a dawn-dusk, sun-synchronous, elliptical orbit spanning the orbital altitudes of interest (e.g., between 500 and 1200 km of altitude). The satellite should be equipped with a 3-axis accelerometer package with an acceleration resolution better than (10^{-11} g) (with (g) the acceleration at the Earth's surface). The expected measurement range is (10^{-8} - 10^{-11} g) considering estimates of drag forces at minimum and maximum solar activity conditions in the altitude range of interest and a preliminary estimate of the satellite (A/m) ratio. The overall concept of the mission will be discussed, concentrating on the fundamental aspects and main scientific return. The main instrumentation to be hosted on-board the spacecraft will be then reviewed, with a focus on current and projected capabilities.

Summary of the Orbiter mechanical systems

Science.gov (United States)

Kiker, J.; Hinson, K.

1979-01-01

Major mechanical systems of the Orbiter space vehicle are summarized with respect to general design details, manner of operation, expected performance, and, where applicable, unique features. A synopsis of data obtained during the five atmospheric flight tests of spacecraft OV-101 and status of the systems for the first orbital spacecraft STS-1 are presented.

Definition of the topological structure of the automatic control system of spacecrafts

International Nuclear Information System (INIS)

KrasnoyarskiyRabochiy prospect, Krasnoyarsk, 660014 (Russian Federation))" data-affiliation=" (Siberian State Aerospace University named after Academician M.F.Reshetnev 31 KrasnoyarskiyRabochiy prospect, Krasnoyarsk, 660014 (Russian Federation))" >Zelenkov, P V; KrasnoyarskiyRabochiy prospect, Krasnoyarsk, 660014 (Russian Federation))" data-affiliation=" (Siberian State Aerospace University named after Academician M.F.Reshetnev 31 KrasnoyarskiyRabochiy prospect, Krasnoyarsk, 660014 (Russian Federation))" >Karaseva, M V; KrasnoyarskiyRabochiy prospect, Krasnoyarsk, 660014 (Russian Federation))" data-affiliation=" (Siberian State Aerospace University named after Academician M.F.Reshetnev 31 KrasnoyarskiyRabochiy prospect, Krasnoyarsk, 660014 (Russian Federation))" >Tsareva, E A; Tsarev, R Y

2015-01-01

The paper considers the problem of selection the topological structure of the automated control system of spacecrafts. The integer linear model of mathematical programming designed to define the optimal topological structure for spacecraft control is proposed. To solve the determination problem of topological structure of the control system of spacecrafts developed the procedure of the directed search of some structure variants according to the scheme 'Branch and bound'. The example of the automated control system of spacecraft development included the combination of ground control stations, managing the spacecraft of three classes with a geosynchronous orbit with constant orbital periods is presented

Some New Locally Optimal Control Laws for Sailcraft Dynamics in Heliocentric Orbits

Directory of Open Access Journals (Sweden)

F. A. Abd El-Salam

2013-01-01

Full Text Available The concept of solar sailing and its developing spacecraft is presented. The gravitational and solar radiation forces are considered. The effect of source of radiation pressure and the force due to coronal mass ejections and solar wind on the sailcraft configurations is modeled. Some analytical control laws with some mentioned input constraints for optimizing sailcraft dynamics in heliocentric orbit using lagrange’s planetary equations are obtained. Optimum force vector in a required direction is maximized by deriving optimal sail cone angle. Ignoring the absorbed and diffusely reflected parts of the radiation, some special cases are obtained. New control laws that maximize thrust to obtain certain required maximization in some particular orbital element are obtained.

Charge transfer dynamics from adsorbates to surfaces with single active electron and configuration interaction based approaches

Energy Technology Data Exchange (ETDEWEB)

Ramakrishnan, Raghunathan, E-mail: r.ramakrishnan@unibas.ch [Institute of Physical Chemistry, National Center for Computational Design and Discovery of Novel Materials (MARVEL), Department of Chemistry, University of Basel, Klingelbergstrasse 80, CH-4056 Basel (Switzerland); Nest, Mathias [Theoretische Chemie, Technische Universität München, Lichtenbergstr. 4, 85747 Garching (Germany)

2015-01-13

Highlights: • We model electron dynamics across cyano alkanethiolates attached to gold cluster. • We present electron transfer time scales from TD-DFT and TD-CI based simulations. • Both DFT and CI methods qualitatively predict the trend in time scales. • TD-CI predicts the experimental relative time scale very accurately. - Abstract: We employ wavepacket simulations based on many-body time-dependent configuration interaction (TD-CI), and single active electron theories, to predict the ultrafast molecule/metal electron transfer time scales, in cyano alkanethiolates bonded to model gold clusters. The initial states represent two excited states where a valence electron is promoted to one of the two virtual π{sup ∗} molecular orbitals localized on the cyanide fragment. The ratio of the two time scales indicate the efficiency of one charge transfer channel over the other. In both our one-and many-electron simulations, this ratio agree qualitatively with each other as well as with the previously reported experimental time scales (Blobner et al., 2012), measured for a macroscopic metal surface. We study the effect of cluster size and the description of electron correlation on the charge transfer process.

Orbit-attitude coupled motion around small bodies: Sun-synchronous orbits with Sun-tracking attitude motion

Science.gov (United States)

Kikuchi, Shota; Howell, Kathleen C.; Tsuda, Yuichi; Kawaguchi, Jun'ichiro

2017-11-01

The motion of a spacecraft in proximity to a small body is significantly perturbed due to its irregular gravity field and solar radiation pressure. In such a strongly perturbed environment, the coupling effect of the orbital and attitude motions exerts a large influence that cannot be neglected. However, natural orbit-attitude coupled dynamics around small bodies that are stationary in both orbital and attitude motions have yet to be observed. The present study therefore investigates natural coupled motion that involves both a Sun-synchronous orbit and Sun-tracking attitude motion. This orbit-attitude coupled motion enables a spacecraft to maintain its orbital geometry and attitude state with respect to the Sun without requiring active control. Therefore, the proposed method can reduce the use of an orbit and attitude control system. This paper first presents analytical conditions to achieve Sun-synchronous orbits and Sun-tracking attitude motion. These analytical solutions are then numerically propagated based on non-linear coupled orbit-attitude equations of motion. Consequently, the possibility of implementing Sun-synchronous orbits with Sun-tracking attitude motion is demonstrated.

Lunar Gravity-Assist Maneuver As a Way of Reducing the Orbit Amplitude in the Spectrum-Röntgen-Gamma Project

Science.gov (United States)

Kovalenko, I. D.; Eismont, N. A.

2018-04-01

Spectrum-Röntgen-Gamma (SRG) is a space observatory designed to observe astrophysical objects in the X-ray range of the electromagnetic spectrum. SRG is planned to be launched in 2019 by a Proton-M launch vehicle with a DM3 upper stage. The spacecraft will be delivered to an orbit around the Sun-Earth collinear libration point L2 located at a distance of 1.5 million km from the Earth. Although the SRG launch scheme has already been determined at present, in this paper we consider an alternative spacecraft transfer scenario using a lunar gravity-assist maneuver. The proposed scenario allows a oneimpulse transfer from a low Earth orbit to a small-amplitude orbit around the libration point to be performed while fulfilling the technical constraints and the scientific requirements of the mission.

The analysis of single-electron orbits in a free electron laser based upon a rectangular hybrid wiggler

International Nuclear Information System (INIS)

Kordbacheh, A.; Ghahremaninezhad, Roghayeh; Maraghechi, B.

2012-01-01

A three-dimensional analysis of a novel free-electron laser (FEL) based upon a rectangular hybrid wiggler (RHW) is presented. This RHW is designed in a configuration composed of rectangular rings with alternating ferrite and dielectric spacers immersed in a solenoidal magnetic field. An analytic model of RHW is introduced by solution of Laplace's equation for the magnetostatic fields under the appropriate boundary conditions. The single-electron orbits in combined RHW and axial guide magnetic fields are studied when only the first and the third spatial harmonic components of the RHW field are taken into account and the higher order terms are ignored. The results indicate that the third spatial harmonic leads to group III orbits with a strong negative mass regime particularly in large solenoidal magnetic fields. RHW is found to be a promising candidate with favorable characteristics to be used in microwave FEL.

The analysis of single-electron orbits in a free electron laser based upon a rectangular hybrid wiggler

Science.gov (United States)

Kordbacheh, A.; Ghahremaninezhad, Roghayeh; Maraghechi, B.

2012-09-01

A three-dimensional analysis of a novel free-electron laser (FEL) based upon a rectangular hybrid wiggler (RHW) is presented. This RHW is designed in a configuration composed of rectangular rings with alternating ferrite and dielectric spacers immersed in a solenoidal magnetic field. An analytic model of RHW is introduced by solution of Laplace's equation for the magnetostatic fields under the appropriate boundary conditions. The single-electron orbits in combined RHW and axial guide magnetic fields are studied when only the first and the third spatial harmonic components of the RHW field are taken into account and the higher order terms are ignored. The results indicate that the third spatial harmonic leads to group III orbits with a strong negative mass regime particularly in large solenoidal magnetic fields. RHW is found to be a promising candidate with favorable characteristics to be used in microwave FEL.

Evaluation of bearing configurations using the single bearing tester in liquid nitrogen

Science.gov (United States)

Jett, T.; Hall, P.; Thom, R.

1991-01-01

Various bearing configurations were tested using the Marshall Space Flight Center single bearing tester with LN2 as the cryogenic coolant. The baseline was one Rocketdyne phase one high pressure oxidizer turbopump (HPOTP) pump end 45-mm bore bearing. The bearing configurations that were tested included a Salox/M cage configuration, a silicon nitride ball configuration, an elongated cage configuration, and a Bray 601 grease configuration.

Modes of uncontrolled rotational motion of the Progress M-29M spacecraft

Science.gov (United States)

Belyaev, M. Yu.; Matveeva, T. V.; Monakhov, M. I.; Rulev, D. N.; Sazonov, V. V.

2018-01-01

We have reconstructed the uncontrolled rotational motion of the Progress M-29M transport cargo spacecraft in the single-axis solar orientation mode (the so-called sunward spin) and in the mode of the gravitational orientation of a rotating satellite. The modes were implemented on April 3-7, 2016 as a part of preparation for experiments with the DAKON convection sensor onboard the Progress spacecraft. The reconstruction was performed by integral statistical techniques using the measurements of the spacecraft's angular velocity and electric current from its solar arrays. The measurement data obtained in a certain time interval have been jointly processed using the least-squares method by integrating the equations of the spacecraft's motion relative to the center of mass. As a result of processing, the initial conditions of motion and parameters of the mathematical model have been estimated. The motion in the sunward spin mode is the rotation of the spacecraft with an angular velocity of 2.2 deg/s about the normal to the plane of solar arrays; the normal is oriented toward the Sun or forms a small angle with this direction. The duration of the mode is several orbit passes. The reconstruction has been performed over time intervals of up to 1 h. As a result, the actual rotational motion of the spacecraft relative to the Earth-Sun direction was obtained. In the gravitational orientation mode, the spacecraft was rotated about its longitudinal axis with an angular velocity of 0.1-0.2 deg/s; the longitudinal axis executed small oscillated relative to the local vertical. The reconstruction of motion relative to the orbital coordinate system was performed in time intervals of up to 7 h using only the angularvelocity measurements. The measurements of the electric current from solar arrays were used for verification.

Nuclear reactor power as applied to a space-based radar mission

Science.gov (United States)

Jaffe, L.; Beatty, R.; Bhandari, P.; Chow, E.; Deininger, W.; Ewell, R.; Fujita, T.; Grossman, M.; Bloomfield, H.; Heller, J.

1988-01-01

A space-based radar mission and spacecraft are examined to determine system requirements for a 300 kWe space nuclear reactor power system. The spacecraft configuration and its orbit, launch vehicle, and propulsion are described. Mission profiles are addressed, and storage in assembly orbit is considered. Dynamics and attitude control and the problems of nuclear and thermal radiation are examined.

IUS/TUG orbital operations and mission support study. Volume 4: Project planning data

Science.gov (United States)

1975-01-01

Planning data are presented for the development phases of interim upper stage (IUS) and tug systems. Major project planning requirements, major event schedules, milestones, system development and operations process networks, and relevant support research and technology requirements are included. Topics discussed include: IUS flight software; tug flight software; IUS/tug ground control center facilities, personnel, data systems, software, and equipment; IUS mission events; tug mission events; tug/spacecraft rendezvous and docking; tug/orbiter operations interface, and IUS/orbiter operations interface.

Effects of Space Weather on Geosynchronous Electromagnetic Spacecraft Perturbations Using Statistical Fluxes

Science.gov (United States)

Hughes, J.; Schaub, H.

2017-12-01

Spacecraft can charge to very negative voltages at GEO due to interactions with the space plasma. This can cause arcing which can damage spacecraft electronics or solar panels. Recently, it has been suggested that spacecraft charging may lead to orbital perturbations which change the orbits of lightweight uncontrolled debris orbits significantly. The motions of High Area to Mass Ratio objects are not well explained with just perturbations from Solar Radiation Pressure (SRP) and earth, moon, and sun gravity. A charged spacecraft will experience a Lorentz force as the spacecraft moves relative to Earth's magnetic field, as well as a Lorentz torque and eddy current torques if the object is rotating. Prior work assuming a constant "worst case" voltage has shown that Lorentz and eddy torques can cause quite large orbital changes by rotating the object to experience more or less SRP. For some objects, including or neglecting these electromagnetic torques can lead to differences of thousands of kilometers after only two orbits. This paper will further investigate the effects of electromagnetic perturbations by using a charging model that uses measured flux distributions to better simulate natural charging. This differs from prior work which used a constant voltage or Maxwellian distributions. This is done to a calm space weather case of Kp = 2 and a stormy case where Kp = 8. Preliminary analysis suggests that electrostatics will still cause large orbital changes even with the more realistic charging model.

Cassini Spacecraft In-Flight Swap to Backup Attitude Control Thrusters

Science.gov (United States)

Bates, David M.

2010-01-01

NASA's Cassini Spacecraft, launched on October 15th, 1997 and arrived at Saturn on June 30th, 2004, is the largest and most ambitious interplanetary spacecraft in history. In order to meet the challenging attitude control and navigation requirements of the orbit profile at Saturn, Cassini is equipped with a monopropellant thruster based Reaction Control System (RCS), a bipropellant Main Engine Assembly (MEA) and a Reaction Wheel Assembly (RWA). In 2008, after 11 years of reliable service, several RCS thrusters began to show signs of end of life degradation, which led the operations team to successfully perform the swap to the backup RCS system, the details and challenges of which are described in this paper. With some modifications, it is hoped that similar techniques and design strategies could be used to benefit other spacecraft.

Space Situational Awareness of Large Numbers of Payloads From a Single Deployment

Science.gov (United States)

Segerman, A.; Byers, J.; Emmert, J.; Nicholas, A.

2014-09-01

The nearly simultaneous deployment of a large number of payloads from a single vehicle presents a new challenge for space object catalog maintenance and space situational awareness (SSA). Following two cubesat deployments last November, it took five weeks to catalog the resulting 64 orbits. The upcoming Kicksat mission will present an even greater SSA challenge, with its deployment of 128 chip-sized picosats. Although all of these deployments are in short-lived orbits, future deployments will inevitably occur at higher altitudes, with a longer term threat of collision with active spacecraft. With such deployments, individual scientific payload operators require rapid precise knowledge of their satellites' locations. Following the first November launch, the cataloguing did not initially associate a payload with each orbit, leaving this to the satellite operators. For short duration missions, the time required to identify an experiment's specific orbit may easily be a large fraction of the spacecraft's lifetime. For a Kicksat-type deployment, present tracking cannot collect enough observations to catalog each small object. The current approach is to treat the chip cloud as a single catalog object. However, the cloud dissipates into multiple subclouds and, ultimately, tiny groups of untrackable chips. One response to this challenge may be to mandate installation of a transponder on each spacecraft. Directional transponder transmission detections could be used as angle observations for orbit cataloguing. Of course, such an approach would only be employable with cooperative spacecraft. In other cases, a probabilistic association approach may be useful, with the goal being to establish the probability of an element being at a given point in space. This would permit more reliable assessment of the probability of collision of active spacecraft with any cloud element. This paper surveys the cataloguing challenges presented by large scale deployments of small spacecraft

Manifold dynamics in the Earth-Moon system via isomorphic mapping with application to spacecraft end-of-life strategies

Science.gov (United States)

Pontani, Mauro; Giancotti, Marco; Teofilatto, Paolo

2014-12-01

Recently, manifold dynamics has assumed an increasing relevance for analysis and design of low-energy missions, both in the Earth-Moon system and in alternative multibody environments. With regard to lunar missions, exterior and interior transfers, based on the transit through the regions where the collinear libration points L1 and L2 are located, have been studied for a long time and some space missions have already taken advantage of the results of these studies. This paper is focused on the definition and use of a special isomorphic mapping for low-energy mission analysis. A convenient set of cylindrical coordinates is employed to describe the spacecraft dynamics (i.e. position and velocity), in the context of the circular restricted three-body problem, used to model the spacecraft motion in the Earth-Moon system. This isomorphic mapping of trajectories allows the identification and intuitive representation of periodic orbits and of the related invariant manifolds, which correspond to tubes that emanate from the curve associated with the periodic orbit. Heteroclinic connections, i.e. the trajectories that belong to both the stable and the unstable manifolds of two distinct periodic orbits, can be easily detected by means of this representation. This paper illustrates the use of isomorphic mapping for finding (a) periodic orbits, (b) heteroclinic connections between trajectories emanating from two Lyapunov orbits, the first at L1, and the second at L2, and (c) heteroclinic connections between trajectories emanating from the Lyapunov orbit at L1 and from a particular unstable lunar orbit. Heteroclinic trajectories are asymptotic trajectories that travels at zero-propellant cost. In practical situations, a modest delta-v budget is required to perform transfers along the manifolds. This circumstance implies the possibility of performing complex missions, by combining different types of trajectory arcs belonging to the manifolds. This work studies also the possible

Thermal analysis of hybrid single-phase, two-phase and heat pump thermal control system (TCS) for future spacecraft

International Nuclear Information System (INIS)

Lee, S.H.; Mudawar, I.; Hasan, Mohammad M.

2016-01-01

Highlights: • Hybrid Thermal Control System (H-TCS) is proposed for future spacecraft. • Thermodynamic performance of H-TCS is examined for different space missions. • Operational modes including single-phase, two-phase and heat pump are explored. • R134a is deemed most appropriate working fluid. - Abstract: An urgent need presently exists to develop a new class of versatile spacecraft capable of conducting different types of missions and enduring varying gravitational and temperature environments, including Lunar, Martian and Near Earth Object (NEOs). This study concerns the spacecraft's Thermal Control System (TCS), which tackles heat acquisition, especially from crew and avionics, heat transport, and ultimate heat rejection by radiation. The primary goal of the study is to explore the design and thermal performance of a Hybrid Thermal Control System (H-TCS) that would satisfy the diverse thermal requirements of the different space missions. The H-TCS must endure both ‘cold’ and ‘hot’ environments, reduce weight and size, and enhance thermodynamic performance. Four different operational modes are considered: single-phase, two-phase, basic heat pump and heat pump with liquid-side, suction-side heat exchanger. A thermodynamic trade study is conducted for six different working fluids to assess important performance parameters including mass flow rate of the working fluid, maximum pressure, radiator area, compressor/pump work, and coefficient of performance (COP). R134a is determined to be most suitable based on its ability to provide a balanced compromise between reducing flow rate and maintaining low system pressure, and a moderate coefficient of performance (COP); this fluid is also both nontoxic and nonflammable, and features zero ozone depletion potential (ODP) and low global warming potential (GWP). It is shown how specific mission stages dictate which mode of operation is most suitable, and this information is used to size the radiator for the

Testing programs for the Multimission Modular Spacecraft

Science.gov (United States)

Greenwell, T. J.

1978-01-01

The Multimission Modular Spacecraft (MMS) provides a standard spacecraft bus to a user for a variety of space missions ranging from near-earth to synchronous orbits. The present paper describes the philosophy behind the MMS module test program and discusses the implementation of the test program. It is concluded that the MMS module test program provides an effective and comprehensive customer buy-off at the subsystem contractor's plant, is an optimum approach for checkout of the subsystems prior to use for on-orbit servicing in the Shuttle Cargo Bay, and is a cost-effective technique for environmental testing.

Sorting photons of different rotational Doppler shifts (RDS) by orbital angular momentum of single-photon with spin-orbit-RDS entanglement.

Science.gov (United States)

Chen, Lixiang; She, Weilong

2008-09-15

We demonstrate that single photons from a rotating q-plate exhibit an entanglement in three degrees of freedom of spin, orbital angular momentum, and the rotational Doppler shift (RDS) due to the nonconservation of total spin and orbital angular momenta. We find that the rotational Doppler shift deltaomega = Omega((delta)s + deltal) , where s, l and Omega are quantum numbers of spin, orbital angular momentum, and rotating velocity of the q-plate, respectively. Of interest is that the rotational Doppler shift directly reflects the rotational symmetry of q-plates and can be also expressed as deltaomega = (Omega)n , where n = 2(q-1) denotes the fold number of rotational symmetry. Besides, based on this single-photon spin-orbit-RDS entanglement, we propose an experimental scheme to sort photons of different frequency shifts according to individual orbital angular momentum.

Viking orbiter and its Mariner inheritance

Science.gov (United States)

1984-01-01

Improvements to the design of the Mariner spacecraft resulted in the Viking spacecraft. The Viking spacecraft would consist of two major systems - an orbiter and a lander, while the lander would provide the means for safely delivering the scientific instruments to the surface, house, and provide the necessary power source and communication links for those experiments, the orbiter would transport the lander to Mars, rovide a platform for the Viking imaging system so that proposed landing sites could be surveyed and certified, relay lander science information back to Earth, and conduct scientific observations in its own right.

Application of the NASCAP Spacecraft Simulation Tool to Investigate Electrodynamic Tether Current Collection in LEO

Science.gov (United States)

Adams, Mitzi; HabashKrause, Linda

2012-01-01

Recent interest in using electrodynamic tethers (EDTs) for orbital maneuvering in Low Earth Orbit (LEO) has prompted the development of the Marshall ElectroDynamic Tether Orbit Propagator (MEDTOP) model. The model is comprised of several modules which address various aspects of EDT propulsion, including calculation of state vectors using a standard orbit propagator (e.g., J2), an atmospheric drag model, realistic ionospheric and magnetic field models, space weather effects, and tether librations. The natural electromotive force (EMF) attained during a radially-aligned conductive tether results in electrons flowing down the tether and accumulating on the lower-altitude spacecraft. The energy that drives this EMF is sourced from the orbital energy of the system; thus, EDTs are often proposed as de-orbiting systems. However, when the current is reversed using satellite charged particle sources, then propulsion is possible. One of the most difficult challenges of the modeling effort is to ascertain the equivalent circuit between the spacecraft and the ionospheric plasma. The present study investigates the use of the NASA Charging Analyzer Program (NASCAP) to calculate currents to and from the tethered satellites and the ionospheric plasma. NASCAP is a sophisticated set of computational tools to model the surface charging of three-dimensional (3D) spacecraft surfaces in a time-varying space environment. The model's surface is tessellated into a collection of facets, and NASCAP calculates currents and potentials for each one. Additionally, NASCAP provides for the construction of one or more nested grids to calculate space potential and time-varying electric fields. This provides for the capability to track individual particles orbits, to model charged particle wakes, and to incorporate external charged particle sources. With this study, we have developed a model of calculating currents incident onto an electrodynamic tethered satellite system, and first results are shown

Single photon laser altimeter simulator and statistical signal processing

Science.gov (United States)

Vacek, Michael; Prochazka, Ivan

2013-05-01

Spaceborne altimeters are common instruments onboard the deep space rendezvous spacecrafts. They provide range and topographic measurements critical in spacecraft navigation. Simultaneously, the receiver part may be utilized for Earth-to-satellite link, one way time transfer, and precise optical radiometry. The main advantage of single photon counting approach is the ability of processing signals with very low signal-to-noise ratio eliminating the need of large telescopes and high power laser source. Extremely small, rugged and compact microchip lasers can be employed. The major limiting factor, on the other hand, is the acquisition time needed to gather sufficient volume of data in repetitive measurements in order to process and evaluate the data appropriately. Statistical signal processing is adopted to detect signals with average strength much lower than one photon per measurement. A comprehensive simulator design and range signal processing algorithm are presented to identify a mission specific altimeter configuration. Typical mission scenarios (celestial body surface landing and topographical mapping) are simulated and evaluated. The high interest and promising single photon altimeter applications are low-orbit (˜10 km) and low-radial velocity (several m/s) topographical mapping (asteroids, Phobos and Deimos) and landing altimetry (˜10 km) where range evaluation repetition rates of ˜100 Hz and 0.1 m precision may be achieved. Moon landing and asteroid Itokawa topographical mapping scenario simulations are discussed in more detail.

The first preliminary experiments on an 84 GHz gyrotron with a single-stage depressed collector

International Nuclear Information System (INIS)

Shimozuma, T.; Sato, M.; Takita, Y.

1997-10-01

We fabricated and tested an 84GHz gyrotron with a single-stage depressed collector. The gyrotron has a high-voltage insulating section made of a low loss silicon nitride composite. In this preliminary experiment in the depressed collector configuration, we obtained 591kW, 41% operation with a depression voltage of 22.5kV. Access to the higher efficiency region was inhibited by an increase in anode current. (author)

The Coupled Orbit-Attitude Dynamics and Control of Electric Sail in Displaced Solar Orbits

Directory of Open Access Journals (Sweden)

Mingying Huo

2017-01-01

Full Text Available Displaced solar orbits for spacecraft propelled by electric sails are investigated. Since the propulsive thrust is induced by the sail attitude, the orbital and attitude dynamics of electric-sail-based spacecraft are coupled and required to be investigated together. However, the coupled dynamics and control of electric sails have not been discussed in most published literatures. In this paper, the equilibrium point of the coupled dynamical system in displaced orbit is obtained, and its stability is analyzed through a linearization. The results of stability analysis show that only some of the orbits are marginally stable. For unstable displaced orbits, linear quadratic regulator is employed to control the coupled attitude-orbit system. Numerical simulations show that the proposed strategy can control the coupled system and a small torque can stabilize both the attitude and orbit. In order to generate the control force and torque, the voltage distribution problem is studied in an optimal framework. The numerical results show that the control force and torque of electric sail can be realized by adjusting the voltage distribution of charged tethers.

Orbital express capture system: concept to reality

Science.gov (United States)

Stamm, Shane; Motaghedi, Pejmun

2004-08-01

The development of autonomous servicing of on-orbit spacecraft has been a sought after objective for many years. A critical component of on-orbit servicing involves the ability to successfully capture, institute mate, and perform electrical and fluid transfers autonomously. As part of a Small Business Innovation Research (SBIR) grant, Starsys Research Corporation (SRC) began developing such a system. Phase I of the grant started in 1999, with initial work focusing on simultaneously defining the parameters associated with successful docking while designing to those parameters. Despite the challenge of working without specific requirements, SRC completed development of a prototype design in 2000. Throughout the following year, testing was conducted on the prototype to characterize its performance. Having successfully completed work on the prototype, SRC began a Phase II SBIR effort in mid-2001. The focus of the second phase was a commercialization effort designed to augment the prototype model into a more flight-like design. The technical requirements, however, still needed clear definition for the design to progress. The advent of the Orbital Express (OE) program provided much of that definition. While still in the proposal stages of the OE program, SRC began tailoring prototype redesign efforts to the OE program requirements. A primary challenge involved striking a balance between addressing the technical requirements of OE while designing within the scope of the SBIR. Upon award of the OE contract, the Phase II SBIR design has been fully developed. This new design, designated the Mechanical Docking System (MDS), successfully incorporated many of the requirements of the OE program. SRC is now completing dynamic testing on the MDS hardware, with a parallel effort of developing a flight design for OE. As testing on the MDS progresses, the design path that was once common to both SBIR effort and the OE program begins to diverge. The MDS will complete the scope of the

Spacecraft electrical power subsystem: Failure behavior, reliability, and multi-state failure analyses

International Nuclear Information System (INIS)

Kim, So Young; Castet, Jean-Francois; Saleh, Joseph H.

2012-01-01

This article investigates the degradation and failure behavior of spacecraft electrical power subsystem (EPS) on orbit. First, this work provides updated statistical reliability and multi-state failure analyses of spacecraft EPS and its different constituents, namely the batteries, the power distribution, and the solar arrays. The EPS is shown to suffer from infant mortality and to be a major driver of spacecraft unreliability. Over 25% of all spacecraft failures are the result of EPS failures. As a result, satellite manufacturers may wish to pursue targeted improvement to this subsystem, either through better testing or burn-in procedures, better design or parts selection, or additional redundancy. Second, this work investigates potential differences in the EPS degradation and failure behavior for spacecraft in low earth orbits (LEO) and geosynchronous orbits (GEO). This analysis was motivated by the recognition that the power/load cycles and the space environment are significantly different in LEO and GEO, and as such, they may result in different failure behavior for the EPS in these two types of orbits. The results indicate, and quantify the extent to which, the EPS fails differently in LEO and GEO, both in terms of frequency and severity of failure events. A casual summary of the findings can be stated as follows: the EPS fails less frequently but harder (with fatal consequences to the spacecraft) in LEO than in GEO.

Pursuit/evasion in orbit

Science.gov (United States)

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.

Small Spacecraft Constellation Concept for Mars Atmospheric Radio Occultations

Science.gov (United States)

Asmar, S. W.; Mannucci, A. J.; Ao, C. O.; Kobayashi, M. M.; Lazio, J.; Marinan, A.; Massone, G.; McCandless, S. E.; Preston, R. A.; Seubert, J.; Williamson, W.

2017-12-01

First demonstrated in 1965 when Mariner IV flew by Mars and determined the salient features of its atmosphere, radio occultation experiments have been carried out on numerous planetary missions with great discoveries. These experiments utilize the now classic configuration of a signal from a single planetary spacecraft to Earth receiving stations, where the science data are acquired. The Earth science community advanced the technique to utilizing a constellation of spacecraft with the radio occultation links between the spacecraft, enabled by the infrastructure of the Global Positioning System. With the advent of small and less costly spacecraft, such as planetary CubeSats and other variations, such as the anticipated innovative Mars Cube One mission, crosslinks among small spacecraft can be used to study other planets in the near future. Advantages of this type of experiment include significantly greater geographical coverage, which could reach global coverage over a few weeks with a small number of spacecraft. Repeatability of the global coverage can lead to examining temperature-pressure profiles and ionospheric electron density profiles, on daily, seasonal, annual, or other time scales of interest. The higher signal-to-noise ratio for inter-satellite links, compared to a link to Earth, decreases the design demands on the instrumentation (smaller antennas and transmitters, etc.). After an actual Mars crosslink demonstration, this concept has been in development using Mars as a possible target. Scientific objectives, delivery methods, operational scenarios and end-to-end configuration have been documented. Science objectives include determining the state and variability of the lower Martian atmosphere, which has been an identified as a high priority objective by the Mars Exploration Program Analysis Group, particularly as it relates to entry, descent, and landing and ascent for future crewed and robotic missions. This paper will present the latest research on the

Compressed gas combined single- and two-stage light-gas gun

Science.gov (United States)

Lamberson, L. E.; Boettcher, P. A.

2018-02-01

With more than 1 trillion artificial objects smaller than 1 μm in low and geostationary Earth orbit, space assets are subject to the constant threat of space debris impact. These collisions occur at hypervelocity or speeds greater than 3 km/s. In order to characterize material behavior under this extreme event as well as study next-generation materials for space exploration, this paper presents a unique two-stage light-gas gun capable of replicating hypervelocity impacts. While a limited number of these types of facilities exist, they typically are extremely large and can be costly and dangerous to operate. The design presented in this paper is novel in two distinct ways. First, it does not use a form of combustion in the first stage. The projectile is accelerated from a pressure differential using air and inert gases (or purely inert gases), firing a projectile in a nominal range of 1-4 km/s. Second, the design is modular in that the first stage sits on a track sled and can be pulled back and used in itself to study lower speed impacts without any further modifications, with the first stage piston as the impactor. The modularity of the instrument allows the ability to investigate three orders of magnitude of impact velocities or between 101 and 103 m/s in a single, relatively small, cost effective instrument.

A localized swarm of low-resource CubeSat-class spacecraft for auroral ionospheric science

Science.gov (United States)

Clayton, R.; Lynch, K. A.; Gayetsky, L.; Guinther, J.; Slagle, A.; Currey, S.

2012-12-01

In interesting and dynamic auroral ionospheric plasmas, single-point in situ measurements are insufficient. Changes in measurements recorded from a single probe can be ascribed to either changes in position or to changes over time, and gradient scales can only be inferred. A localized array of sensors deployed as a low-resource swarm from a main deployer, can address these issues. We consider two aspects of designing such a swarm: (a) maintaining the localization in a low-cost manner, and (b) creating an extremely low-resource spacecraft by taking advantage of commercially available technologies. For a few-week low-altitude mission, STK (SatelliteToolKit) studies show that with proper deployment, an array of CubeSat-class spacecraft near 350 km altitude can regroup once per orbit to within a few 10s of km. Kepler's laws and Hill's equations allow us to put constraints on the capability of the deployer needed, in order to deploy the array with a minimal component of the ejection velocity along the orbital track. In order to keep the cost of each spacecraft low, we are exploring commercially available technologies such as Arduino controllers and video-game sensors. The Arduino on each payload will take in information from the sensors on the payload, and will send the information to a DNT-900MHz local area communications system. We show an example experiment measuring river flows on the Connecticut river, and discuss the design of our payload swarm.

Design considerations for single-stage and two-stage pneumatic pellet injectors

International Nuclear Information System (INIS)

Gouge, M.J.; Combs, S.K.; Fisher, P.W.; Milora, S.L.

1988-09-01

Performance of single-stage pneumatic pellet injectors is compared with several models for one-dimensional, compressible fluid flow. Agreement is quite good for models that reflect actual breech chamber geometry and incorporate nonideal effects such as gas friction. Several methods of improving the performance of single-stage pneumatic pellet injectors in the near term are outlined. The design and performance of two-stage pneumatic pellet injectors are discussed, and initial data from the two-stage pneumatic pellet injector test facility at Oak Ridge National Laboratory are presented. Finally, a concept for a repeating two-stage pneumatic pellet injector is described. 27 refs., 8 figs., 3 tabs

Optimal Weighting of Multi-Spacecraft Data to Estimate Gradients of Physical Fields

Science.gov (United States)

Chanteur, G. M.; Le Contel, O.; Sahraoui, F.; Retino, A.; Mirioni, L.

2016-12-01

Multi-spacecraft missions like the ESA mission CLUSTER and the NASA mission MMS are essential to improve our understanding of physical processes in space plasmas. Several methods were designed in the 90's during the preparation phase of the CLUSTER mission to estimate gradients of physical fields from simultaneous multi-points measurements [1, 2]. Both CLUSTER and MMS involve four spacecraft with identical full scientific payloads including various sensors of electromagnetic fields and different type of particle detectors. In the standard methods described in [1, 2], which are presently in use, data from the four spacecraft have identical weights and the estimated gradients are most reliable when the tetrahedron formed by the four spacecraft is regular. There are three types of errors affecting the estimated gradients (see chapter 14 in [1]) : i) truncature errors are due to local non-linearity of spatial variations, ii) physical errors are due to instruments, and iii) geometrical errors are due to uncertainties on the positions of the spacecraft. An assessment of truncature errors for a given observation requires a theoretical model of the measured field. Instrumental errors can easily be taken into account for a given geometry of the cluster but are usually less than the geometrical errors which diverge quite fast when the tetrahedron flattens, a circumstance occurring twice per orbit of the cluster. Hence reliable gradients can be estimated only on part of the orbit. Reciprocal vectors of the tetrahedron were presented in chapter 4 of [1], they have the advantage over other methods to treat the four spacecraft symmetrically and to allow a theoretical analysis of the errors (see chapters 4 of [1] and 4 of [2]). We will present Generalized Reciprocal Vectors for weighted data and an optimization procedure to improve the reliability of the estimated gradients when the tetrahedron is not regular. A brief example using CLUSTER or MMS data will be given. This approach

Propulsion Trade Studies for Spacecraft Swarm Mission Design

Science.gov (United States)

Dono, Andres; Plice, Laura; Mueting, Joel; Conn, Tracie; Ho, Michael

2018-01-01

Spacecraft swarms constitute a challenge from an orbital mechanics standpoint. Traditional mission design involves the application of methodical processes where predefined maneuvers for an individual spacecraft are planned in advance. This approach does not scale to spacecraft swarms consisting of many satellites orbiting in close proximity; non-deterministic maneuvers cannot be preplanned due to the large number of units and the uncertainties associated with their differential deployment and orbital motion. For autonomous small sat swarms in LEO, we investigate two approaches for controlling the relative motion of a swarm. The first method involves modified miniature phasing maneuvers, where maneuvers are prescribed that cancel the differential delta V of each CubeSat's deployment vector. The second method relies on artificial potential functions (APFs) to contain the spacecraft within a volumetric boundary and avoid collisions. Performance results and required delta V budgets are summarized, indicating that each method has advantages and drawbacks for particular applications. The mini phasing maneuvers are more predictable and sustainable. The APF approach provides a more responsive and distributed performance, but at considerable propellant cost. After considering current state of the art CubeSat propulsion systems, we conclude that the first approach is feasible, but the modified APF method of requires too much control authority to be enabled by current propulsion systems.

A comparison between the response of compound semiconductor detectors in single and back-to-back configuration

International Nuclear Information System (INIS)

Auricchio, N.; Donati, A.; Dusi, W.; Perillo, E.; Siffert, P.

2004-01-01

CdTe detectors (3 x 5 mm 2 electrode area and 1 mm thick), mounted in a back-to-back configuration with common anode, have been characterized, aiming at determining the actual regions of these devices giving the best and almost constant spectroscopic performance. This configuration ensures that the maximum paths of the charge carriers correspond to the inter-electrode distances and consequently the spectroscopic performance is essentially the same as for the single coupled detectors, but allows one: - to double the useful absorbing thickness when irradiated in the classical planar parallel field geometry; - to double the pixel area when irradiated in the planar transverse field (PTF) geometry, while the absorbing thickness can be freely adjusted according to the experimental requirements. The tests were performed in PTF geometry by using well collimated photon beams (10-300 keV energy range) obtained by a 20 mm thick tungsten collimator having a 0.1 x 1 or a 0.3 x 1 mm 2 collimating channel. Experimental results concerning the spectroscopic behaviour of the coupled detectors have been compared with those obtained by single devices having the same thickness as the single components or the same total thickness as the coupled device. The results show that the back-to-back configuration does not introduce further significant noise with respect to 1 mm thick single detectors and their spectroscopic performance is similar or better than the one shown by 2 mm thick single devices for fine and wide collimator measurements, and mostly in the uncollimated case

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

Science.gov (United States)

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.

Possible Periodic Orbit Control Maneuvers for an eLISA Mission

Science.gov (United States)

Bender, Peter L.; Welter, Gary L.

2012-01-01

This paper investigates the possible application of periodic orbit control maneuvers for so-called evolved-LISA (eLISA) missions, i.e., missions for which the constellation arm lengths and mean distance from the Earth are substantially reduced. We find that for missions with arm lengths of 106 km and Earth-trailing distance ranging from approx. 12deg to 20deg over the science lifetime, the occasional use of the spacecraft micro-Newton thrusters for constellation configuration maintenance should be able to essentially eliminate constellation distortion caused by Earth-induced tidal forces at a cost to science time of only a few percent. With interior angle variation kept to approx. +/-0:1deg, the required changes in the angles between the laser beam pointing directions for the two arms from any spacecraft could be kept quite small. This would considerably simplify the apparatus necessary for changing the transmitted beam directions.

A group arrival retrial G - queue with multi optional stages of service, orbital search and server breakdown

Science.gov (United States)

Radha, J.; Indhira, K.; Chandrasekaran, V. M.

2017-11-01

A group arrival feedback retrial queue with k optional stages of service and orbital search policy is studied. Any arriving group of customer finds the server free, one from the group enters into the first stage of service and the rest of the group join into the orbit. After completion of the i th stage of service, the customer under service may have the option to choose (i+1)th stage of service with θi probability, with pI probability may join into orbit as feedback customer or may leave the system with {q}i=≤ft\\{\\begin{array}{l}1-{p}i-{θ }i,i=1,2,\\cdots k-1\\ 1-{p}i,i=k\\end{array}\\right\\} probability. Busy server may get to breakdown due to the arrival of negative customers and the service channel will fail for a short interval of time. At the completion of service or repair, the server searches for the customer in the orbit (if any) with probability α or remains idle with probability 1-α. By using the supplementary variable method, steady state probability generating function for system size, some system performance measures are discussed.

In-target rare nuclei production rates with EURISOL single-stage configuration

CERN Document Server

Chabod, S P; Ene, D; Doré, D; Blideanu, V; David, J.-Ch; Ridikas, D

2010-01-01

We conducted calculations of exotic nuclei production rates for 320 configurations of EURISOL (European Isotope Separation On-Line Radioactive Ion Beam Facility) direct spallation targets. The nuclei yields were evaluated using neutron generation-transport codes, completed with evolution calculations to account for nuclei decays and low energy neutron interactions. The yields were optimized for 11 selected elements (Li, Be, Ne, Mg, Ar, Ni, Ga, Kr, Sn, Hg, Fr) and 23 of their isotopes, as function of the target compositions and geometries as well as the incident proton beam energies. For the considered elements, we evaluated the yield distributions as functions of the charge and mass numbers using two different spallation models.

An Integrated Vision-Based System for Spacecraft Attitude and Topology Determination for Formation Flight Missions

Science.gov (United States)

Rogers, Aaron; Anderson, Kalle; Mracek, Anna; Zenick, Ray

2004-01-01

With the space industry's increasing focus upon multi-spacecraft formation flight missions, the ability to precisely determine system topology and the orientation of member spacecraft relative to both inertial space and each other is becoming a critical design requirement. Topology determination in satellite systems has traditionally made use of GPS or ground uplink position data for low Earth orbits, or, alternatively, inter-satellite ranging between all formation pairs. While these techniques work, they are not ideal for extension to interplanetary missions or to large fleets of decentralized, mixed-function spacecraft. The Vision-Based Attitude and Formation Determination System (VBAFDS) represents a novel solution to both the navigation and topology determination problems with an integrated approach that combines a miniature star tracker with a suite of robust processing algorithms. By combining a single range measurement with vision data to resolve complete system topology, the VBAFDS design represents a simple, resource-efficient solution that is not constrained to certain Earth orbits or formation geometries. In this paper, analysis and design of the VBAFDS integrated guidance, navigation and control (GN&C) technology will be discussed, including hardware requirements, algorithm development, and simulation results in the context of potential mission applications.

Charging in the environment of large spacecraft

International Nuclear Information System (INIS)

Lai, S.T.

1993-01-01

This paper discusses some potential problems of spacecraft charging as a result of interactions between a large spacecraft, such as the Space Station, and its environment. Induced electric field, due to VXB effect, may be important for large spacecraft at low earth orbits. Differential charging, due to different properties of surface materials, may be significant when the spacecraft is partly in sunshine and partly in shadow. Triple-root potential jump condition may occur because of differential charging. Sudden onset of severe differential charging may occur when an electron or ion beam is emitted from the spacecraft. The beam may partially return to the ''hot spots'' on the spacecraft. Wake effects, due to blocking of ambient ion trajectories, may result in an undesirable negative potential region in the vicinity of a large spacecraft. Outgassing and exhaust may form a significant spacecraft induced environment; ionization may occur. Spacecraft charging and discharging may affect the electronic components on board

Protecting Spacecraft Fragments from Exposure to Small Debris

Directory of Open Access Journals (Sweden)

V. V. Zelentsov

2015-01-01

Full Text Available Since the launch of the first artificial Earth satellite a large amount of space debris has been accumulated in near-earth space. This debris comprises the exhausted spacecrafts, final stages of rocket-carriers and boosters, technological space junk, consisting of the structure elements, which are separated when deploying the solar arrays, antennas etc., as well as when undocking a booster and a spacecraft. All the debris is divided into observable one of over 100 mm in size and unobservable debris. In case of possible collision with the observed debris an avoidance manoeuvre is provided. The situation with unobservable debris is worse, its dimensions ranging from 100 mm to several microns. This debris is formed as a result of explosions of dead space objects and at collisions of destroyed spacecraft fragments against each other. This debris moves along arbitrary trajectories at different speeds.At collision of a spacecraft with fragments of small-size space debris, various consequences are possible: the device can immediately fail, suffer damages, which will have effect later and damages, which break no bones to the aircraft. Anyway, the spacecraft collision with small-size debris particles is undesirable. The protective shields are used to protect the aircraft from damage. Development of shield construction is complicated because the high cost of launch makes it impossible to conduct field tests of shields in space. All the work is carried out in the laboratory, with particles having co-impact speeds up to 10 km/s (possible speeds are up to 20 km/s and spherically shaped particles of 0.8 ... 3 mm in diameter.Various materials are used to manufacture shields. These are aluminum sheet, sandwich panels, metal mesh, metal foam, and woven materials (ballistic fabric. The paper considers single-layer (from sheet metal sandwich materials and multilayer shield designs. As experimental studies show, a single-layer shield protects colliding at speeds

Spacecraft Environmental Interactions Technology, 1983

Science.gov (United States)

1985-01-01

State of the art of environment interactions dealing with low-Earth-orbit plasmas; high-voltage systems; spacecraft charging; materials effects; and direction of future programs are contained in over 50 papers.

LDEF materials results for spacecraft applications: Executive summary

Science.gov (United States)

Whitaker, A. F.; Dooling, D.

1995-03-01

To address the challenges of space environmental effects, NASA designed the Long Duration Exposure Facility (LDEF) for an 18-month mission to expose thousands of samples of candidate materials that might be used on a space station or other orbital spacecraft. LDEF was launched in April 1984 and was to have been returned to Earth in 1985. Changes in mission schedules postponed retrieval until January 1990, after 69 months in orbit. Analyses of the samples recovered from LDEF have provided spacecraft designers and managers with the most extensive data base on space materials phenomena. Many LDEF samples were greatly changed by extended space exposure. Among even the most radially altered samples, NASA and its science teams are finding a wealth of surprising conclusions and tantalizing clues about the effects of space on materials. Many were discussed at the first two LDEF results conferences and subsequent professional papers. The LDEF Materials Results for Spacecraft Applications Conference was convened in Huntsville to discuss implications for spacecraft design. Already, paint and thermal blanket selections for space station and other spacecraft have been affected by LDEF data. This volume synopsizes those results.

Searching for Orbits with Minimum Fuel Consumption for Station-Keeping Maneuvers: An Application to Lunisolar Perturbations

Directory of Open Access Journals (Sweden)

Antonio Fernando Bertachini de Almeida Prado

2013-01-01

Full Text Available The present paper has the goal of developing a new criterion to search for orbits that minimize the fuel consumption for station-keeping maneuvers. This approach is based on the integral over the time of the perturbing forces. This integral measures the total variation of velocity caused by the perturbations in the spacecraft, which corresponds to the equivalent variation of velocity that an engine should deliver to the spacecraft to compensate the perturbations and to keep its orbit Keplerian all the time. This integral is a characteristic of the orbit and the set of perturbations considered and does not depend on the type of engine used. In this sense, this integral can be seen as a criterion to select the orbit of the spacecraft. When this value becomes larger, more consumption of fuel is required for the station keeping, and, in this sense, less interesting is the orbit. This concept can be applied to any perturbation. In the present research, as an example, the perturbation caused by a third body is considered. Then, numerical simulations considering the effects of the Sun and the Moon in a satellite around the Earth are shown to exemplify the method.

Computing in the presence of soft bit errors. [caused by single event upset on spacecraft

Science.gov (United States)

Rasmussen, R. D.

1984-01-01

It is shown that single-event-upsets (SEUs) due to cosmic rays are a significant source of single bit error in spacecraft computers. The physical mechanism of SEU, electron hole generation by means of Linear Energy Transfer (LET), it discussed with reference made to the results of a study of the environmental effects on computer systems of the Galileo spacecraft. Techniques for making software more tolerant of cosmic ray effects are considered, including: reducing the number of registers used by the software; continuity testing of variables; redundant execution of major procedures for error detection; and encoding state variables to detect single-bit changes. Attention is also given to design modifications which may reduce the cosmic ray exposure of on-board hardware. These modifications include: shielding components operating in LEO; removing low-power Schottky parts; and the use of CMOS diodes. The SEU parameters of different electronic components are listed in a table.

Orbital debris: a technical assessment

National Research Council Canada - National Science Library

Committee on Space Debris, National Research Council

..., and other debris created as a byproduct of space operations. Orbital Debris examines the methods we can use to characterize orbital debris, estimates the magnitude of the debris population, and assesses the hazard that this population poses to spacecraft...

Comparing a single-stage geocoding method to a multi-stage geocoding method: how much and where do they disagree?

Directory of Open Access Journals (Sweden)

Rice Kenneth

2007-03-01

Full Text Available Abstract Background Geocoding methods vary among spatial epidemiology studies. Errors in the geocoding process and differential match rates may reduce study validity. We compared two geocoding methods using 8,157 Washington State addresses. The multi-stage geocoding method implemented by the state health department used a sequence of local and national reference files. The single-stage method used a single national reference file. For each address geocoded by both methods, we measured the distance between the locations assigned by each method. Area-level characteristics were collected from census data, and modeled as predictors of the discordance between geocoded address coordinates. Results The multi-stage method had a higher match rate than the single-stage method: 99% versus 95%. Of 7,686 addresses were geocoded by both methods, 96% were geocoded to the same census tract by both methods and 98% were geocoded to locations within 1 km of each other by the two methods. The distance between geocoded coordinates for the same address was higher in sparsely populated and low poverty areas, and counties with local reference files. Conclusion The multi-stage geocoding method had a higher match rate than the single-stage method. An examination of differences in the location assigned to the same address suggested that study results may be most sensitive to the choice of geocoding method in sparsely populated or low-poverty areas.

Protecting Spacecraft Fragments from Exposure to Small Debris

OpenAIRE

V. V. Zelentsov

2015-01-01

Since the launch of the first artificial Earth satellite a large amount of space debris has been accumulated in near-earth space. This debris comprises the exhausted spacecrafts, final stages of rocket-carriers and boosters, technological space junk, consisting of the structure elements, which are separated when deploying the solar arrays, antennas etc., as well as when undocking a booster and a spacecraft. All the debris is divided into observable one of over 100 mm in size and unobservable ...

Circular revisit orbits design for responsive mission over a single target

Science.gov (United States)

Li, Taibo; Xiang, Junhua; Wang, Zhaokui; Zhang, Yulin

2016-10-01

The responsive orbits play a key role in addressing the mission of Operationally Responsive Space (ORS) because of their capabilities. These capabilities are usually focused on supporting specific targets as opposed to providing global coverage. One subtype of responsive orbits is repeat coverage orbit which is nearly circular in most remote sensing applications. This paper deals with a special kind of repeating ground track orbit, referred to as circular revisit orbit. Different from traditional repeat coverage orbits, a satellite on circular revisit orbit can visit a target site at both the ascending and descending stages in one revisit cycle. This typology of trajectory allows a halving of the traditional revisit time and does a favor to get useful information for responsive applications. However the previous reported numerical methods in some references often cost lots of computation or fail to obtain such orbits. To overcome this difficulty, an analytical method to determine the existence conditions of the solutions to revisit orbits is presented in this paper. To this end, the mathematical model of circular revisit orbit is established under the central gravity model and the J2 perturbation. A constraint function of the circular revisit orbit is introduced, and the monotonicity of that function has been studied. The existent conditions and the number of such orbits are naturally worked out. Taking the launch cost into consideration, optimal design model of circular revisit orbit is established to achieve a best orbit which visits a target twice a day in the morning and in the afternoon respectively for several days. The result shows that it is effective to apply circular revisit orbits in responsive application such as reconnoiter of natural disaster.

Special Semaphore Scheme for UHF Spacecraft Communications

Science.gov (United States)

Butman, Stanley; Satorius, Edgar; Ilott, Peter

2006-01-01

A semaphore scheme has been devised to satisfy a requirement to enable ultrahigh- frequency (UHF) radio communication between a spacecraft descending from orbit to a landing on Mars and a spacecraft, in orbit about Mars, that relays communications between Earth and the lander spacecraft. There are also two subsidiary requirements: (1) to use UHF transceivers, built and qualified for operation aboard the spacecraft that operate with residual-carrier binary phase-shift-keying (BPSK) modulation at a selectable data rate of 8, 32, 128, or 256 kb/s; and (2) to enable low-rate signaling even when received signals become so weak as to prevent communication at the minimum BPSK rate of 8 kHz. The scheme involves exploitation of Manchester encoding, which is used in conjunction with residual-carrier modulation to aid the carrier-tracking loop. By choosing various sequences of 1s, 0s, or 1s alternating with 0s to be fed to the residual-carrier modulator, one would cause the modulator to generate sidebands at a fundamental frequency of 4 or 8 kHz and harmonics thereof. These sidebands would constitute the desired semaphores. In reception, the semaphores would be detected by a software demodulator.

Dynamics and control of Lorentz-augmented spacecraft relative motion

CERN Document Server

Yan, Ye; Yang, Yueneng

2017-01-01

This book develops a dynamical model of the orbital motion of Lorentz spacecraft in both unperturbed and J2-perturbed environments. It explicitly discusses three kinds of typical space missions involving relative orbital control: spacecraft hovering, rendezvous, and formation flying. Subsequently, it puts forward designs for both open-loop and closed-loop control schemes propelled or augmented by the geomagnetic Lorentz force. These control schemes are entirely novel and represent a significantly departure from previous approaches.

Convergence of configuration-interaction single-center calculations of positron-atom interactions

International Nuclear Information System (INIS)

Mitroy, J.; Bromley, M. W. J.

2006-01-01

The configuration interaction (CI) method using orbitals centered on the nucleus has recently been applied to calculate the interactions of positrons interacting with atoms. Computational investigations of the convergence properties of binding energy, phase shift, and annihilation rate with respect to the maximum angular momentum of the orbital basis for the e + Cu and PsH bound states, and the e + -H scattering system were completed. The annihilation rates converge very slowly with angular momentum, and moreover the convergence with radial basis dimension appears to be slower for high angular momentum. A number of methods of completing the partial wave sum are compared; an approach based on a ΔX J =a(J+(1/2)) -n +b(J+(1/2)) -(n+1) form [with n=4 for phase shift (or energy) and n=2 for the annihilation rate] seems to be preferred on considerations of utility and underlying physical justification

Spacecraft Thermal Management

Science.gov (United States)

Hurlbert, Kathryn Miller

2009-01-01

In the 21st century, the National Aeronautics and Space Administration (NASA), the Russian Federal Space Agency, the National Space Agency of Ukraine, the China National Space Administration, and many other organizations representing spacefaring nations shall continue or newly implement robust space programs. Additionally, business corporations are pursuing commercialization of space for enabling space tourism and capital business ventures. Future space missions are likely to include orbiting satellites, orbiting platforms, space stations, interplanetary vehicles, planetary surface missions, and planetary research probes. Many of these missions will include humans to conduct research for scientific and terrestrial benefits and for space tourism, and this century will therefore establish a permanent human presence beyond Earth s confines. Other missions will not include humans, but will be autonomous (e.g., satellites, robotic exploration), and will also serve to support the goals of exploring space and providing benefits to Earth s populace. This section focuses on thermal management systems for human space exploration, although the guiding principles can be applied to unmanned space vehicles as well. All spacecraft require a thermal management system to maintain a tolerable thermal environment for the spacecraft crew and/or equipment. The requirements for human rating and the specified controlled temperature range (approximately 275 K - 310 K) for crewed spacecraft are unique, and key design criteria stem from overall vehicle and operational/programatic considerations. These criteria include high reliability, low mass, minimal power requirements, low development and operational costs, and high confidence for mission success and safety. This section describes the four major subsystems for crewed spacecraft thermal management systems, and design considerations for each. Additionally, some examples of specialized or advanced thermal system technologies are presented

Using Solar Radiation Pressure to Control L2 Orbits

Science.gov (United States)

Tene, Noam; Richon, Karen; Folta, David

1998-01-01

The main perturbations at the Sun-Earth Lagrange points L1 and L2 are from solar radiation pressure (SRP), the Moon and the planets. Traditional approaches to trajectory design for Lagrange-point orbits use maneuvers every few months to correct for these perturbations. The gravitational effects of the Moon and the planets are small and periodic. However, they cannot be neglected because small perturbations in the direction of the unstable eigenvector are enough to cause exponential growth within a few months. The main effect of a constant SRP is to shift the center of the orbit by a small distance. For spacecraft with large sun-shields like the Microwave Anisotropy Probe (MAP) and the Next Generation Space Telescope (NGST), the SRP effect is larger than all other perturbations and depends mostly on spacecraft attitude. Small variations in the spacecraft attitude are large enough to excite or control the exponential eigenvector. A closed-loop linear controller based on the SRP variations would eliminate one of the largest errors to the orbit and provide a continuous acceleration for use in controlling other disturbances. It is possible to design reference trajectories that account for the periodic lunar and planetary perturbations and still satisfy mission requirements. When such trajectories are used the acceleration required to control the unstable eigenvector is well within the capabilities of a continuous linear controller. Initial estimates show that by using attitude control it should be possible to minimize and even eliminate thruster maneuvers for station keeping.

Lifetime of a spacecraft around a synchronous system of asteroids using a dipole model

Science.gov (United States)

dos Santos, Leonardo Barbosa Torres; de Almeida Prado, Antonio F. Bertachini; Sanchez, Diogo Merguizo

2017-11-01

Space missions allow us to expand our knowledge about the origin of the solar system. It is believed that asteroids and comets preserve the physical characteristics from the time that the solar system was created. For this reason, there was an increase of missions to asteroids in the past few years. To send spacecraft to asteroids or comets is challenging, since these objects have their own characteristics in several aspects, such as size, shape, physical properties, etc., which are often only discovered after the approach and even after the landing of the spacecraft. These missions must be developed with sufficient flexibility to adjust to these parameters, which are better determined only when the spacecraft reaches the system. Therefore, conducting a dynamic investigation of a spacecraft around a multiple asteroid system offers an extremely rich environment. Extracting accurate information through analytical approaches is quite challenging and requires a significant number of restrictive assumptions. For this reason, a numerical approach to the dynamics of a spacecraft in the vicinity of a binary asteroid system is offered in this paper. In the present work, the equations of the Restricted Synchronous Four-Body Problem (RSFBP) are used to model a binary asteroid system. The main objective of this work is to construct grids of initial conditions, which relates semi-major axis and eccentricity, in order to quantify the lifetime of a spacecraft when released close to the less massive body of the binary system (modeled as a rotating mass dipole). We performed an analysis of the lifetime of the spacecraft considering several mass ratios of a binary system of asteroids and investigating the behavior of a spacecraft in the vicinity of this system. We analyze direct and retrograde orbits. This study investigated orbits that survive for at least 500 orbital periods of the system (which is approximately one year), then not colliding or escaping from the system during this

New current control based MPPT technique for single stage grid connected PV systems

International Nuclear Information System (INIS)

Jain, Sachin; Agarwal, Vivek

2007-01-01

This paper presents a new maximum power point tracking algorithm based on current control for a single stage grid connected photovoltaic system. The main advantage of this algorithm comes from its ability to predict the approximate amplitude of the reference current waveform or power that can be derived from the PV array with the help of an intermediate variable β. A variable step size for the change in reference amplitude during initial tracking helps in fast tracking. It is observed that if the reference current amplitude is greater than the array capacity, the system gets unstable (i.e. moves into the positive slope region of the p-v characteristics of the array). The proposed algorithm prevents the PV system from entering the positive slope region of the p-v characteristics. It is also capable of restoring stability if the system goes unstable due to a sudden environmental change. The proposed algorithm has been tested on a new single stage grid connected PV configuration recently developed by the authors to feed sinusoidal current into the grid. The system is operated in a continuous conduction mode to realize advantages such as low device current stress, high efficiency and low EMI. A fast MPPT tracker with single stage inverter topology operating in CCM makes the overall system highly efficient. Specific cases of the system, operating in just discontinuous current mode and discontinuous current mode and their relative merits and demerits are also discussed

Development of solid-gas equilibrium propulsion system for small spacecraft

Science.gov (United States)

Chujo, Toshihiro; Mori, Osamu; Kubo, Yuki

2017-11-01

A phase equilibrium propulsion system is a kind of cold-gas jet in which the phase equilibrium state of the fuel is maintained in a tank and its vapor is ejected when a valve is opened. One such example is a gas-liquid equilibrium propulsion system that uses liquefied gas as fuel. This system was mounted on the IKAROS solar sail and has been demonstrated in orbit. The system has a higher storage efficiency and a lighter configuration than a high-pressure cold-gas jet because the vapor pressure is lower, and is suitable for small spacecraft. However, the system requires a gas-liquid separation device in order to avoid leakage of the liquid, which makes the system complex. As another example of a phase equilibrium propulsion system, we introduce a solid-gas equilibrium propulsion system, which uses a sublimable substance as fuel and ejects its vapor. This system has an even lower vapor pressure and does not require such a separation device, instead requiring only a filter to keep the solid inside the tank. Moreover, the system is much simpler and lighter, making it more suitable for small spacecraft, especially CubeSat-class spacecraft, and the low thrust of the system allows spacecraft motion to be controlled precisely. In addition, the thrust level can be controlled by controlling the temperature of the fuel, which changes the vapor pressure. The present paper introduces the concept of the proposed system, and describes ejection experiments and its evaluation. The basic function of the proposed system is demonstrated in order to verify its usefulness.

Injection of a microsatellite in circular orbits using a three-stage launch vehicle

Science.gov (United States)

Marchi, L. O.; Murcia, J. O.; Prado, A. F. B. A.; Solórzano, C. R. H.

2017-10-01

The injection of a satellite into orbit is usually done by a multi-stage launch vehicle. Nowadays, the space market demonstrates a strong tendency towards the use of smaller satellites, because the miniaturization of the systems improve the cost/benefit of a mission. A study to evaluate the capacity of the Brazilian Microsatellite Launch Vehicle (VLM) to inject payloads into Low Earth Orbits is presented in this paper. All launches are selected to be made to the east side of the Alcântara Launch Center (CLA). The dynamical model to calculate the trajectory consists of the three degrees of freedom (3DOF) associated with the translational movement of the rocket. Several simulations are performed according to a set of restrictions imposed to the flight. The altitude reached in the separation of the second stage, the altitude and velocity of injection, the flight path angle at the moment of the activation of the third stage and the duration of the ballistic flight are presented as a function of the payload carried.

Saturn V Second Stage (S-II) Ready for Static Test

Science.gov (United States)

1965-01-01

Two workers are dwarfed by the five J-2 engines of the Saturn V second stage (S-II) as they make final inspections prior to a static test firing by North American Space Division. These five hydrogen -fueled engines produced one million pounds of thrust, and placed the Apollo spacecraft into earth orbit before departing for the moon. The towering 363-foot Saturn V was a multi-stage, multi-engine launch vehicle standing taller than the Statue of Liberty. Altogether, the Saturn V engines produced as much power as 85 Hoover Dams.

Mars Molniya Orbit Atmospheric Resource Mining

Science.gov (United States)

Mueller, Robert P.; Braun, Robert D.; Sibille, Laurent; Sforzo, Brandon; Gonyea, Keir; Ali, Hisham

2016-01-01

This NIAC (NASA Advanced Innovative Concepts) work will focus on Mars and will build on previous efforts at analyzing atmospheric mining at Earth and the outer solar system. Spacecraft systems concepts will be evaluated and traded, to assess feasibility. However the study will primarily examine the architecture and associated missions to explore the closure, constraints and critical parameters through sensitivity studies. The Mars atmosphere consists of 95.5 percent CO2 gas which can be converted to methane fuel (CH4) and Oxidizer (O2) for chemical rocket propulsion, if hydrogen is transported from electrolyzed water on the Mars surface or from Earth. By using a highly elliptical Mars Molniya style orbit, the CO2 atmosphere can be scooped, ram-compressed and stored while the spacecraft dips into the Mars atmosphere at periapsis. Successive orbits result in additional scooping of CO2 gas, which also serves to aerobrake the spacecraft, resulting in a decaying Molniya orbit.

Fiber-Optic Micrometeoroid/Orbital Debris Impact Detector System

Science.gov (United States)

Christiansen, Eric L.; Tennyson, R. C.; Morison, W. D.

2012-01-01

A document describes a reliable, lightweight micrometeoroid/orbital debris (MMOD) detection system that can be located at strategic positions of "high consequence" to provide real-time warning of a penetration, its location, and the extent of the damage to a spacecraft. The concept is to employ fiber-optic sensors to detect impact damage and penetration of spacecraft structures. The fibers are non-electrical, employ light waves, and are immune to electromagnetic interference. The fiber-optic sensor array can be made as a stand-alone product, being bonded to a flexible membrane material or a structure that is employed as a MMOD shield material. The optical sensors can also be woven into hybrid MMOD shielding fabrics. The glass fibers of the fiber-optic sensor provide a dual purpose in contributing to the breakup of MMOD projectiles. The grid arrays can be made in a modular configuration to provide coverage over any area desired. Each module can be connected to a central scanner instrument and be interrogated in a continuous or periodic mode.

Coordinated polar spacecraft, geosynchronous spacecraft, and ground-based observations of magnetopause processes and their coupling to the ionosphere

Directory of Open Access Journals (Sweden)

G. Le

2004-12-01

Full Text Available In this paper, we present in-situ observations of processes occurring at the magnetopause and vicinity, including surface waves, oscillatory magnetospheric field lines, and flux transfer events, and coordinated observations at geosynchronous orbit by the GOES spacecraft, and on the ground by CANOPUS and 210° Magnetic Meridian (210MM magnetometer arrays. On 7 February 2002, during a high-speed solar wind stream, the Polar spacecraft was skimming the magnetopause in a post-noon meridian plane for ~3h. During this interval, it made two short excursions and a few partial crossings into the magnetosheath and observed quasi-periodic cold ion bursts in the region adjacent to the magnetopause current layer. The multiple magnetopause crossings, as well as the velocity of the cold ion bursts, indicate that the magnetopause was oscillating with an ~6-min period. Simultaneous observations of Pc5 waves at geosynchronous orbit by the GOES spacecraft and on the ground by the CANOPUS magnetometer array reveal that these magnetospheric pulsations were forced oscillations of magnetic field lines directly driven by the magnetopause oscillations. The magnetospheric pulsations occurred only in a limited longitudinal region in the post-noon dayside sector, and were not a global phenomenon, as one would expect for global field line resonance. Thus, the magnetopause oscillations at the source were also limited to a localized region spanning ~4h in local time. These observations suggest that it is unlikely that the Kelvin-Helmholz instability and/or fluctuations in the solar wind dynamic pressure were the direct driving mechanisms for the observed boundary oscillations. Instead, the likely mechanism for the localized boundary oscillations was pulsed reconnection at the magnetopause occurring along the X-line extending over the same 4-h region. The Pc5 band pressure fluctuations commonly seen in high-speed solar wind streams may modulate the reconnection rate as an

Organizational design configurations in the early stages of firm’s life cycle

OpenAIRE

Mosca, Luigi; Gianecchini, Martina; Campagnolo, Diego

2017-01-01

http://ocs.editorial.upv.es/index.php/CARMA/CARMA2016 Mosca, L.; Gianecchini, M.; Campagnolo, D. (2017). Organizational design configurations in the early stages of firm’s life cycle. Editorial Universitat Politècnica de València. doi:10.4995/CARMA2016.2015.3104. OCS

The Effect of Orbital Configuration on the Possible Climates and Habitability of Kepler-62f.

Science.gov (United States)

Shields, Aomawa L; Barnes, Rory; Agol, Eric; Charnay, Benjamin; Bitz, Cecilia; Meadows, Victoria S

2016-06-01

As lower-mass stars often host multiple rocky planets, gravitational interactions among planets can have significant effects on climate and habitability over long timescales. Here we explore a specific case, Kepler-62f (Borucki et al., 2013 ), a potentially habitable planet in a five-planet system with a K2V host star. N-body integrations reveal the stable range of initial eccentricities for Kepler-62f is 0.00 ≤ e ≤ 0.32, absent the effect of additional, undetected planets. We simulate the tidal evolution of Kepler-62f in this range and find that, for certain assumptions, the planet can be locked in a synchronous rotation state. Simulations using the 3-D Laboratoire de Météorologie Dynamique (LMD) Generic global climate model (GCM) indicate that the surface habitability of this planet is sensitive to orbital configuration. With 3 bar of CO2 in its atmosphere, we find that Kepler-62f would only be warm enough for surface liquid water at the upper limit of this eccentricity range, providing it has a high planetary obliquity (between 60° and 90°). A climate similar to that of modern-day Earth is possible for the entire range of stable eccentricities if atmospheric CO2 is increased to 5 bar levels. In a low-CO2 case (Earth-like levels), simulations with version 4 of the Community Climate System Model (CCSM4) GCM and LMD Generic GCM indicate that increases in planetary obliquity and orbital eccentricity coupled with an orbital configuration that places the summer solstice at or near pericenter permit regions of the planet with above-freezing surface temperatures. This may melt ice sheets formed during colder seasons. If Kepler-62f is synchronously rotating and has an ocean, CO2 levels above 3 bar would be required to distribute enough heat to the nightside of the planet to avoid atmospheric freeze-out and permit a large enough region of open water at the planet's substellar point to remain stable. Overall, we find multiple plausible combinations of

Properties of Coronal Shocks at the Origin of SEP events Observed by Only One Single Spacecraft

Science.gov (United States)

Lario, D.; Kwon, R.

2017-12-01

The simultaneous observation of solar energetic particle (SEP) events by multiple spacecraft distributed in the interplanetary medium depends not only on the spatial separation among the different spacecraft, but also on the properties of the particle sources and the characteristics of the SEP transport in interplanetary space. Among the SEP events observed by STEREO-A, STEREO-B and/or near-Earth spacecraft during solar cycle 24, we select SEP events observed by a single spacecraft (specifically, the SEP events observed only by near-Earth spacecraft on 2012 April 5, 2011 September 4, and 2013 August 17). We analyze whether the properties of the coronal shock associated with the origin of the events (as seen in extreme-ultraviolet and white-light coronal images) differ from those associated with SEP events observed by two or three spacecraft. For the selected events we find that the associated CMEs are, in general, narrower than those associated with SEP events observed by two or three spacecraft. The confined extension of the parent coronal shock and the absence of magnetic connection between distant spacecraft and the regions of the expanding coronal shock able to efficiently accelerate SEPs seem to be the conditions leading to intense SEP events observed only over narrow regions of interplanetary space by spacecraft magnetically connected to regions close to the parent eruption site. Weak and gradual intensity increases observed in extended regions of space might involve transport processes and/or later connections established with interplanetary shocks. Systematic analyses of a larger number of events are required before drawing firm conclusions.

Performance Evaluation of the Single-Phase Split-Source Inverter Using an Alternative DC-AC Configuration

DEFF Research Database (Denmark)

Abdelhakim, Ahmed; Mattavelli, Paolo; Davari, Pooya

2018-01-01

This paper investigates and evaluates the performance of a single-phase split-source inverter (SSI), where an alternative unidirectional dc-ac configuration is used. Such configuration is utilized in order to use two common-cathode diodes in a single-device instead of using two separate diodes, r...

Thermodynamic analysis and economical evaluation of two 310-80 K pre-cooling stage configurations for helium refrigeration and liquefaction cycle

Science.gov (United States)

Zhu, Z. G.; Zhuang, M.; Jiang, Q. F.; Y Zhang, Q.; Feng, H. S.

2017-12-01

In 310-80 K pre-cooling stage, the temperature of the HP helium stream reduces to about 80 K where nearly 73% of the enthalpy drop from room temperature to 4.5 K occurs. Apart from the most common liquid nitrogen pre-cooling, another 310-80 K pre-cooling configuration with turbine is employed in some helium cryoplants. In this paper, thermodynamic and economical performance of these two kinds of 310-80 K pre-cooling stage configurations has been studied at different operating conditions taking discharge pressure, isentropic efficiency of turbines and liquefaction rate as independent parameters. The exergy efficiency, total UA of heat exchangers and operating cost of two configurations are computed. This work will provide a reference for choosing 310-80 K pre-cooling stage configuration during design.

Development of the solar array deployment and drive system for the XTE spacecraft

Science.gov (United States)

Farley, Rodger; Ngo, Son

1995-01-01

The X-ray Timing Explorer (XTE) spacecraft is a NASA science low-earth orbit explorer-class satellite to be launched in 1995, and is an in-house Goddard Space Flight Center (GSFC) project. It has two deployable aluminum honeycomb solar array wings with each wing being articulated by a single axis solar array drive assembly. This paper will address the design, the qualification testing, and the development problems as they surfaced of the Solar Array Deployment and Drive System.

Formation flying for electric sails in displaced orbits. Part I: Geometrical analysis

Science.gov (United States)

Wang, Wei; Mengali, Giovanni; Quarta, Alessandro A.; Yuan, Jianping

2017-09-01

We present a geometrical methodology for analyzing the formation flying of electric solar wind sail based spacecraft that operate in heliocentric, elliptic, displaced orbits. The spacecraft orbit is maintained by adjusting its propulsive acceleration modulus, whose value is estimated using a thrust model that takes into account a variation of the propulsive performance with the sail attitude. The properties of the relative motion of the spacecraft are studied in detail and a geometrical solution is obtained in terms of relative displaced orbital elements, assumed to be small quantities. In particular, for the small eccentricity case (i.e. for a near-circular displaced orbit), the bounds characterized by the extreme values of relative distances are analytically calculated, thus providing an useful mathematical tool for preliminary design of the spacecraft formation structure.

Tracking the ultrafast motion of a single molecule by femtosecond orbital imaging

Science.gov (United States)

Cocker, Tyler L.; Peller, Dominik; Yu, Ping; Repp, Jascha; Huber, Rupert

2016-11-01

Watching a single molecule move on its intrinsic timescale has been one of the central goals of modern nanoscience, and calls for measurements that combine ultrafast temporal resolution with atomic spatial resolution. Steady-state experiments access the requisite spatial scales, as illustrated by direct imaging of individual molecular orbitals using scanning tunnelling microscopy or the acquisition of tip-enhanced Raman and luminescence spectra with sub-molecular resolution. But tracking the intrinsic dynamics of a single molecule directly in the time domain faces the challenge that interactions with the molecule must be confined to a femtosecond time window. For individual nanoparticles, such ultrafast temporal confinement has been demonstrated by combining scanning tunnelling microscopy with so-called lightwave electronics, which uses the oscillating carrier wave of tailored light pulses to directly manipulate electronic motion on timescales faster even than a single cycle of light. Here we build on ultrafast terahertz scanning tunnelling microscopy to access a state-selective tunnelling regime, where the peak of a terahertz electric-field waveform transiently opens an otherwise forbidden tunnelling channel through a single molecular state. It thereby removes a single electron from an individual pentacene molecule’s highest occupied molecular orbital within a time window shorter than one oscillation cycle of the terahertz wave. We exploit this effect to record approximately 100-femtosecond snapshot images of the orbital structure with sub-ångström spatial resolution, and to reveal, through pump/probe measurements, coherent molecular vibrations at terahertz frequencies directly in the time domain. We anticipate that the combination of lightwave electronics and the atomic resolution of our approach will open the door to visualizing ultrafast photochemistry and the operation of molecular electronics on the single-orbital scale.

Comparison of single-stage and temperature-phased two-stage anaerobic digestion of oily food waste

International Nuclear Information System (INIS)

Wu, Li-Jie; Kobayashi, Takuro; Li, Yu-You; Xu, Kai-Qin

2015-01-01

Highlights: • A single-stage and two two-stage anaerobic systems were synchronously operated. • Similar methane production 0.44 L/g VS_a_d_d_e_d from oily food waste was achieved. • The first stage of the two-stage process became inefficient due to serious pH drop. • Recycle favored the hythan production in the two-stage digestion. • The conversion of unsaturated fatty acids was enhanced by recycle introduction. - Abstract: Anaerobic digestion is an effective technology to recover energy from oily food waste. A single-stage system and temperature-phased two-stage systems with and without recycle for anaerobic digestion of oily food waste were constructed to compare the operation performances. The synchronous operation indicated the similar ability to produce methane in the three systems, with a methane yield of 0.44 L/g VS_a_d_d_e_d. The pH drop to less than 4.0 in the first stage of two-stage system without recycle resulted in poor hydrolysis, and methane or hydrogen was not produced in this stage. Alkalinity supplement from the second stage of two-stage system with recycle improved pH in the first stage to 5.4. Consequently, 35.3% of the particulate COD in the influent was reduced in the first stage of two-stage system with recycle according to a COD mass balance, and hydrogen was produced with a percentage of 31.7%, accordingly. Similar solids and organic matter were removed in the single-stage system and two-stage system without recycle. More lipid degradation and the conversion of long-chain fatty acids were achieved in the single-stage system. Recycling was proved to be effective in promoting the conversion of unsaturated long-chain fatty acids into saturated fatty acids in the two-stage system.

Trace Contaminant Control During the International Space Station's On-Orbit Assembly and Outfitting

Science.gov (United States)

Perry, J. L.

2017-01-01

Achieving acceptable cabin air quality must balance competing elements during spacecraft design, assembly, ground processing, and flight operations. Among the elements that contribute to the trace chemical contaminant load and, therefore, the cabin air quality aboard crewed spacecraft are the vehicle configuration, crew size and activities, mission duration and objectives, materials selection, and vehicle manufacturing and preflight ground processing methods. Trace chemical contaminants produced from pervasive sources such as equipment offgassing, human metabolism, and cleaning fluids during preflight ground processing present challenges to maintaining acceptable cabin air quality. To address these challenges, both passive and active contamination control techniques are used during a spacecraft's design, manufacturing, preflight preparation, and operational phases. Passive contamination control methods seek to minimize the equipment offgassing load by selecting materials, manufacturing processes, preflight preparation processes, and in-flight operations that have low chemical offgassing characteristics. Passive methods can be employed across the spacecraft's entire life cycle from conceptual design through flight operations. However, because the passive contamination control techniques cannot fully eliminate the contaminant load, active contamination control equipment must be deployed aboard the spacecraft to purify and revitalize the cabin atmosphere during in-flight operations. Verifying that the passive contamination control techniques have successfully maintained the total trace contaminant load within the active contamination control equipment's capabilities occurs late in the preflight preparation stages. This verification consists of subjecting the spacecraft to an offgassing test to determine the trace contaminant load. This load is then assessed versus the active contamination control equipment's capabilities via trace contaminant control (TCC) engineering

Programs To Optimize Spacecraft And Aircraft Trajectories

Science.gov (United States)

Brauer, G. L.; Petersen, F. M.; Cornick, D.E.; Stevenson, R.; Olson, D. W.

1994-01-01

POST/6D POST is set of two computer programs providing ability to target and optimize trajectories of powered or unpowered spacecraft or aircraft operating at or near rotating planet. POST treats point-mass, three-degree-of-freedom case. 6D POST treats more-general rigid-body, six-degree-of-freedom (with point masses) case. Used to solve variety of performance, guidance, and flight-control problems for atmospheric and orbital vehicles. Applications include computation of performance or capability of vehicle in ascent, or orbit, and during entry into atmosphere, simulation and analysis of guidance and flight-control systems, dispersion-type analyses and analyses of loads, general-purpose six-degree-of-freedom simulation of controlled and uncontrolled vehicles, and validation of performance in six degrees of freedom. Written in FORTRAN 77 and C language. Two machine versions available: one for SUN-series computers running SunOS(TM) (LAR-14871) and one for Silicon Graphics IRIS computers running IRIX(TM) operating system (LAR-14869).

Increased electric sail thrust through removal of trapped shielding electrons by orbit chaotisation due to spacecraft body

Directory of Open Access Journals (Sweden)

P. Janhunen

2009-08-01

Full Text Available An electric solar wind sail is a recently introduced propellantless space propulsion method whose technical development has also started. The electric sail consists of a set of long, thin, centrifugally stretched and conducting tethers which are charged positively and kept in a high positive potential of order 20 kV by an onboard electron gun. The positively charged tethers deflect solar wind protons, thus tapping momentum from the solar wind stream and producing thrust. The amount of obtained propulsive thrust depends on how many electrons are trapped by the potential structures of the tethers, because the trapped electrons tend to shield the charged tether and reduce its effect on the solar wind. Here we present physical arguments and test particle calculations indicating that in a realistic three-dimensional electric sail spacecraft there exist a natural mechanism which tends to remove the trapped electrons by chaotising their orbits and causing them to eventually collide with the conducting tethers. We present calculations which indicate that if these mechanisms were able to remove trapped electrons nearly completely, the electric sail performance could be about five times higher than previously estimated, about 500 nN/m, corresponding to 1 N thrust for a baseline construction with 2000 km total tether length.

Orbital Express fluid transfer demonstration system

Science.gov (United States)

Rotenberger, Scott; SooHoo, David; Abraham, Gabriel

2008-04-01

Propellant resupply of orbiting spacecraft is no longer in the realm of high risk development. The recently concluded Orbital Express (OE) mission included a fluid transfer demonstration that operated the hardware and control logic in space, bringing the Technology Readiness Level to a solid TRL 7 (demonstration of a system prototype in an operational environment). Orbital Express (funded by the Defense Advanced Research Projects Agency, DARPA) was launched aboard an Atlas-V rocket on March 9th, 2007. The mission had the objective of demonstrating technologies needed for routine servicing of spacecraft, namely autonomous rendezvous and docking, propellant resupply, and orbital replacement unit transfer. The demonstration system used two spacecraft. A servicing vehicle (ASTRO) performed multiple dockings with the client (NextSat) spacecraft, and performed a variety of propellant transfers in addition to exchanges of a battery and computer. The fluid transfer and propulsion system onboard ASTRO, in addition to providing the six degree-of-freedom (6 DOF) thruster system for rendezvous and docking, demonstrated autonomous transfer of monopropellant hydrazine to or from the NextSat spacecraft 15 times while on orbit. The fluid transfer system aboard the NextSat vehicle was designed to simulate a variety of client systems, including both blowdown pressurization and pressure regulated propulsion systems. The fluid transfer demonstrations started with a low level of autonomy, where ground controllers were allowed to review the status of the demonstration at numerous points before authorizing the next steps to be performed. The final transfers were performed at a full autonomy level where the ground authorized the start of a transfer sequence and then monitored data as the transfer proceeded. The major steps of a fluid transfer included the following: mate of the coupling, leak check of the coupling, venting of the coupling, priming of the coupling, fluid transfer, gauging

The Effect of Air Drag in Optimal Power-Limited Rendezvous Between Coplanar Low-Earth Orbits

Directory of Open Access Journals (Sweden)

Gil-Young Maeng

1998-06-01

Full Text Available The effect of air drag was researched when a low-earth orbit spacecraft using power-limited thruster rendezvoused another low-earth orbit spacecraft. The air density was assumed to decrease exponentially. The radius of parking orbit was 6655.935 km and that of target orbit was 7321.529 km. From the trajectories of active vehicles, the fuel consumption and the magnitude of thrust acceleration, we could conclude that the effect of air drag had to be considered in fuel optimal rendezvous problem between low-earth orbit spacecrafts. In multiple-revolution rendezvous case, the air drag was more effective.

Displaced Electric Sail Orbits Design and Transition Trajectory Optimization

Directory of Open Access Journals (Sweden)

Naiming Qi

2014-01-01

Full Text Available Displaced orbits for spacecraft propelled by electric sails are investigated as an alternative to the use of solar sails. The orbital dynamics of electric sails based spacecraft are studied within a spherical coordinate system, which permits finding the solutions of displaced electric sail orbits and optimize transfer trajectory. Transfer trajectories from Earth's orbit to displaced orbit are also studied in an optimal framework, by using genetic algorithm and Gauss pseudospectral method. The initial guesses for the state and control histories used in the Gauss pseudospectral method are interpolated from the best solution of a genetic algorithm. Numerical simulations show that the electric sail is able to perform the transfer from Earth’s orbit to displaced orbit in acceptable time, and the hybrid optimization method has the capability to search the feasible and optimal solution without any initial value guess.

Gravity Probe B spacecraft description

International Nuclear Information System (INIS)

Bennett, Norman R; Burns, Kevin; Katz, Russell; Kirschenbaum, Jon; Mason, Gary; Shehata, Shawky

2015-01-01

The Gravity Probe B spacecraft, developed, integrated, and tested by Lockheed Missiles and Space Company and later Lockheed Martin Corporation, consisted of structures, mechanisms, command and data handling, attitude and translation control, electrical power, thermal control, flight software, and communications. When integrated with the payload elements, the integrated system became the space vehicle. Key requirements shaping the design of the spacecraft were: (1) the tight mission timeline (17 months, 9 days of on-orbit operation), (2) precise attitude and translational control, (3) thermal protection of science hardware, (4) minimizing aerodynamic, magnetic, and eddy current effects, and (5) the need to provide a robust, low risk spacecraft. The spacecraft met all mission requirements, as demonstrated by dewar lifetime meeting specification, positive power and thermal margins, precision attitude control and drag-free performance, reliable communications, and the collection of more than 97% of the available science data. (paper)

Review of current activities to model and measure the orbital debris environment in low-earth orbit

Science.gov (United States)

Reynolds, R. C.

A very active orbital debris program is currently being pursued at the NASA/Johnson Space Center (JSC), with projects designed to better define the current environment, to project future environments, to model the processes contributing to or constraining the growth of debris in the environment, and to gather supporting data needed to improve the understanding of the orbital debris problem and the hazard it presents to spacecraft. This paper is a review of the activity being conducted at JSC, by NASA, Lockheed Engineering and Sciences Company, and other support contractors, and presents a review of current activity, results of current research, and a discussion of directions for future development.

Attitude Control Tradeoff Study Between the Use of a Flexible Beam and a Tether Configuration for the Connection of Two Bodies in Orbit

Science.gov (United States)

Graff, S. H.

1985-01-01

Sometimes it is necessary to mount a payload remotely from the main body of a spacecraft or space station. The reasons for this vary from vibration isolation to avoidance of measurement contamination. For example the SP-100 project, which grew out of the increased interest in nuclear power in space for space stations and for deep space explorations, requires separation of the nuclear reactor from the user because of vibration, heat and radiation. The different attitude control problems for beam and tether configurations are discussed. The beam configuration uses a conservative design approach. The vibration, beam flexibility and deployment concerns are analyzed. The tether configuration offers some very attractive design features, but not without several thorny problems. These problems are analyzed. One configuration will be recommended for the main thrust of the SP-100 design effort based on attitude control considerations.

Spacecraft attitude determination using the earth's magnetic field

Science.gov (United States)

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.

On the atmospheric drag in orbit determination for low Earth orbit

Science.gov (United States)

Tang, Jingshi; Liu, Lin; Miao, Manqian

2012-07-01

The atmosphere model is always a major limitation for low Earth orbit (LEO) in orbit prediction and determination. The accelerometer can work around the non-gravitational perturbations in orbit determination, but it helps little to improve the atmosphere model or to predict the orbit. For certain satellites, there may be some specific software to handle the orbit problem. This solution can improve the orbit accuracy for both prediction and determination, yet it always contains empirical terms and is exclusive for certain satellites. This report introduces a simple way to handle the atmosphere drag for LEO, which does not depend on instantaneous atmosphere conditions and improves accuracy of predicted orbit. This approach, which is based on mean atmospheric density, is supported by two reasons. One is that although instantaneous atmospheric density is very complicated with time and height, the major pattern is determined by the exponential variation caused by hydrostatic equilibrium and periodic variation caused by solar radiation. The mean density can include the major variations while neglect other minor details. The other reason is that the predicted orbit is mathematically the result from integral and the really determinant factor is the mean density instead of instantaneous density for every time and spot. Using the mean atmospheric density, which is mainly determined by F10.7 solar flux and geomagnetic index, can be combined into an overall parameter B^{*} = C_{D}(S/m)ρ_{p_{0}}. The combined parameter contains several less accurate parameters and can be corrected during orbit determination. This approach has been confirmed in various LEO computations and an example is given below using Tiangong-1 spacecraft. Precise orbit determination (POD) is done using one-day GPS positioning data without any accurate a-priori knowledge on spacecraft or atmosphere conditions. Using the corrected initial state vector of the spacecraft and the parameter B^* from POD, the

Wave-Based Attitude Control of Spacecraft with Fuel Sloshing Dynamics

Directory of Open Access Journals (Sweden)

Thompson Joseph William

2016-06-01

Full Text Available Wave-Based Control has been previously applied successfully to simple under-actuated flexible mechanical systems. Spacecraft and rockets with structural flexibility and sloshing are examples of such systems but have added difficulties due to non-uniform structure, external disturbing forces and non-ideal actuators and sensors. The aim of this paper is to extend the application of WBC to spacecraft systems, to compare the performance of WBC to other popular controllers and to carry out experimental validation of the designed control laws. A mathematical model is developed for an upper stage accelerating rocket moving in a single plane. Fuel sloshing is represented by an equivalent mechanical pendulum model. A wave-based controller is designed for the upper stage AVUM of the European launcher Vega. In numerical simulations the controller successfully suppresses the sloshing motion. A major advantage of the strategy is that no measurement of the pendulum states (sloshing motion is required.

Man-Made Debris In and From Lunar Orbit

Science.gov (United States)

Johnson, Nicholas L.; McKay, Gordon A. (Technical Monitor)

1999-01-01

During 1966-1976, as part of the first phase of lunar exploration, 29 manned and robotic missions placed more than 40 objects into lunar orbit. Whereas several vehicles later successfully landed on the Moon and/or returned to Earth, others were either abandoned in orbit or intentionally sent to their destruction on the lunar surface. The former now constitute a small population of lunar orbital debris; the latter, including four Lunar Orbiters and four Lunar Module ascent stages, have contributed to nearly 50 lunar sites of man's refuse. Other lunar satellites are known or suspected of having fallen from orbit. Unlike Earth satellite orbital decays and deorbits, lunar satellites impact the lunar surface unscathed by atmospheric burning or melting. Fragmentations of lunar satellites, which would produce clouds of numerous orbital debris, have not yet been detected. The return to lunar orbit in the 1990's by the Hagoromo, Hiten, Clementine, and Lunar Prospector spacecraft and plans for increased lunar exploration early in the 21st century, raise questions of how best to minimize and to dispose of lunar orbital debris. Some of the lessons learned from more than 40 years of Earth orbit exploitation can be applied to the lunar orbital environment. For the near-term, perhaps the most important of these is postmission passivation. Unique solutions, e.g., lunar equatorial dumps, may also prove attractive. However, as with Earth satellites, debris mitigation measures are most effectively adopted early in the concept and design phase, and prevention is less costly than remediation.

REQUIREMENTS FOR IMAGE QUALITY OF EMERGENCY SPACECRAFTS

Directory of Open Access Journals (Sweden)

A. I. Altukhov

2015-05-01

Full Text Available The paper deals with the method for formation of quality requirements to the images of emergency spacecrafts. The images are obtained by means of remote sensing of near-earth space orbital deployment in the visible range. of electromagnetic radiation. The method is based on a joint taking into account conditions of space survey, characteristics of surveillance equipment, main design features of the observed spacecrafts and orbital inspection tasks. Method. Quality score is the predicted linear resolution image that gives the possibility to create a complete view of pictorial properties of the space image obtained by electro-optical system from the observing satellite. Formulation of requirements to the numerical value of this indicator is proposed to perform based on the properties of remote sensing system, forming images in the conditions of outer space, and the properties of the observed emergency spacecraft: dimensions, platform construction of the satellite, on-board equipment placement. For method implementation the authors have developed a predictive model of requirements to a linear resolution for images of emergency spacecrafts, making it possible to select the intervals of space shooting and get the satellite images required for quality interpretation. Main results. To verify the proposed model functionality we have carried out calculations of the numerical values for the linear resolution of the image, ensuring the successful task of determining the gross structural damage of the spacecrafts and identifying changes in their spatial orientation. As input data were used with dimensions and geometric primitives corresponding to the shape of deemed inspected spacecrafts: Resurs-P", "Canopus-B", "Electro-L". Numerical values of the linear resolution images have been obtained, ensuring the successful task solution for determining the gross structural damage of spacecrafts.

Seniority and orbital symmetry as tools for establishing a full configuration interaction hierarchy.

Science.gov (United States)

Bytautas, Laimutis; Henderson, Thomas M; Jiménez-Hoyos, Carlos A; Ellis, Jason K; Scuseria, Gustavo E

2011-07-28

We explore the concept of seniority number (defined as the number of unpaired electrons in a determinant) when applied to the problem of electron correlation in atomic and molecular systems. Although seniority is a good quantum number only for certain model Hamiltonians (such as the pairing Hamiltonian), we show that it provides a useful partitioning of the electronic full configuration interaction (FCI) wave function into rapidly convergent Hilbert subspaces whose weight diminishes as its seniority number increases. The primary focus of this study is the adequate description of static correlation effects. The examples considered are the ground states of the helium, beryllium, and neon atoms, the symmetric dissociation of the N(2) and CO(2) molecules, as well as the symmetric dissociation of an H(8) hydrogen chain. It is found that the symmetry constraints that are normally placed on the spatial orbitals greatly affect the convergence rate of the FCI expansion. The energy relevance of the seniority zero sector (determinants with all paired electrons) increases dramatically if orbitals of broken spatial symmetry (as those commonly used for Hubbard Hamiltonian studies) are allowed in the wave function construction. © 2011 American Institute of Physics

Failure analysis of satellite subsystems to define suitable de-orbit devices

Science.gov (United States)

Palla, Chiara; Peroni, Moreno; Kingston, Jennifer

2016-11-01

Space missions in Low Earth Orbit (LEO) are severely affected by the build-up of orbital debris. A key practice, to be compliant with IADC (Inter-Agency Space Debris Coordination Committee) mitigation guidelines, is the removal of space systems that interfere with the LEO region not later than 25 years after the End of Mission. It is important to note that the current guidelines are not generally legally binding, even if different Space Agencies are now looking at the compliance for their missions. If the guidelines will change in law, it will be mandatory to have a postmission disposal strategy for all satellites, including micro and smaller classes. A potential increased number of these satellites is confirmed by different projections, in particular in the commercial sector. Micro and smaller spacecraft are, in general, not provided with propulsion capabilities to achieve a controlled re-entry, so they need different de-orbit disposal methods. When considering the utility of different debris mitigation methods, it is useful to understand which spacecraft subsystems are most likely to fail and how this may affect the operation of a de-orbit system. This also helps the consideration of which components are the most relevant or should be redundant depending on the satellite mass class. This work is based on a sample of LEO and MEO satellites launched between January 2000 and December 2014 with mass lower than 1000 kg. Failure analysis of satellite subsystems is performed by means of the Kaplan-Meier survival analysis; the parametric fits are conducted with Weibull distributions. The study is carried out by using the satellite database SpaceTrak™ which provides anomalies, failures, and trends information for spacecraft subsystems and launch vehicles. The database identifies five states for each satellite subsystem: three degraded states, one fully operational state, and one failed state (complete failure). The results obtained can guide the identification of the

System concepts and design examples for optical communication with planetary spacecraft

Science.gov (United States)

Lesh, James R.

Systems concepts for optical communication with future deep-space (planetary) spacecraft are described. These include not only the optical transceiver package aboard the distant spacecraft, but the earth-vicinity optical-communications receiving station as well. Both ground-based, and earth-orbiting receivers are considered. Design examples for a number of proposed or potential deep-space missions are then presented. These include an orbital mission to Saturn, a Lander and Rover mission to Mars, and an astronomical mission to a distance of 1000 astronomical units.

Spacecraft Charging Modeling -- Nascap-2k 2014 Annual Report

Science.gov (United States)

2014-09-19

appears to work similarly in Internet Explorer, FireFox , and Opera, but fails in Safari and Chrome. Note that the SEE Spacecraft Charging Handbook is... Characteristics of Spacecraft Charging in Low Earth Orbit, J Geophys Res. 11 7, doi: 10.1029/20 11JA016875, 2012. 2 M. Cho, K. Saito, T. Hamanaga, Data

ExoMars Trace Gas Orbiter provides atmospheric data during Aerobraking into its final orbit

Science.gov (United States)

Svedhem, Hakan; Vago, Jorge L.; Bruinsma, Sean; Müller-Wodarg, Ingo; ExoMars 2016 Team

2017-10-01

After the arrival of the Trace Gas Orbiter (TGO) at Mars on 19 October 2016 a number of initial orbit change manoeuvres were executed and the spacecraft was put in an orbit with a 24 hour period and 74 degrees inclination. The spacecraft and its four instruments were thoroughly checked out after arrival and a few measurements and images were taken in November 2016 and in Feb-March 2017. The solar occultation observations have however not yet been possible due to lack of the proper geometry.On 15 March a long period of aerobraking to reach the final 400km semi-circular frozen orbit (370x430km, with a fixed pericentre latitude). This orbit is optimised for the payload observations and for the communication relay with the ExoMars Rover, due to arrive in 2021.The aerobraking is proceeding well and the final orbit is expected to be reached in April 2018. A large data set is being acquired for the upper atmosphere of Mars, from the limit of the sensitivity of the accelerometer, down to lowest altitude of the aerobraking at about 105km. Initial analysis has shown a highly variable atmosphere with a slightly lower density then predicted by existing models. Until the time of the abstract writing no dust storms have been observed.The ExoMars programme is a joint activity by the European Space Agency(ESA) and ROSCOSMOS, Russia. ESA is providing the TGO spacecraft and Schiaparelli (EDM) and two of the TGO instruments and ROSCOSMOS is providing the Proton launcher and the other two TGO instruments. After the arrival of the ExoMars 2020 mission, consisting of a Rover and a Surface platform also launched by a Proton rocket, the TGO will handle the communication between the Earth and the Rover and Surface Platform through its (NASA provided) UHF communication system.

Short rendezvous missions for advanced Russian human spacecraft

Science.gov (United States)

Murtazin, Rafail F.; Budylov, Sergey G.

2010-10-01

The two-day stay of crew in a limited inhabited volume of the Soyuz-TMA spacecraft till docking to ISS is one of the most stressful parts of space flight. In this paper a number of possible ways to reduce the duration of the free flight phase are considered. The duration is defined by phasing strategy that is necessary for reduction of the phase angle between the chaser and target spacecraft. Some short phasing strategies could be developed. The use of such strategies creates more comfortable flight conditions for crew thanks to short duration and additionally it allows saving spacecraft's life support resources. The transition from the methods of direct spacecraft rendezvous using one orbit phasing (first flights of " Vostok" and " Soyuz" vehicles) to the currently used methods of two-day rendezvous mission can be observed in the history of Soviet manned space program. For an advanced Russian human rated spacecraft the short phasing strategy is recommended, which can be considered as a combination between the direct and two-day rendezvous missions. The following state of the art technologies are assumed available: onboard accurate navigation; onboard computations of phasing maneuvers; launch vehicle with high accuracy injection orbit, etc. Some operational requirements and constraints for the strategies are briefly discussed. In order to provide acceptable phase angles for possible launch dates the experience of the ISS altitude profile control can be used. As examples of the short phasing strategies, the following rendezvous missions are considered: direct ascent, short mission with the phasing during 3-7 orbits depending on the launch date (nominal or backup). For each option statistical modeling of the rendezvous mission is fulfilled, as well as an admissible phase angle range, accuracy of target state vector and addition fuel consumption coming out of emergency is defined. In this paper an estimation of pros and cons of all options is conducted.

Mercury Orbiter: Report of the Science Working Team

Science.gov (United States)

Belcher, John W.; Slavin, James A.; Armstrong, Thomas P.; Farquhar, Robert W.; Akasofu, Syun I.; Baker, Daniel N.; Cattell, Cynthia A.; Cheng, Andrew F.; Chupp, Edward L.; Clark, Pamela E.

1991-01-01

The results are presented of the Mercury Orbiter Science Working Team which held three workshops in 1988 to 1989 under the auspices of the Space Physics and Planetary Exploration Divisions of NASA Headquarters. Spacecraft engineering and mission design studies at the Jet Propulsion Lab were conducted in parallel with this effort and are detailed elsewhere. The findings of the engineering study, summarized herein, indicate that spin stabilized spacecraft carrying comprehensive particles and fields experiments and key planetology instruments in high elliptical orbits can survive and function in Mercury orbit without costly sun shields and active cooling systems.

Detection of gravitational radiation by the Doppler tracking of spacecraft

International Nuclear Information System (INIS)

Mashhoon, B.

1979-01-01

It has been suggested that the residual Doppler shift in the precision electromagnetic tracking of spacecraft be used to search for gravitational radiation that may be incident on the Earth-spacecraft system. The influence of a gravitational wave on the Doppler shift is calculated, and it is found that the residual shift is dominated by two terms: one is due to the passage of electromagnetic waves through the gravitational radiation field, and the other depends on the change in the relative velocity of the Earth and the spacecraft caused by the external wave. A detailed analysis is given of the influence of gravitational radiation on a binary system with an orbital size small compared to the wavelength of the incident radiation. It is shown that, as a consequence of the interaction with the external wave, the system makes a transition from one Keplerian orbit into another which, in general, has a different energy and angular momentum. It is therefore proposed to search for such effects in the solar system. Observations of the orbit of an artificial Earth satellite, the lunar orbit, and especially the planetary orbits offer exciting possibilities for the detection of gravitational waves of various wavelengths. From the results of the lunar laser ranging experiment and the range measurement to Mars, certain interesting limits may be established on the frequency of incidence of gravitational waves of a given flux on the Earth-Moon and the Earth-Mars systems. This is followed by a brief and preliminary analysis of the possibility of detecting gravitational radiation by measuring a residual secular Doppler shift in the satellite-to-satellite Doppler tracking of two counterorbiting drag-free spacecraft around the Earth as in the Van Patten-Everitt experiment

A Compact Device for Colloidal Crystal Studies on Tiangong-1 Target Spacecraft

Science.gov (United States)

Li, Xiao-Long; Hu, Shu-Xin; Sun, Zhi-Bin; Zhai, Yong-Liang; Wu, Lan-Sheng; Huang, Zhen; Li, Wei-Ning; Yang, Han-Dong; Zhai, Guang-Jie; Li, Ming

2014-07-01

An experimental device with three crystallization cells, each with two working positions, was designed to study growth kinetics and structural transformation of colloidal crystals under microgravity condition. The device is capable of remote control of experimental procedures. It uses direct-space imaging with white light to monitor morphology of the crystals and reciprocal-space laser diffraction (Kossel lines) to reveal lattice structure. The device, intended for colloidal crystal growth kinetics and structural transformation on Tiangong-1 target spacecraft, had run on-orbit for more than one year till the end of the mission. Hundreds of images and diffraction patterns were collected via the on-ground data receiving station. The data showed that single crystalline samples were successfully grown on the orbit. Structural transformation was carefully studied under electric and thermal field. Using a backup device, control experiments were also performed on the ground under similar conditions except for the microgravity. Preliminary results indicated that the on-orbit crystals were more stable than the on-ground ones.

Present status of the Japanese Venus climate orbiter

Science.gov (United States)

Nakamura, M.; Imamura, T.; Abe, T.; Ishii, N.

The code name of 24th science spacecraft of ISAS/JAXA is Planet-C. It is the first Venus Climate Orbiter (VCO) of Japan. The ministry of finance of Japan finally agreed to start phase B study of VCO from this April, 2004. We plan 1-2 years phase B study followed by 2 years of flight model integration. The spacecraft will be launched between 2009 and 2010. After arriving Venus, 2 years of operation is expected. VCO will complemet the ESA's Venus Express mission which have several spectrometers and will reveal the composition of the Venusian atmosphere. On the other hand, VCO is designed to reveal the details of the atmospheric motion on Venus and approach the dynamics of the Venusian climate. Cooperation between Japanese VCO and ESA's Venus Express, in the colaboration framework of U.S., Europian, and Japanese scienctist is very important. To elucidate the driving mechanism of the 4-days super-rotation is one of our main targets. We have 4 cameras to take snap shots of the planets in different wave lengths. They are the IR1 camera (1 micron-meter), the IR2 camera (2.4 micron-meter), the LIR camera (10-12 micron-meter), and the UVI camera (340nm). They are attached to the side panel of the 3-axis stabilized spacecraft, and are directed to Venus with the spacecraft's attitude control. Snap shots are expected to be taken every 2 hours. The spacecraft has an orbit of 300km x 13Rv (Venusian radii) with 172 degrees inclination. Orbital period is 30 hours. The angular position of the spacecraft on this orbit is synchronized for 20 hours at its apoapsis with the global atmospheric circulation at the altitude of 50km, thus the snap shots of every 2 hours will be the images of the same side of the atmosphere. In addition to these 4 cameras, we have a Lightning and Airglow camera (LAC) in visible range. This will be operated when the orbiter is close to the planet.

Spin-orbit coupling and the static polarizability of single-wall carbon nanotubes

International Nuclear Information System (INIS)

Diniz, Ginetom S.; Ulloa, Sergio E.

2014-01-01

We calculate the static longitudinal polarizability of single-wall carbon tubes in the long wavelength limit taking into account spin-orbit effects. We use a four-orbital orthogonal tight-binding formalism to describe the electronic states and the random phase approximation to calculate the dielectric function. We study the role of both the Rashba as well as the intrinsic spin-orbit interactions on the longitudinal dielectric response, i.e., when the probing electric field is parallel to the nanotube axis. The spin-orbit interaction modifies the nanotube electronic band dispersions, which may especially result in a small gap opening in otherwise metallic tubes. The bandgap size and state features, the result of competition between Rashba and intrinsic spin-orbit interactions, result in drastic changes in the longitudinal static polarizability of the system. We discuss results for different nanotube types and the dependence on nanotube radius and spin-orbit couplings.

Spin-orbit coupling and the static polarizability of single-wall carbon nanotubes

Energy Technology Data Exchange (ETDEWEB)

Diniz, Ginetom S., E-mail: ginetom@gmail.com; Ulloa, Sergio E. [Department of Physics and Astronomy and Nanoscale and Quantum Phenomena Institute, Ohio University, Athens, Ohio 45701-2979 (United States)

2014-07-14

We calculate the static longitudinal polarizability of single-wall carbon tubes in the long wavelength limit taking into account spin-orbit effects. We use a four-orbital orthogonal tight-binding formalism to describe the electronic states and the random phase approximation to calculate the dielectric function. We study the role of both the Rashba as well as the intrinsic spin-orbit interactions on the longitudinal dielectric response, i.e., when the probing electric field is parallel to the nanotube axis. The spin-orbit interaction modifies the nanotube electronic band dispersions, which may especially result in a small gap opening in otherwise metallic tubes. The bandgap size and state features, the result of competition between Rashba and intrinsic spin-orbit interactions, result in drastic changes in the longitudinal static polarizability of the system. We discuss results for different nanotube types and the dependence on nanotube radius and spin-orbit couplings.

Project Overview of the Naval Postgraduate School Spacecraft Architecture and Technology Demonstration Experiment

National Research Council Canada - National Science Library

Reuer, Charles

2001-01-01

The Naval Postgraduate School's current attempt at getting another spacecraft into orbit is focusing on Naval Postgraduate School Spacecraft Architecture and Technology Demonstration Experiment (NPSAT1...

Deployable reflector configurations

Science.gov (United States)

Meinel, A. B.; Meinel, M. P.; Woolf, N. J.

Both the theoretical reasons for considering a non-circular format for the Large Deployable Reflector, and a potentially realizable concept for such a device, are discussed. The optimum systems for diffraction limited telescopes with incoherent detection have either a single filled aperture, or two such apertures as an interferometer to synthesize a larger aperture. For a single aperture of limited area, a reflector in the form of a slot can be used to give increased angular resolution. It is shown how a 20 x 8 meter telescope can be configured to fit the Space Shuttle bay, and deployed with relatively simple operations. The relationship between the sunshield design and the inclination of the orbit is discussed. The possible use of the LDR as a basic module to permit the construction of supergiant space telescopes and interferometers both for IR/submm studies and for the entire ultraviolet through mm wave spectral region is discussed.

Single-Particle Spin-Orbit Splittings in Nuclei

OpenAIRE

Kazuhiko, ANDO; Hiroharu, BANDO; Department of Physics, Kyoto University; Division of Mathematical Physics, Fukui University

1981-01-01

Single-particle spin-orbit splittings (Δ^) in ^O and ^Ca nuclei are evaluated within the framework of the effective interaction theory by employing the Reid soft-core potential and meson-exchange three-body forces (TBF). Among the two-body force contributions, the Pauli-rearrangement effect on Δ^ is studied with special care. The TBF contribution to Δ^ is found to be significant. The G-matrix, the second-order pauli-rearrangement and the TBF contribute to Δ^ by the amount of ～1/2, ～1/5 and ～1...

Feasibility Study for a Near Term Demonstration of Laser-Sail Propulsion from the Ground to Low Earth Orbit

Science.gov (United States)

Montgomery, Edward E., IV; Johnson, Les; Thomas, Herbert D.

2016-01-01

This paper adds to the body of research related to the concept of propellant-less in-space propulsion utilizing an external high energy laser (HEL) to provide momentum to an ultra-lightweight (gossamer) spacecraft. It has been suggested that the capabilities of Space Situational Awareness assets and the advanced analytical tools available for fine resolution orbit determination make it possible to investigate the practicalities of a ground to Low Earth Orbit (LEO) demonstration at delivered power levels that only illuminate a spacecraft without causing damage to it. The degree to which this can be expected to produce a measurable change in the orbit of a low ballistic coefficient spacecraft is investigated. Key system characteristics and estimated performance are derived for a near term mission opportunity involving the LightSail 2 spacecraft and laser power levels modest in comparison to those proposed previously by Forward, Landis, or Marx. [1,2,3] A more detailed investigation of accessing LightSail 2 from Santa Rosa Island on Eglin Air Force Base on the United States coast of the Gulf of Mexico is provided to show expected results in a specific case.

The Mars Reconnaissance Orbiter Mission: 10 Years of Exploration from Mars Orbit

Science.gov (United States)

Johnston, M. Daniel; Zurek, Richard W.

2016-01-01

The Mars Reconnaissance Orbiter ( MRO ) entered Mars orbit on March 10, 2006. After five months of aerobraking, a series of propulsive maneuvers were used to establish the desired low -altitude science orbit. The spacecraft has been on station in its 255 x 320 k m, sun -synchronous (approximately 3 am -pm ), primary science orbit since September 2006 performing both scientific and Mars programmatic support functions. This paper will provide a summary of the major achievements of the mission to date and the major flight activities planned for the remainder of its third Extended Mission (EM3). Some of the major flight challenges the flight team has faced are also discussed.

Magnus Effect on a Spinning Satellite in Low Earth Orbit

Science.gov (United States)

Ramjatan, Sahadeo; Fitz-Coy, Norman; Yew, Alvin Garwai

2016-01-01

A spinning body in a flow field generates an aerodynamic lift or Magnus effect that displaces the body in a direction normal to the freestream flow. Earth orbiting satellites with substantial body rotation in appreciable atmospheric densities may generate a Magnus force to perturb orbital dynamics. We investigate the feasibility of using this effect for spacecraft at a perigee of 80km using the Systems Tool Kit (STK). Results show that for a satellite of reasonable properties, the Magnus effect doubles the amount of time in orbit. Orbital decay was greatly mitigated for satellites spinning at 10000 and 15000RPM. This study demonstrates that the Magnus effect has the potential to sustain a spacecraft's orbit at a low perigee altitude and could also serve as an orbital maneuver capability.

Small Orbital Stereo Tracking Camera Technology Development

Science.gov (United States)

Gagliano, L.; Bryan, T.; MacLeod, T.

On-Orbit Small Debris Tracking and Characterization is a technical gap in the current National Space Situational Awareness necessary to safeguard orbital assets and crew. This poses a major risk of MOD damage to ISS and Exploration vehicles. In 2015 this technology was added to NASAs Office of Chief Technologist roadmap. For missions flying in or assembled in or staging from LEO, the physical threat to vehicle and crew is needed in order to properly design the proper level of MOD impact shielding and proper mission design restrictions. Need to verify debris flux and size population versus ground RADAR tracking. Use of ISS for In-Situ Orbital Debris Tracking development provides attitude, power, data and orbital access without a dedicated spacecraft or restricted operations on-board a host vehicle as a secondary payload. Sensor Applicable to in-situ measuring orbital debris in flux and population in other orbits or on other vehicles. Could enhance safety on and around ISS. Some technologies extensible to monitoring of extraterrestrial debris as well To help accomplish this, new technologies must be developed quickly. The Small Orbital Stereo Tracking Camera is one such up and coming technology. It consists of flying a pair of intensified megapixel telephoto cameras to evaluate Orbital Debris (OD) monitoring in proximity of International Space Station. It will demonstrate on-orbit optical tracking (in situ) of various sized objects versus ground RADAR tracking and small OD models. The cameras are based on Flight Proven Advanced Video Guidance Sensor pixel to spot algorithms (Orbital Express) and military targeting cameras. And by using twin cameras we can provide Stereo images for ranging & mission redundancy. When pointed into the orbital velocity vector (RAM), objects approaching or near the stereo camera set can be differentiated from the stars moving upward in background.

Standardized spacecraft: a methodology for decision making. AMS report No. 1199

International Nuclear Information System (INIS)

Greenberg, J.S.; Nichols, R.A.

1974-01-01

As the space program matures, more and more attention is being focused on ways to reduce the costs of performing space missions. Standardization has been suggested as a way of providing cost reductions. The question of standardization at the system level, in particular, the question of the desirability of spacecraft standardization for geocentric space missions is addressed. The spacecraft is considered to be a bus upon which mission oriented equipment, the payload, is mounted. Three basic questions are considered: (1) is spacecraft standardization economically desirable; (2) if spacecraft standardization is economically desirable, what standardized spacecraft configuration or mix of configurations and technologies should be developed; and (3) if standardized spacecraft are to be developed, what power levels should they be designed for. A methodology which has been developed and which is necessary to follow if the above questions are to be answered and informed decisions made relative to spacecraft standardization is described. To illustrate the decision making problems and the need for the developed methodology and the data requirements, typical standardized spacecraft have been considered. Both standardized solar and nuclear-powered spacecraft and mission specialized spacecraft have been conceptualized and performance and cost estimates have been made. These estimates are not considered to be of sufficient accuracy to allow decisions regarding spacecraft mix and power levels to be made at this time. The estimates are deemed of sufficient accuracy so as to demonstrate the desirability of spacecraft standardization and the methodology (as well as the need for the methodology) which is necessary to decide upon the best mix of standardized spacecraft and their design power levels. (U.S.)

Control of Single-Stage Single-Phase PV inverter

DEFF Research Database (Denmark)

Ciobotaru, Mihai; Teodorescu, Remus; Blaabjerg, Frede

2005-01-01

In this paper the issue of control strategies for single-stage photovoltaic (PV) inverter is addressed. Two different current controllers have been implemented and an experimental comparison between them has been made. A complete control structure for the single-phase PV system is also presented......-forward; - and the grid current controller implemented in two different ways, using the classical proportional integral (PI) and the novel proportional resonant (PR) controllers. The control strategy was tested experimentally on 1.5 kW PV inverter....

Spin flip in single quantum ring with Rashba spin–orbit interation

Science.gov (United States)

Liu, Duan-Yang; Xia, Jian-Bai

2018-03-01

We theoretically investigate spin transport in the elliptical ring and the circular ring with Rashba spin–orbit interaction. It is shown that when Rashba spin–orbit interaction is relatively weak, a single circular ring can not realize spin flip, however an elliptical ring may work as a spin-inverter at this time, and the influence of the defect of the geometry is not obvious. Howerver if a giant Rashba spin–orbit interaction strength has been obtained, a circular ring can work as a spin-inverter with a high stability. Project supported by the National Natural Science Foundation of China (Grant No. 11504016).

New results for single stage low energy carbon AMS

International Nuclear Information System (INIS)

Klody, G.M.; Schroeder, J.B.; Norton, G.A.; Loger, R.L.; Kitchen, R.L.; Sundquist, M.L.

2005-01-01

A new configuration of the NEC single stage, low energy carbon AMS system (U.S. Patent 6,815,666 B2) has been built and tested. The injector includes two 40-sample ion sources, electrostatic and magnetic analysis, and fast sequential injection. The gas stripper, analyzing magnet, electrostatic analyzer, and detector are on an open air 250 kV deck. Both 12 C and 13 C currents are measured on the deck after the stripper, and an SSB detector is used for 14 C counting. Injected 12 C and mass 13 ( 13 C and 12 CH) currents are also measured. Automated controls follow a user-specified run list for unattended operation. Initial test results show precision for 14 C/ 12 C ratios of better than 5 per mil, and backgrounds for unprocessed graphite of less than 0.005 x modern. We will report final results for precision, background, and throughput and discuss related design features

Link Analysis of High Throughput Spacecraft Communication Systems for Future Science Missions

Science.gov (United States)

Simons, Rainee N.

2015-01-01

NASA's plan to launch several spacecrafts into low Earth Orbit (LEO) to support science missions in the next ten years and beyond requires down link throughput on the order of several terabits per day. The ability to handle such a large volume of data far exceeds the capabilities of current systems. This paper proposes two solutions, first, a high data rate link between the LEO spacecraft and ground via relay satellites in geostationary orbit (GEO). Second, a high data rate direct to ground link from LEO. Next, the paper presents results from computer simulations carried out for both types of links taking into consideration spacecraft transmitter frequency, EIRP, and waveform; elevation angle dependent path loss through Earths atmosphere, and ground station receiver GT.

Validation of Galileo orbits using SLR with a focus on satellites launched into incorrect orbital planes

Science.gov (United States)

Sośnica, Krzysztof; Prange, Lars; Kaźmierski, Kamil; Bury, Grzegorz; Drożdżewski, Mateusz; Zajdel, Radosław; Hadas, Tomasz

2018-02-01

The space segment of the European Global Navigation Satellite System (GNSS) Galileo consists of In-Orbit Validation (IOV) and Full Operational Capability (FOC) spacecraft. The first pair of FOC satellites was launched into an incorrect, highly eccentric orbital plane with a lower than nominal inclination angle. All Galileo satellites are equipped with satellite laser ranging (SLR) retroreflectors which allow, for example, for the assessment of the orbit quality or for the SLR-GNSS co-location in space. The number of SLR observations to Galileo satellites has been continuously increasing thanks to a series of intensive campaigns devoted to SLR tracking of GNSS satellites initiated by the International Laser Ranging Service. This paper assesses systematic effects and quality of Galileo orbits using SLR data with a main focus on Galileo satellites launched into incorrect orbits. We compare the SLR observations with respect to microwave-based Galileo orbits generated by the Center for Orbit Determination in Europe (CODE) in the framework of the International GNSS Service Multi-GNSS Experiment for the period 2014.0-2016.5. We analyze the SLR signature effect, which is characterized by the dependency of SLR residuals with respect to various incidence angles of laser beams for stations equipped with single-photon and multi-photon detectors. Surprisingly, the CODE orbit quality of satellites in the incorrect orbital planes is not worse than that of nominal FOC and IOV orbits. The RMS of SLR residuals is even lower by 5.0 and 1.5 mm for satellites in the incorrect orbital planes than for FOC and IOV satellites, respectively. The mean SLR offsets equal -44.9, -35.0, and -22.4 mm for IOV, FOC, and satellites in the incorrect orbital plane. Finally, we found that the empirical orbit models, which were originally designed for precise orbit determination of GNSS satellites in circular orbits, provide fully appropriate results also for highly eccentric orbits with variable linear

Nanoscale multireference quantum chemistry: full configuration interaction on graphical processing units.

Science.gov (United States)

Fales, B Scott; Levine, Benjamin G

2015-10-13

Methods based on a full configuration interaction (FCI) expansion in an active space of orbitals are widely used for modeling chemical phenomena such as bond breaking, multiply excited states, and conical intersections in small-to-medium-sized molecules, but these phenomena occur in systems of all sizes. To scale such calculations up to the nanoscale, we have developed an implementation of FCI in which electron repulsion integral transformation and several of the more expensive steps in σ vector formation are performed on graphical processing unit (GPU) hardware. When applied to a 1.7 × 1.4 × 1.4 nm silicon nanoparticle (Si72H64) described with the polarized, all-electron 6-31G** basis set, our implementation can solve for the ground state of the 16-active-electron/16-active-orbital CASCI Hamiltonian (more than 100,000,000 configurations) in 39 min on a single NVidia K40 GPU.

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

Science.gov (United States)

Guglielmo, David; Omar, Sanny R.; 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

High Gain Antenna Calibration on Three Spacecraft

Science.gov (United States)

Hashmall, Joseph A.

2011-01-01

This paper describes the alignment calibration of spacecraft High Gain Antennas (HGAs) for three missions. For two of the missions (the Lunar Reconnaissance Orbiter and the Solar Dynamics Observatory) the calibration was performed on orbit. For the third mission (the Global Precipitation Measurement core satellite) ground simulation of the calibration was performed in a calibration feasibility study. These three satellites provide a range of calibration situations-Lunar orbit transmitting to a ground antenna for LRO, geosynchronous orbit transmitting to a ground antenna fer SDO, and low Earth orbit transmitting to TDRS satellites for GPM The calibration results depend strongly on the quality and quantity of calibration data. With insufficient data the calibration Junction may give erroneous solutions. Manual intervention in the calibration allowed reliable parameters to be generated for all three missions.

A Memory/Immunology-Based Control Approach with Applications to Multiple Spacecraft Formation Flying

Directory of Open Access Journals (Sweden)

Liguo Weng

2013-01-01

Full Text Available This paper addresses the problem of formation control for multiple spacecrafts in Planetary Orbital Environment (POE. Due to the presence of diverse interferences and uncertainties in the outer space, such as the changing spacecraft mass, unavailable space parameters, and varying gravity forces, traditional control methods encounter great difficulties in this area. A new control approach inspired by human memory and immune system is proposed, and this approach is shown to be capable of learning from past control experience and current behavior to improve its performance. It demands much less system dynamic information as compared with traditional controls. Both theoretic analysis and computer simulation verify its effectiveness.

On the economics of staging for reusable launch vehicles

Science.gov (United States)

Griffin, Michael D.; Claybaugh, William R.

1996-03-01

There has been much recent discussion concerning possible replacement systems for the current U.S. fleet of launch vehicles, including both the shuttle and expendable vehicles. Attention has been focused upon the feasibility and potential benefits of reusable single-stage-to-orbit (SSTO) launch systems for future access to low Earth orbit (LEO). In this paper we assume the technical feasibility of such vehicles, as well as the benefits to be derived from system reusability. We then consider the benefits of launch vehicle staging from the perspective of economic advantage rather than performance necessity. Conditions are derived under which two-stage-to-orbit (TSTO) launch systems, utilizing SSTO-class vehicle technology, offer a relative economic advantage for access to LEO.

Streamlined Modeling for Characterizing Spacecraft Anomalous Behavior

Science.gov (United States)

Klem, B.; Swann, D.

2011-09-01

Anomalous behavior of on-orbit spacecraft can often be detected using passive, remote sensors which measure electro-optical signatures that vary in time and spectral content. Analysts responsible for assessing spacecraft operational status and detecting detrimental anomalies using non-resolved imaging sensors are often presented with various sensing and identification issues. Modeling and measuring spacecraft self emission and reflected radiant intensity when the radiation patterns exhibit a time varying reflective glint superimposed on an underlying diffuse signal contribute to assessment of spacecraft behavior in two ways: (1) providing information on body component orientation and attitude; and, (2) detecting changes in surface material properties due to the space environment. Simple convex and cube-shaped spacecraft, designed to operate without protruding solar panel appendages, may require an enhanced level of preflight characterization to support interpretation of the various physical effects observed during on-orbit monitoring. This paper describes selected portions of the signature database generated using streamlined signature modeling and simulations of basic geometry shapes apparent to non-imaging sensors. With this database, summarization of key observable features for such shapes as spheres, cylinders, flat plates, cones, and cubes in specific spectral bands that include the visible, mid wave, and long wave infrared provide the analyst with input to the decision process algorithms contained in the overall sensing and identification architectures. The models typically utilize baseline materials such as Kapton, paints, aluminum surface end plates, and radiators, along with solar cell representations covering the cylindrical and side portions of the spacecraft. Multiple space and ground-based sensors are assumed to be located at key locations to describe the comprehensive multi-viewing aspect scenarios that can result in significant specular reflection

Configuration interaction wave functions: A seniority number approach

International Nuclear Information System (INIS)

Alcoba, Diego R.; Torre, Alicia; Lain, Luis; Massaccesi, Gustavo E.; Oña, Ofelia B.

2014-01-01

This work deals with the configuration interaction method when an N-electron Hamiltonian is projected on Slater determinants which are classified according to their seniority number values. We study the spin features of the wave functions and the size of the matrices required to formulate states of any spin symmetry within this treatment. Correlation energies associated with the wave functions arising from the seniority-based configuration interaction procedure are determined for three types of molecular orbital basis: canonical molecular orbitals, natural orbitals, and the orbitals resulting from minimizing the expectation value of the N-electron seniority number operator. The performance of these bases is analyzed by means of numerical results obtained from selected N-electron systems of several spin symmetries. The comparison of the results highlights the efficiency of the molecular orbital basis which minimizes the mean value of the seniority number for a state, yielding energy values closer to those provided by the full configuration interaction procedure

Configuration interaction wave functions: A seniority number approach

Energy Technology Data Exchange (ETDEWEB)

Alcoba, Diego R. [Departamento de Física, Facultad de Ciencias Exactas y Naturales, Universidad de Buenos Aires and Instituto de Física de Buenos Aires, Consejo Nacional de Investigaciones Científicas y Técnicas, Ciudad Universitaria, 1428 Buenos Aires (Argentina); Torre, Alicia; Lain, Luis, E-mail: qfplapel@lg.ehu.es [Departamento de Química Física, Facultad de Ciencia y Tecnología, Universidad del País Vasco, Apdo. 644, E-48080 Bilbao (Spain); Massaccesi, Gustavo E. [Departamento de Ciencias Exactas, Ciclo Básico Común, Universidad de Buenos Aires, Ciudad Universitaria, 1428 Buenos Aires (Argentina); Oña, Ofelia B. [Instituto de Investigaciones Fisicoquímicas Teóricas y Aplicadas, Universidad Nacional de La Plata, CCT La Plata, Consejo Nacional de Investigaciones Científicas y Técnicas, Diag. 113 y 64 (S/N), Sucursal 4, CC 16, 1900 La Plata (Argentina)

2014-06-21

This work deals with the configuration interaction method when an N-electron Hamiltonian is projected on Slater determinants which are classified according to their seniority number values. We study the spin features of the wave functions and the size of the matrices required to formulate states of any spin symmetry within this treatment. Correlation energies associated with the wave functions arising from the seniority-based configuration interaction procedure are determined for three types of molecular orbital basis: canonical molecular orbitals, natural orbitals, and the orbitals resulting from minimizing the expectation value of the N-electron seniority number operator. The performance of these bases is analyzed by means of numerical results obtained from selected N-electron systems of several spin symmetries. The comparison of the results highlights the efficiency of the molecular orbital basis which minimizes the mean value of the seniority number for a state, yielding energy values closer to those provided by the full configuration interaction procedure.

Mission operations for unmanned nuclear electric propulsion outer planet exploration with a thermionic reactor spacecraft.

Science.gov (United States)

Spera, R. J.; Prickett, W. Z.; Garate, J. A.; Firth, W. L.

1971-01-01

Mission operations are presented for comet rendezvous and outer planet exploration NEP spacecraft employing in-core thermionic reactors for electric power generation. The selected reference missions are the Comet Halley rendezvous and a Jupiter orbiter at 5.9 planet radii, the orbit of the moon Io. The characteristics of the baseline multi-mission NEP spacecraft are presented and its performance in other outer planet missions, such as Saturn and Uranus orbiters and a Neptune flyby, are discussed. Candidate mission operations are defined from spacecraft assembly to mission completion. Pre-launch operations are identified. Shuttle launch and subsequent injection to earth escape by the Centaur D-1T are discussed, as well as power plant startup and the heliocentric mission phases. The sequence and type of operations are basically identical for all missions investigated.

A Preliminary Formation Flying Orbit Dynamics Analysis for Leonardo-BRDF

Science.gov (United States)

Hughes, Steven P.; Mailhe, Laurie M.

2001-01-01

Leonardo-BRDF is a NASA mission concept proposed to allow the investigation of radiative transfer and its effect on the Earth's climate and atmospheric phenomenon. Enabled by the recent developments in small-satellite and formation flying technology, the mission is envisioned to be composed of an array of spacecraft in carefully designed orbits. The different perspectives provided by a distributed array of spacecraft offer a unique advantage to study the Earth's albedo. This paper presents the orbit dynamics analysis performed in the context of the Leonardo-BRDF science requirements. First, the albedo integral is investigated and the effect of viewing geometry on science return is studied. The method used in this paper, based on Gauss quadrature, provides the optimal formation geometry to ensure that the value of the integral is accurately approximated. An orbit design approach is presented to achieve specific relative orbit geometries while simultaneously satisfying orbit dynamics constraints to reduce formation-keeping fuel expenditure. The relative geometry afforded by the design is discussed in terms of mission requirements. An optimal two-burn initialization scheme is presented with the required delta-V to distribute all spacecraft from a common parking orbit into their appropriate orbits in the formation. Finally, formation-keeping strategies are developed and the associated delta-V's are calculated to maintain the formation in the presence of perturbations.

Configuration mixing for spin-isospin modes

International Nuclear Information System (INIS)

Ichimura, Munetake

2005-01-01

Development of theories of configuration mixing is reviewed, concentrating on their application to spin-isospin modes, especially to the Gamow-Teller transitions. This talk is divided into three historical stages, the first order configuration mixing as the first stage, the second order configuration mixing as the second stage, and the delta-isobar-hole mixing as the third stage

Estimating spacecraft attitude based on in-orbit sensor measurements

DEFF Research Database (Denmark)

Jakobsen, Britt; Lyn-Knudsen, Kevin; Mølgaard, Mathias

2014-01-01

of 2014/15. To better evaluate the performance of the payload, it is desirable to couple the payload data with the satellite's orientation. With AAUSAT3 already in orbit it is possible to collect data directly from space in order to evaluate the performance of the attitude estimation. An extended kalman...... filter (EKF) is used for quaternion-based attitude estimation. A Simulink simulation environment developed for AAUSAT3, containing a "truth model" of the satellite and the orbit environment, is used to test the performance The performance is tested using different sensor noise parameters obtained both...... from a controlled environment on Earth as well as in-orbit. By using sensor noise parameters obtained on Earth as the expected parameters in the attitude estimation, and simulating the environment using the sensor noise parameters from space, it is possible to assess whether the EKF can be designed...

Theory and practice of uncommon molecular electronic configurations.

Science.gov (United States)

Gryn'ova, Ganna; Coote, Michelle L; Corminboeuf, Clemence

2015-01-01

The electronic configuration of the molecule is the foundation of its structure and reactivity. The spin state is one of the key characteristics arising from the ordering of electrons within the molecule's set of orbitals. Organic molecules that have open-shell ground states and interesting physicochemical properties, particularly those influencing their spin alignment, are of immense interest within the up-and-coming field of molecular electronics. In this advanced review, we scrutinize various qualitative rules of orbital occupation and spin alignment, viz., the aufbau principle, Hund's multiplicity rule, and dynamic spin polarization concept, through the prism of quantum mechanics. While such rules hold in selected simple cases, in general the spin state of a system depends on a combination of electronic factors that include Coulomb and Pauli repulsion, nuclear attraction, kinetic energy, orbital relaxation, and static correlation. A number of fascinating chemical systems with spin states that fluctuate between triplet and open-shell singlet, and are responsive to irradiation, pH, and other external stimuli, are highlighted. In addition, we outline a range of organic molecules with intriguing non-aufbau orbital configurations. In such quasi-closed-shell systems, the singly occupied molecular orbital (SOMO) is energetically lower than one or more doubly occupied orbitals. As a result, the SOMO is not affected by electron attachment to or removal from the molecule, and the products of such redox processes are polyradicals. These peculiar species possess attractive conductive and magnetic properties, and a number of them that have already been developed into molecular electronics applications are highlighted in this review. WIREs Comput Mol Sci 2015, 5:440-459. doi: 10.1002/wcms.1233 For further resources related to this article, please visit the WIREs website.

Nano-Satellite Secondary Spacecraft on Deep Space Missions

Science.gov (United States)

Klesh, Andrew T.; Castillo-Rogez, Julie C.

2012-01-01

NanoSat technology has opened Earth orbit to extremely low-cost science missions through a common interface that provides greater launch accessibility. They have also been used on interplanetary missions, but these missions have used one-off components and architectures so that the return on investment has been limited. A natural question is the role that CubeSat-derived NanoSats could play to increase the science return of deep space missions. We do not consider single instrument nano-satellites as likely to complete entire Discovery-class missions alone,but believe that nano-satellites could augment larger missions to significantly increase science return. The key advantages offered by these mini-spacecrafts over previous planetary probes is the common availability of advanced subsystems that open the door to a large variety of science experiments, including new guidance, navigation and control capabilities. In this paper, multiple NanoSat science applications are investigated, primarily for high risk/high return science areas. We also address the significant challenges and questions that remain as obstacles to the use of nano-satellites in deep space missions. Finally, we provide some thoughts on a development roadmap toward interplanetary usage of NanoSpacecraft.

Indirect evidences for a gas/dust torus along the Phobos orbit

International Nuclear Information System (INIS)

Dubinin, E.M.; Lundin, R.; Pissarenko, N.F.; Barabash, S.V.; Zakharov, A.V.; Koskinen, H.; Schwingenshuh, K.; Yeroshenko, Ye.G.

1990-01-01

Observations from the PHOBOS-2 space-craft of plasma and magnetic field effects in the solar wind near Mars suggest that a neutral gas (dust?)torus/ring resides along the orbit of the Martian satellite Phobos. Magnetic cavities (strong decreases of the magnetic field strength) coincident with strong plasma density increases (up to a factor of ten) are observed during the first elliptic transition orbits when the spacecraft approached the Phobos orbits. The characteristic transverse dimension of the structures along the spacecraft orbit is in the range 100-1,000 km. Torus effects also have characteristics similar to the formation of a bow shock with increases of plasma density and ion temperature, and a characteristic deflection of the ion flow. This suggests a rather strong interaction between the solar wind plasma and plasma near Phobos orbit. The interaction appears quite similar to that of the solar wind with a comet. The outgassing of matter from Phobos (and Deimos) is also suggested by plasma observations in the wake/tail of the Martian satellites. Altogether, the authors observations imply that a neutral gas cloud - possibly also associated with a faint dust ring - exists along the Phobos orbit

Preliminary thermal design of the COLD-SAT spacecraft

Science.gov (United States)

Arif, Hugh

1991-01-01

The COLD-SAT free-flying spacecraft was to perform experiments with LH2 in the cryogenic fluid management technologies of storage, supply and transfer in reduced gravity. The Phase A preliminary design of the Thermal Control Subsystem (TCS) for the spacecraft exterior and interior surfaces and components of the bus subsystems is described. The TCS was composed of passive elements which were augmented with heaters. Trade studies to minimize the parasitic heat leakage into the cryogen storage tanks are described. Selection procedure for the thermally optimum on-orbit spacecraft attitude was defined. TRASYS-2 and SINDA'85 verification analysis was performed on the design and the results are presented.

Topography of the Lunar Poles and Application to Geodesy with the Lunar Reconnaissance Orbiter

Science.gov (United States)

Mazarico, Erwan; Neumann, Gregory A.; Rowlands, David D.; Smith, David E.; Zuber, Maria T.

2012-01-01

The Lunar Orbiter Laser Altimeter (LOLA) [1] onboard the Lunar Reconnaissance Orbiter (LRO) [2] has been operating continuously since July 2009 [3], accumulating approx.5.4 billion measurements from 2 billion on-orbit laser shots. LRO s near-polar orbit results in very high data density in the immediate vicinity of the lunar poles, which are each sampled every 2h. With more than 10,000 orbits, high-resolution maps can be constructed [4] and studied [5]. However, this requires careful processing of the raw data, as subtle errors in the spacecraft position and pointing can lead to visible artifacts in the final map. In other locations on the Moon, ground tracks are subparallel and longitudinal separations are typically a few hundred meters. Near the poles, the track intersection angles can be large and the inter-track spacing is small (above 80 latitude, the effective resolution is better than 50m). Precision Orbit Determination (POD) of the LRO spacecraft [6] was performed to satisfy the LOLA and LRO mission requirements, which lead to a significant improvement in the orbit position knowledge over the short-release navigation products. However, with pixel resolutions of 10 to 25 meters, artifacts due to orbit reconstruction still exist. Here, we show how the complete LOLA dataset at both poles can be adjusted geometrically to produce a high-accuracy, high-resolution maps with minimal track artifacts. We also describe how those maps can then feedback to the POD work, by providing topographic base maps with which individual LOLA altimetric measurements can be contributing to orbit changes. These direct altimetry constraints improve accuracy and can be used more simply than the altimetric crossovers [6].

Envisioning a 21st Century, National, Spacecraft Servicing and Protection Infrastructure and Demand Potential: A Logical Development of the Earth Orbit Economy

Science.gov (United States)

Horsham, Gary A.

2003-01-01

The modern world is extremely dependent on thin strings of several hundred civil, military, and commercial spacecraft/satellites currently stationed in space. They provide a steady stream of commerce, defense, and knowledge data. This dependency will in all likelihood increase significantly during this century. A major disruption of any kind in these essential systems and networks could be socially, economically, and politically catastrophic, on a global scale. The development of a space-based, robotic services economy could be useful in mitigating this growing risk, from an efficiency and security standpoint. This paper attempts to suggest what makes sense to invest in next for the logical, economic development of Earth orbit i.e., after ISS completion. It expands on the results of an advanced market research and analysis study that sampled the opinions of several satellite industry executives and presents these results within a broad policy context. The concept of a spacecraft carrier that serves as the nucleus of a national, space-based or on-orbit, robotic services infrastructure is introduced as the next logical step for United States leadership in space. This is viewed as a reasonable and appropriate followon to the development of ELVs and satellites in the 1950s and 1960s, the Space Shuttle/PRLV in the 1970s and 1980s, and the International Space Station (ISS) in the 1980s, 1990s and 2000s. Large-scale experience in LEO-to-GEO spacecraft/satellite servicing and protection by robotic means is assumed to be an indispensable prerequisite or stepping-stone toward the development and preservation of the large scientific exploration facilities that are envisioned by NASA for operation beyond GEO. A balanced, return on national investment (RONI) strategy for space, focused on the provision of enhanced national/homeland security for increased protection, national economic/industrial expansion for increased revenue, and national scientific exploration for increased

A Survey of Cost Estimating Methodologies for Distributed Spacecraft Missions

Science.gov (United States)

Foreman, Veronica L.; Le Moigne, Jacqueline; de Weck, Oliver

2016-01-01

Satellite constellations present unique capabilities and opportunities to Earth orbiting and near-Earth scientific and communications missions, but also present new challenges to cost estimators. An effective and adaptive cost model is essential to successful mission design and implementation, and as Distributed Spacecraft Missions (DSM) become more common, cost estimating tools must become more representative of these types of designs. Existing cost models often focus on a single spacecraft and require extensive design knowledge to produce high fidelity estimates. Previous research has examined the limitations of existing cost practices as they pertain to the early stages of mission formulation, for both individual satellites and small satellite constellations. Recommendations have been made for how to improve the cost models for individual satellites one-at-a-time, but much of the complexity in constellation and DSM cost modeling arises from constellation systems level considerations that have not yet been examined. This paper constitutes a survey of the current state-of-theart in cost estimating techniques with recommendations for improvements to increase the fidelity of future constellation cost estimates. To enable our investigation, we have developed a cost estimating tool for constellation missions. The development of this tool has revealed three high-priority shortcomings within existing parametric cost estimating capabilities as they pertain to DSM architectures: design iteration, integration and test, and mission operations. Within this paper we offer illustrative examples of these discrepancies and make preliminary recommendations for addressing them. DSM and satellite constellation missions are shifting the paradigm of space-based remote sensing, showing promise in the realms of Earth science, planetary observation, and various heliophysical applications. To fully reap the benefits of DSM technology, accurate and relevant cost estimating capabilities

Delayed Single Stage Perineal Posterior Urethroplasty.

Science.gov (United States)

Ali, Shahzad; Shahnawaz; Shahzad, Iqbal; Baloch, Muhammad Umar

2015-06-01

To determine the delayed single stage perineal posterior urethroplasty for treatment of posterior urethral stricture/distraction defect. Descriptive case series. Department of Urology, Jinnah Postgraduate Medical Centre, Karachi, from January 2009 to December 2011. Patients were selected for delayed single stage perineal posterior urethroplasty for treatment of posterior urethral stricture / distraction defect. All were initially suprapubically catheterized followed by definitive surgery after at least 3 months. Thirty male patients were analyzed with a mean follow-up of 10 months, 2 patients were excluded as they developed failure in first 3 months postoperatively. Mean patient's age was 26.25 ± 7.9 years. On follow-up, 7 patients (23.3%) experienced recurrent stricture during first 10 months. Five (16.6%) patients were treated successfully with single direct visual internal urethrotomy. Two patients (6.6%) had more than one direct visual internal urethrotomy and considered failed. Re-do perineal urethroplasty was eventually performed. The overall success rate was 93.3% with permissive criteria allowing single direct visual internal urethrotomy and 76.6% with strict criteria allowing no more procedures postoperatively. Posterior anastomotic urethroplasty offers excellent long-term results to patients with posterior urethral trauma and distraction defect even after multiple prior procedures.

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

Science.gov (United States)

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.

Analysis on coverage ability of BeiDou navigation satellite system for manned spacecraft

Science.gov (United States)

Zhao, Sihao; Yao, Zheng; Zhuang, Xuebin; Lu, Mingquan

2014-12-01

To investigate the service ability of the BeiDou Navigation Satellite System (BDS) for manned spacecraft, both the current regional and the future-planned global constellations of BDS are introduced and simulated. The orbital parameters of the International Space Station and China's Tiangong-1 spacelab are used to create the simulation scenario and evaluate the performance of the BDS constellations. The number of visible satellites and the position dilution (PDOP) of precision at the spacecraft-based receiver are evaluated. Simulation and analysis show quantitative results on the coverage ability and time percentages of both the current BDS regional and future global constellations for manned-space orbits which can be a guideline to the applications and mission design of BDS receivers on manned spacecraft.

Contrast in low-cost operational concepts for orbiting satellites

Science.gov (United States)

Walyus, Keith D.; Reis, James; Bradley, Arthur J.

2002-12-01

Older spacecraft missions, especially those in low Earth orbit with telemetry intensive requirements, required round-the-clock control center staffing. The state of technology relied on control center personnel to continually examine data, make decisions, resolve anomalies, and file reports. Hubble Space Telescope (HST) is a prime example of this description. Technological advancements in hardware and software over the last decade have yielded increases in productivity and operational efficiency, which result in lower cost. The re-engineering effort of HST, which has recently concluded, utilized emerging technology to reduce cost and increase productivity. New missions, of which NASA's Transition Region and Coronal Explorer Satellite (TRACE) is an example, have benefited from recent technological advancements and are more cost-effective than when HST was first launched. During its launch (1998) and early orbit phase, the TRACE Flight Operations Team (FOT) employed continually staffed operations. Yet once the mission entered its nominal phase, the FOT reduced their staffing to standard weekday business hours. Operations were still conducted at night and during the weekends, but these operations occurred autonomously without compromising their high standards for data collections. For the HST, which launched in 1990, reduced cost operations will employ a different operational concept, when the spacecraft enters its low-cost phase after its final servicing mission in 2004. Primarily due to the spacecraft"s design, the HST Project has determined that single-shift operations will introduce unacceptable risks for the amount of dollars saved. More importantly, significant cost-savings can still be achieved by changing the operational concept for the FOT, while still maintaining round-the-clock staffing. It"s important to note that the low-cost solutions obtained for one satellite may not be applicable for other satellites. This paper will contrast the differences between

Role of bumpy fields on single particle orbit in near quasihelically symmetric stellarators

International Nuclear Information System (INIS)

Seol, JaeChun; Hegna, C.C.

2004-01-01

The role of symmetry breaking on single particle orbits in near helically symmetric stellarators is investigated. In particular, the effect of a symmetry-breaking bumpy term is included in the analysis of trapped particle orbits. It is found that all trapped particle drift orbits are determined by surfaces on which vertical bar B vertical bar min is constant. Trapped particle orbits reside on these surfaces regardless of pitch angle and are determined solely by the initial position and the shape of the vertical bar B vertical bar min contour. Since vertical bar B vertical bar min contours do not depend on the direction of the banana center motion, superbanana orbits do not appear

Autonomous orbit determination and its error analysis for deep space using X-ray pulsar

International Nuclear Information System (INIS)

Feng, Dongzhu; Yuan, Xiaoguang; Guo, Hehe; Wang, Xin

2014-01-01

Autonomous orbit determination (OD) is a complex process using filtering method to integrate observation and orbit dynamic model effectively and estimate the position and velocity of a spacecraft. As a novel technology for autonomous interplanetary OD, X-ray pulsar holds great promise for deep space exploration. The position and velocity of spacecraft should be estimated accurately during the OD process. However, under the same condition, the accuracy of OD can be greatly reduced by the error of the initial orbit value and the orbit mutation. To resolve this problem, we propose a novel OD method, which is based on the X-ray pulsar measurement and Adaptive Unscented Kalman Filter (AUKF). The accuracy of OD can be improved obviously because the AUKF estimates the orbit of spacecraft using measurement residual. During the simulation, the orbit of Phoenix Mars Lander, Deep Impact Probe, and Voyager 1 are selected. Compared with Unscented Kalman Filter (UKF) and Extended Kalman Filter (EKF), the simulation results demonstrate that the proposed OD method based on AUKF can accurately determinate the velocity and position and effectively decrease the orbit estimated errors which is caused by the orbit mutation and orbit initial errors. (authors)

Study to adapt solar electric propulsion to the Pioneer F and G spacecraft

Science.gov (United States)

1972-01-01

The addition of an electric thrust subsystem to the spin-stabilized Pioneer F and G spacecraft to improve performance capability for certain missions is discussed. The evaluation was performed for the Atlas and Titan launch vehicles with Centaur and TE-364-4 stages and for electric thrust stages of 8- and 5-kw with three 30- and five 15-cm thrusters respectively. The combination of a spinning spacecraft with electric propulsion is a concept only recently evaluated and the penalty from spinning over three-axis stabilized is not as significant as might initally be thought. There are major gains in weight, cost, and reliability, the disadvantages being lower data rate during the thrust phase and less efficient pointing. A variety of missions were evaluated from a solar approach mission into 0.14 AU to a flyby mission of Neptune at approximately 30 AU. Performance improvements were present for all missions evaluated.

Thermodynamic analysis of single-stage and multi-stage adsorption refrigeration cycles with activated carbon–ammonia working pair

International Nuclear Information System (INIS)

Xu, S.Z.; Wang, L.W.; Wang, R.Z.

2016-01-01

Highlights: • Activated carbon–ammonia multi-stage adsorption refrigerator was analyzed. • COP, exergetic efficiency and entropy production of cycles were calculated. • Single-stage cycle usually has the advantages of simple structure and high COP. • Multi-stage cycles adapt to critical conditions better than single-stage cycle. • Boundary conditions for choosing optimal cycle were summarized as tables. - Abstract: Activated carbon–ammonia multi-stage adsorption refrigeration cycle was analyzed in this article, which realized deep-freezing for evaporating temperature under −18 °C with heating source temperature much lower than 100 °C. Cycle mathematical models for single, two and three-stage cycles were established on the basis of thorough thermodynamic analysis. According to simulation results of thermodynamic evaluation indicators such as COP (coefficient of performance), exergetic efficiency and cycle entropy production, multi-stage cycle adapts to high condensing temperature, low evaporating temperature and low heating source temperature well. Proposed cycle with selected working pair can theoretically work under very severe conditions, such as −25 °C evaporating temperature, 40 °C condensing temperature, and 70 °C heating source temperature, but under these working conditions it has the drawback of low cycle adsorption quantity. It was found that both COP and exergetic efficiency are of great reference value in the choice of cycle, whereas entropy production is not so useful for cycle stage selection. Finally, the application boundary conditions of single-stage, two-stage, and three-stage cycles were summarized as tables according to the simulation results, which provides reference for choosing optimal cycle under different conditions.

Transition matrices and orbitals from reduced density matrix theory

Energy Technology Data Exchange (ETDEWEB)

Etienne, Thibaud [Université de Lorraine – Nancy, Théorie-Modélisation-Simulation, SRSMC, Boulevard des Aiguillettes 54506, Vandoeuvre-lès-Nancy (France); CNRS, Théorie-Modélisation-Simulation, SRSMC, Boulevard des Aiguillettes 54506, Vandoeuvre-lès-Nancy (France); Unité de Chimie Physique Théorique et Structurale, Université de Namur, Rue de Bruxelles 61, 5000 Namur (Belgium)

2015-06-28

In this contribution, we report two different methodologies for characterizing the electronic structure reorganization occurring when a chromophore undergoes an electronic transition. For the first method, we start by setting the theoretical background necessary to the reinterpretation through simple tensor analysis of (i) the transition density matrix and (ii) the natural transition orbitals in the scope of reduced density matrix theory. This novel interpretation is made more clear thanks to a short compendium of the one-particle reduced density matrix theory in a Fock space. The formalism is further applied to two different classes of excited states calculation methods, both requiring a single-determinant reference, that express an excited state as a hole-particle mono-excited configurations expansion, to which particle-hole correlation is coupled (time-dependent Hartree-Fock/time-dependent density functional theory) or not (configuration interaction single/Tamm-Dancoff approximation). For the second methodology presented in this paper, we introduce a novel and complementary concept related to electronic transitions with the canonical transition density matrix and the canonical transition orbitals. Their expression actually reflects the electronic cloud polarisation in the orbital space with a decomposition based on the actual contribution of one-particle excitations from occupied canonical orbitals to virtual ones. This approach validates our novel interpretation of the transition density matrix elements in terms of the Euclidean norm of elementary transition vectors in a linear tensor space. A proper use of these new concepts leads to the conclusion that despite the different principles underlying their construction, they provide two equivalent excited states topological analyses. This connexion is evidenced through simple illustrations of (in)organic dyes electronic transitions analysis.

Benefits of a Single-Person Spacecraft for Weightless Operations. [(Stop Walking and Start Flying)

Science.gov (United States)

Griffin, Brand N.

2012-01-01

Historically, less than 20 percent of crew time related to extravehicular activity (EVA) is spent on productive external work.1 A single-person spacecraft with 90 percent efficiency provides productive new capabilities for maintaining the International Space Station (ISS), exploring asteroids, and servicing telescopes or satellites. With suits, going outside to inspect, service or repair a spacecraft is time-consuming, requiring pre-breathe time, donning a fitted space suit, and pumping down an airlock. For ISS, this is between 12.5 and 16 hours for each EVA, not including translation and work-site set up. The work is physically demanding requiring a day of rest between EVAs and often results in suit-induced trauma with frequent injury to astronauts fingers2. For maximum mobility, suits use a low pressure, pure oxygen atmosphere. This represents a fire hazard and requires pre-breathing to reduce the risk of decompression sickness (bends). With virtually no gravity, humans exploring asteroids cannot use legs for walking. The Manned Maneuvering Unit offers a propulsive alternative however it is no longer in NASA s flight inventory. FlexCraft is a single person spacecraft operating at the same cabin atmosphere as its host so there is no risk of the bends and no pre-breathing. This allows rapid, any-time access to space for repeated short or long EVAs by different astronauts. Integrated propulsion eliminates hand-over-hand translation or having another crew member operate the robotic arm. The one-size-fits-all FlexCraft interior eliminates the suit part inventory and crew time required to fit all astronauts. With a shirtsleeve cockpit, conventional displays and controls are used and because the work is not strenuous no rest days are required. Furthermore, there is no need for hand tools because manipulators are equipped with force multiplying end-effectors that can deliver the precise torque for the job.

Deadly Sunflower Orbits

Science.gov (United States)

Hamilton, Douglas P.

2018-04-01

Solar radiation pressure is usually very effective at removing hazardous millimeter-sized debris from distant orbits around asteroidsand other small solar system bodies (Hamilton and Burns 1992). Theprimary loss mechanism, driven by the azimuthal component of radiationpressure, is eccentricity growth followed by a forced collision withthe central body. One large class of orbits, however, neatly sidestepsthis fate. Orbits oriented nearly perpendicular to the solar directioncan maintain their face-on geometry, oscillating slowly around a stableequilibrium orbit. These orbits, designated sunflower orbits, arerelated to terminator orbits studied by spacecraft mission designers(Broschart etal. 2014).Destabilization of sunflower orbits occurs only for particles smallenough that radiation pressure is some tens of percent the strength ofthe central body's direct gravity. This greatly enhanced stability,which follows from the inability of radiation incident normal to theorbit to efficiently drive eccentricities, presents a threat tospacecraft missions, as numerous dangerous projectiles are potentiallyretained in orbit. We have investigated sunflower orbits insupport of the New Horizons, Aida, and Lucy missions and find thatthese orbits are stable for hazardous particle sizes at asteroids,comets, and Kuiper belt objects of differing dimensions. Weinvestigate the sources and sinks for debris that might populate suchorbits, estimate timescales and equilibrium populations, and willreport on our findings.

Automated Spacecraft Conjunction Assessment at Mars and the Moon

Science.gov (United States)

Berry, David; Guinn, Joseph; Tarzi, Zahi; Demcak, Stuart

2012-01-01

Conjunction assessment and collision avoidance are areas of current high interest in space operations. Most current conjunction assessment activity focuses on the Earth orbital environment. Several of the world's space agencies have satellites in orbit at Mars and the Moon, and avoiding collisions there is important too. Smaller number of assets than Earth, and smaller number of organizations involved, but consequences similar to Earth scenarios.This presentation will examine conjunction assessment processes implemented at JPL for spacecraft in orbit at Mars and the Moon.

Spectral properties of minimal-basis-set orbitals: Implications for molecular electronic continuum states

Science.gov (United States)

Langhoff, P. W.; Winstead, C. L.

Early studies of the electronically excited states of molecules by John A. Pople and coworkers employing ab initio single-excitation configuration interaction (SECI) calculations helped to simulate related applications of these methods to the partial-channel photoionization cross sections of polyatomic molecules. The Gaussian representations of molecular orbitals adopted by Pople and coworkers can describe SECI continuum states when sufficiently large basis sets are employed. Minimal-basis virtual Fock orbitals stabilized in the continuous portions of such SECI spectra are generally associated with strong photoionization resonances. The spectral attributes of these resonance orbitals are illustrated here by revisiting previously reported experimental and theoretical studies of molecular formaldehyde (H2CO) in combination with recently calculated continuum orbital amplitudes.